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Home , Apollo Lunar Module, European Venus Explorer, Hard landing, NetLander, Raman Laser Spectrometer, Samad (crater)

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Ethiraj Venkatapathy1, Chief Technologist, Entry Systems and Technology Division, NASA ARC, 2 Ali Gülhan , Department Head, Supersonic and Hypersonic Technologies Department, DLR, Cologne, and Michelle Munk3, Principal Technologist, EDL, Space Technology Mission Directorate, NASA.

1 NASA Ames Research Center, Moffett Field, CA [email protected]. 2 Deutsches Zentrum für Luft- und Raumfahrt e.V. (DLR), German Aerospace Center, [email protected] 3 NASA Langley Research Center, Hampron, VA. Michelle.m.Munk@nasa.gov

Abstract of the Proposed Talk:

One of the goals of IPPW has been to bring about international collaboration. Establishing collaboration, especially in the area of EDL, can present numerous frustrating challenges. IPPW presents opportunities to present advances in various technology areas. It allows for opportunity for general discussion. Evaluating collaboration potential requires open dialogue as to the needs of the parties and what critical capabilities each party possesses. Understanding opportunities for collaboration as well as the rules and regulations that govern collaboration are essential. The authors of this proposed talk have explored and established collaboration in multiple areas of interest to IPPW community. The authors will present examples that illustrate the motivations for the partnership, our common goals, and the unique capabilities of each party.

The first example involves earth entry of a large asteroid and break-up. NASA Ames is leading an effort for the agency to assess and estimate the threat posed by large asteroids under the Asteroid Threat Assessment Project (ATAP). Asteroids of size ranging from 10 m to 50 m, much larger, heavier and higher velocity than typical spacecrafts, can cause significant damage as a result of break-up in the atmosphere and sending shock waves towards the ground. Depending on the compositional characteristics of the asteroid and size, typically less than ~5% of the asteroid original mass reaches the surface as a result of break-up and spreading. The spreading of the broken material is postulated to deliver energy to the atmosphere in the form of strong shock wave that travels towards earth. This is akin to sonic boom created by supersonic aircrafts but the strength of the shockwave is much more devastating. NASA and DLR have partnered together to understand the physics behind this phenomenon and have been conducting experiments and computations for over a year. The study of the asteroid entry and break-up problem is also relevant to spacecraft break-up, such as large satellites, during their demise. They are also relevant for designing entry systems for mission assurance during separation of heat-shield and back-shell prior to the deployment of parachute along with science payload.

The second example involves planetary entry into planets across solar system. The entry system design requires a better understanding of the aero-thermodynamic characteristics of the flow during entry in different atmospheres on the heatshield and more importantly on the back-shell. In the past, opportunities to instrument flight articles have been neglected due to many reasons. Strong advocacy by the EDL community in the US led to NASA leadership instituting requirements for Engineering Science Investigation (ESI). A direct result of this was instrumented MSL heatshield and now the instrumented aeroshell of Mars 2020. ESA instrument the Mars EDM backshell with COMARS+ sensors. Though the landing of the lander Schiaparelli was not successful, COMARS+ sensor returned data during entry and descent. NASA’s New Frontier proposals, due at the end of April, 2017, requires proposers to respond to the ESI requirements and encourages cost-effective sensors to be part of the proposal. The New Frontiers is a perfect opportunity for international partnership. NASA Ames is working in partnership with DLR and New Frontiers-4 proposal teams in evaluating the viability of the COMARS+ sensor for a host of destinations. This is a first step in the competitive proposal process. Successful selection of a mission that can utilize COMARS+ sensors is necessary and in order to have successful mission implementation has many challenges that need to be tackled.

The third example involves Mars. Dust storms at Mars is a well observed phenomenon and yet, the effect of dust, especially high altitude dust that has the potential to cause significant damage to entry heat-shield is poorly understood. Modeling of the heat-shield performance and risk under dusty environment requires not only the characteristics of the dust and also the response of the specific heat-shield material including system aspects such as seams and gaps. NASA’s human missions to Mars will involve entry system that could be an order of magnitude larger and may involve materials that are more susceptible to dust and if robustness is a requirement, then understanding and accounting for dust effects become very important. NASA’s robotic science missions to Mars use very light weight ablative systems such as SLA or PICA. The response of these systems to dust are poor at best. DLR has developed specialized arc jet test facility and test methodology to study dust erosion. A potential for collaboration is currently being explored to develop fundamental understanding of the dust erosion and validate models with data from DLR facility.

In this talk, in addition to presenting the goals and technical problem dimensions, focus will be on what makes an international collaboration feasible and acceptable. We will also show how successful collaboration can bring about greater understanding that benefits the community at large. AMBIENT MAGNETIC ENERGY HARVESTING AS AN ASSISTING POWER SOURCE A. Aguilar1, T. Larson2 and A. Davies3 1Student, University of Idaho, 709 S Deakin St, Moscow, ID 83844, [email protected], 2Student, University of Idaho, 709 S Deakin St, Moscow, ID 83844 [email protected], 3Student, University of Idaho, 709 S Deakin St, Moscow, ID 83844, da- [email protected].

Introduction: The purpose of this project is to investigate the viability of an ambient magnetic energy harvesting and conversion system as an assisting power source applied to CubeSat and SmallSat technology. As CubeSats and SmallSats become increasingly popu- lar for space exploration and usage as probes beyond Earth, alternative power sources should be investigated in order to keep the costs and weight of these satellites down, while still equipping them with sufficient instrumentation and measurement technology. The ambient magnetic energy harvesting and conversion system would carry less mass than battery packs and function in environments where solar energy is not viable such as beneath the clouds of Venus, or in the outer solar system. While this technology is not meant to be a primary source of power, it will be useful for providing additional power during peak load times as well as charging on-board lithium ion batteries. For a magnetic field, a loop antenna is analogous to an electric dipole in an electric field [1]. Since magnetic fields are more prevalent in space environments and will likely be present in future CubeSat and SmallSat outer planet exploration, a multi-turn loop antenna was chosen for harvesting and conversion analysis. Using a receiving antenna and impedance matching network, planetary magnetic fields incident upon the antenna, assumed to be traveling through the magnetic field over space and time, result in voltage and current in accordance with Faraday’s Law and that can be utilized and transferred to an electrical load. In no load conditions, the antenna has an induced voltage of 1.324 uV, and can transfer 1.306 pW under impedance matched conditions. References: [1] C. A. Balanis (2016) Antenna Theory Analysis and Design

INTERPLANETARY NAVIGATION MISSION SUPPORT SYSTEM. M. Feher*1, T. Renteria*1,J. Rodriguez* 1, P. Wu*1, P. Papadopoulos**1, 1RE: Department of Aerospace Engineering San Jose State University, 1 Washington Square, San Jose, CA 95192.

The current publication presents a potential solution to deep space interplanetary positioning challenges. Introduction: The design of the interplanetary The future of the SNS is the development of navigation mission support system is expanding the Earth’s SNC. The advantage of Earth’s SNC comes capacity for high accuracy research missions in orbits from creating a more robust infrastructure for high beyond Medium Earth Orbit (MEO). Current satellite accuracy travel within the Earth's sphere of influence. positioning systems are generally in MEO, sending Future SNC’s will be expanded into a network of time and position data towards the Earth. These SNC’s that will support missions that travel to any primarily service ground and Low Earth Orbit (LEO) planetary body within the solar system and further into systems. Current satellite systems that operate beyond deep space. MEO are reliant on trajectory prediction models, radio signal delay and inertia measuring units (IMU) to approximate their position and velocity vector. The work conducted on the Solar Navigation Satellite (SNS) is intended to mitigate the high cost of modern navigation subsystems onboard vehicles that travel beyond MEO. The SNS (a high precision, low cost, microsatellite) is designed to provide full 360° spherical coverage with time and position signals travelling on current navigation frequencies. This allows for greater accuracy for missions that have destinations beyond MEO. The SNS is a part of the Solar Navigation Constellation (SNC) architecture which is designed for use around other orbiting bodies (i.e. Mars, Venus, Io, Asteroid 433 Eros, etc.) with a scalable number of SNS’s.

The Structures subsystem is designed to be modified for each mission’s purpose and is modeled after a skeletonized cubesat. The Command and Data Handling (C&DH) subsystem utilizes off the shelf microprocessors and single board computers to manage integrated sensors, maintain orbit, and insure proper satellite functionality. The Communications subsystem is using off the shelf wireless antennae to link two computers together using radio waves as well as a secondary antenna array to broadcast time and position data on current GPS frequencies. The Power subsystem manages the transmission of power collected by the solar panel and delivered to the battery cells as well as the distribution of power to the all subsystems. The SNS integrates with the current GPS system by providing time and position information to standard and specialized GPS receivers that are commercially available. This allows the SNS to aid future missions that travel beyond MEO by providing accurate information on a continuous basis, thereby increasing the mission flight accuracy.

* AIAA Members * SJSU, Aerospace Engineering Senior Students ** SJSU Aerospace Engineering, Faculty Advisor ANALYSIS OF LANDING TRAJECTORIES INTO THE DIDYMOS BINARY SYSTEM OF ASTEROIDS. N. Gerbal1, Ö. Karatekin1, J. Henry de Frahan1 1Royal Observatory of Belgium, Av. Circulaire 3, Brussels/Uccle 1180, Belgium, ([email protected])

The Asteroid Geophysical Explorer (AGEX) concept has been developed in the context of the Cubesat Opportunity Payload (COPINS) for ESA’s Asteroid Impact Mission (AIM). Its aim is to characterize the 65803 Didymos binary asteroid in complement to AIM’s own scientific objectives by unique in-situ investigations as well as to demonstrate new technologies for future deep space missions. The Didymos system is represented by two asteroids: a primary asteroid (“Didymain”) with a diameter of about 775 meters which is orbited by a secondary (“Didymoon”) with a diameter of about 163 meters. AGEX is composed of two 3U cubesats and one of the primary objectives is to study the mechanical properties of the surface material as well as the subsurface. To do so, one of the cubesats will be sent by the AIM main spacecraft toward the system of asteroids in order to land on the secondary. The seismometers and gravimeter on-board the lander cubesat can analyze the average sub-surface seismic properties and surface gravity of Didymoon, as well as characterize the DART (Double Asteroid Redirection Test) impact. The small size of the asteroids and the unknowns about their main characteristics (shape, mass, density…) make it complex to determine a safe landing trajectory, which would ensure not to miss the secondary and not to have a high rebound that would put the cubesat back into orbit. Considering the uncertainties on the deployment mechanism, the variation of the ejection direction, landing velocity or landing location can be critical for the success of the technology demonstration and for ensuring a suitable location on the asteroid for the scientific investigations. In this study we present an analysis of the deployment strategy, and of the descent and landing trajectories. Low-Cost Command and Tracking for Testing of Atmospheric Probes. D. Handy, M. Atkinson, J. Hanson, M. Murbach, A. Zadehgol and the University of Idaho Near-Space Senior Design Team.

Introduction: The University of Idaho Near- devices on or off, and sending various other com- Space Engineering Program, known as the Vandal At- mands. mospheric Science Team (VAST), is made up of five Dial-Up Data: The University of Idaho near- teams: Mechanical/Aerodynamic Design, Science, space senior design team, working closely with NASA Imaging, Radio Communications, and Satellite Com- Ames Research Center, is focused on integrating an munications. VAST designs and builds projects to be Iridium 9523 modem into the high-altitude balloon flown to 80,000 – 100,000 feet underneath high- payload as a proof-of-concept for future atmospheric altitude weather balloons. The Satellite Communica- missions. In addition to using SBD, satellite modem tions team has the responsibility of developing soft- supports dial-up connections, a different method of ware, hardware, as well as procedures and protocols establishing and maintaining a communication link for handling all flight communications between the between the remote payload and a ground control sta- VAST Command Center on the ground and the VAST tion. balloon payload in flight, utilizing the Iridium satellite The communication begins with the modem mak- network and short burst data (SBD) messages. A senior ing a data call to the Iridium Ground Station (IGS), a design team at the University of Idaho is currently gateway to the internet. Once this connection has been working on adding dial-up data capability into the made, the IGS will route any HTTP packets received high-altitude balloons. from the modem on toward their destination, a server Iridium Satellite Network: The network of Iridi- running a program to store, parse, and display the data. um satellites provides truly global coverage of the This dial-up connection provides a constant link Earth. Orbiting 781 km above Earth, the constellation between the remote system and a server on the ground. is capable of providing two-way communication for a This is an improvement in terms of data rate and al- variety of Earth atmosphere and near-Earth probe ap- lowable packet size over SBD, and will allow for the plications. Many satellites in low Earth orbit, sounding transfer of more data, such as pictures from onboard rockets, and high-altitude balloon tests can take ad- cameras, than previously possible. The next generation vantage of the Iridium satellite network for communi- of Iridium satellites, currently being launched into or- cations. bit, will provide much higher data rates and enable Command and Tracking System: The balloon video downlink. payload contains a circuit board with an Arduino- Applications: The Iridium satellite network can compatible microcontroller connected to several sen- be leveraged as a low-cost command and tracking sys- sors measuring temperature and pressure, in addition to tem of any scientific payload in close proximity of a GPS receiver. This data is recorded to a microSD Earth. The platforms developed by the University of card continuously and uploaded at a set interval via Idaho’s VAST program and near-space senior design SBD through an onboard Iridium 9602 modem to the team have proved to be an effective and reliable means satellite network. The Iridium network delivers the of tracking a payload and commanding onboard devic- packets to a remote server where it is automatically es. Use of the Iridium satellite network enables the interpreted, parsed, and uploaded to an online dash- development and flight testing of low power, low board. This allows all data to be seen in real time, and mass, and low cost communications systems that can provides the ability to track and recover the payload be used on future small satellite Earth and planetary during and after its flight. The Satellite Communica- probe missions. tions team is responsible for every step of designing, developing, testing, and operating the Iridium hardware, developing communications protocols, as well as analyzing and interpreting all data received via the Iridium network. In addition to downlinking science, navigation, and engineering telemetry, the system is used to send a payload cutaway command for “range safety” operations, if necessary. The Satellite Communications team plans to expand its ability to send commands to the payload in flight, thereby offering greater control over onboard devices; changing operating modes, powering University of Idaho Near-Space Engineering Program: Engaging Students in Aerospace Engineering, Technol- ogy, and Science with Near-Space Flight. J. Hanson, D. Handy, C. Atkinson, C. Smith, W. Duncan, N. Wagner, B. Perley, M. Murbach, A. Zadehgol, and the University of Idaho Near-Space Engineering Team.

Introduction: The University of Idaho Near-Space surviving the extreme environments of near-space. Stu- Engineering Program, known as the Vandal Atmos- dents must think creatively, and the resulting projects pheric Science Team (VAST), is an undergraduate, in- demonstrate the resourcefulness and consideration of a terdisciplinary, student-led organization that launches a highly constrained, low-cost design environment. high-altitude balloon with a diverse payload of technol- Near-Space flight benefits students in its accessibil- ogies and experiments each semester. ity and learning potential, as well as industry aerospace The program consists of five teams, each focusing research in its utility as a testbed of incremental concept on a critical aspect of the VAST mission. These groups development. are: Aerodynamic & Mechanical Design, Science, Im- Student Engagement: VAST puts students first in aging, Radio Communications, and Satellite Communi- all facets of the program. Being student-led means all cations. Although the majority of students have engi- decisions, from scheduling and budgeting through flight neering backgrounds, this is not a requirement for join- execution, are generated by students. All members, re- ing a team and students are urged to get involved in gardless of experience, are given the freedom to suggest work that interests them regardless of relation to aca- new projects and to rethink and grow existing work. En- demic major. Teams interact heavily, as constant collab- suring students feel ownership of the program makes the oration is a necessary part of integrating in the fast- experience meaningful and encourages members to in- paced design environment. vest themselves. Program Goals: The Near-Space Engineering Pro- The program offers students the opportunity to gram aims to design and test small, low-cost aerospace learn technical and hands-on skills not always afforded technologies, develop experienced, critical-thinking them in a classroom lecture environment. Each flight creators and leaders, and foster partnerships between the cycle demands a design, testing, integration, and execu- University of Idaho and the aerospace industry. tion phase, and this necessitates learning to build off The primary goal of VAST is to enable undergrad- each stage and solve problems procedurally. uate students to experience the hands-on process of de- Collaboration: VAST collaborates with industry signing aerospace technologies, actualizing their de- partners, including NASA Ames Research Center, the signs, and validating them at altitudes up to 100,000 NASA Jet Propulsion Laboratory (JPL), and Pioneer feet. Members get to learn and execute the process of Aerospace, to gather data for emerging technologies for designing, building, testing, flying, and recovering en- use on future planetary probes. Current examples of gineering, science, and technology instrumentation. these partnerships include the development of a guided By involving students of all skill levels, VAST parafoil system for the Small-Payload Quick-Return aims to provide opportunities for experienced students project currently in development at NASA Ames, and to develop their teaching and leading skills, and for use of wireless-sensing technologies to collect para- young students to develop their technical skills and gain chute dynamics data with Pioneer Aerospace. design experience. Engaging with industry teams provides VAST with relevant problems that need solving and connections for internship opportunities, while providing its partners with low-cost near-space flight testing. Near-Space Flight: High-Altitude Balloons are a low-cost, rapid, and safe means for testing a variety of technologies in the extreme conditions of the Earth’s stratosphere. This low-temperature and low-pressure environment approximates the conditions near the Mar- tian surface, and thus serves as an ideal and accessible proving ground for validating planetary probe technologies. VAST members are constrained in the same manner as professional engineers working in small satellite and probe design programs. Flight payloads are necessarily small, lightweight, low-power, and capable of STUDY ON THE MARTIAN EXPLORATION PROBE USING A PARAFOIL-TYPE VEHICLE. T. Mori- yoshi1, H. Kanemaru1, H. Nagano2, K. Yamada3 and H. Nishida1 1Tokyo University of Agriculture and Technology (2-24-16 Naka-cho, Koganei-shi, Tokyo, Japan, ju- [email protected] ) , 2Waseda University, 3JAXA

Introduction: A parafoil type-vehicle have been from altitude of 10 km to the Mars surface. The flight developed for the various purposes and actually used speed to horizontal direction and vertical one are in various field in the past, for example, sky sports, 59.5m/s and 10.8m/s. These results indicated that this cargo transfer and space transfer system. Especially, parafoil type probe have a potential for new Mars ex- this parafoil-type vehicle has a significant advantage ploration. for the space usage, because of its good packing effi- References: [1] K. Yamada, K. Suzuki, D. Akita, ciency. Therefore we proposed the Martian Explora- O. Imamura, Y. Nagata, Y. Takahashi, K. Hiraki, and tion Probe using a parafoil-type vehicle that is frying in S. Higashino. (2013) Innovative Aerodynamic Tech- the Martian atmosphere shown in Fig. 1. The packing niques using Inflatable Structure for Future Martian efficiency of parafoil makes it possible to store that in EDL Mission, APISAT limited volume of a launch vehicle and an entry capsule. In additionally, such type probe can transport relatively heavy observation equipment because a parafoil can realize large wing area. A parafoil-type vehicle is generally known that is easy and simple to control and guidance. Therefore the probe can be landing into deep valley or on the mountainside, which had never been observed in past mission. Innovative parafoil: However, usually parafoil which inflated by ram air might be not sustained wing shape in low density atmosphere, like the Mars. There- fore an innovative parafoil which is a partial closed type parafoil was developed for the Martian probe, Fig. 1 Conceptual image of the Martian Exploration shown fig. 2. The parafoil consists of inflatable spar, Probe using a parafoil-type vehicle rib of airfoil section and skin of thin film. The concept is an intermediate idea between a conventional rigid wing and an inflatable wing. The structure has some advantages that configure the accurate airfoil shape across the full span of the wing and realize an active and quick deployment of the wing by small gas injected. The wind tunnel tests using rigid arched subscale models were carried out to investigate aerodynamic characteristics of the parafoil.(fig. 3). The measurement result of these tests is shown in Fig. 4. The peak of the lift to drag was 12.8 (CL: 0.25, CD: 0.02) at 4 degree angle of attack. In additionally, the wind tunnel Fig. 2 Picture of partial Fig. 3 Picture of wind tun- tests using partial closed type parafoils were also car- closed type parafoil. nel test of rigid arch shape ried out and the lift to drag measured 5.8. This result is subscale model. consistent because the result was including line drag and attitude motion. Concept design: Preliminary concept design of the parafoil-type probe was conducted based on wind tunnel test results. Design constraints are the transport payload mass of 5.0 kg and chord length of wing of 0.9m. The design constraints and the probe specifications as a results of conceptual study are listed table 1. The probe mass was 5.5kg. The probe would have a flight time of about 15minutes, if the vehicle glide

Fig. 4 Measurement result of wind tunnel test of subscale model.

Table 1 Concept design resulte of parafoil type Mar- tian probe.

Design constraint vehicle spesificaition Payload mass 0.49 5.0 Wing mass [kg] [kg] 0.20.3 Payload size Span length [m] 2.7 0.1 Chord length 0.9 Wing area[m2] 2.43 [m] Area density System lift to 0.2 5.5 of parafoil drag Aspect ratio 3 Lift coefficient 0.34 Gravity ac- cerelation 3.2 Drag coefficient 0.07 [m/s2] Atmosphere Glide speed density 0.01 60.5 [m/s] [kg/m3]

SLP: A LANGMUIR PROBE INSTRUMENT ON BOARD A CUBESAT. S. Ranvier1, F. Cipriani2, Andreas Waets2, M. Anciaux1, E. Gamby1, P. Cardoen1, S. Bonnewijn1, J. De Keyser1, D. Pieroux1 and J.P. Lebreton3, 1Royal Belgian Institute for Space Aeronomy (BIRA-IASB), Avenue Circulaire 3, 1180 Brussels, Belgium, Email: [email protected], 2ESA, TEC-EPS (Space Environment and Effects Section), Keplerlaan 1, Noordwijk 2200AG, The Netherlands, 3Laboratoire de Physique et Chimie de l’Environnement et de l’Espace (LPC2E), 3A, Avenue de la Recherche Scientifique, 45071 Orléans cedex 2, France.

Abstract: A Langmuir probe instrument, which where the density of the ambient thermal plasma is will fly on board the Pico-Satellite for Atmospheric rather low, high energetic beams can charge the space- and Space Science Observations (PICASSO), is under craft to dangerously high potential. To assess this risk, development at the Royal Belgian Institute for Space particle-in-cell (PIC) simulations have been performed Aeronomy. PICASSO, an ESA in-orbit demonstrator, with SPIS (Spacecraft Plasma Interaction System). The is a triple unit CubeSat of dimensions 340.5x100x100 results of those simulations together with results of the mm. test of SLP in a plasma chamber will be presented. The sweeping Langmuir probe (SLP) instrument, which includes four thin cylindrical probes whose electrical potential is swept, is designed to measure both plasma density and electron temperature at an altitude between 500 km up to 600 km from a high inclination orbit. Therefore, the plasma density is expected to fluctuate over a wide range, from about 1e8/m³ at high latitude and high altitude up to several times 1e12/m³ at low/mid latitude and low altitude. The electron temperature is expected to lie between approximately 600 K and 10.000 K. Given the high inclination of the orbit, the SLP instrument will allow a global monitoring of the ionosphere with a maximum spatial resolution of the order of 150 m for the electron density and temperature, and up to a few meters for electron density only. The main goals are to study 1) the ionosphere- plasmasphere coupling, 2) the subauroral ionosphere and corresponding magnetospheric features, 3) auroral structures, 4) polar caps, 5) for the density, the multiscale behaviour, spectral properties and turbulence of processes typical for the auroral regions, and 6) ionospheric dynamics via coordinated observations with EISCAT’s heating radar. An important issue implied by the use of a pico-satellite platform for a Langmuir probe instrument is the limited conducting area of the spacecraft which can lead to spacecraft charging. In order to avoid this problem, a specific measurement technique that includes the simultaneous measurement of the potential and current of different probes, has been developed to retrieve consistent current-voltage characteristics that can be used to estimate the plasma parameters mentioned above. Another spacecraft charging effect, which is not due to the Langmuir probe instrument but to the plasma environment, can also alter the results of SLP and, moreover, can lead to the failure of the instrument and can be a threat for the spacecraft itself. When PICASSO will fly above the polar regions, Cold Gas Thruster Attitude Control System to Augment the Exo-Brake Passive Deorbit. R. Rosila1, S. Torres2, L. Murray3, I. Supelana4, and J. Austring5, 1San Jose State University, San Jose CA, 95192 rob- [email protected] , 2San Jose State University, San Jose CA, 95192, 3San Jose State University, San Jose CA, 95192, 4San Jose State University, San Jose CA, 95192, 5San Jose State University, San Jose CA, 95192

With increasing miniaturization and capability of microsatellites, experimentation with micro-propulsive technologies for attitude control is becoming more important. The ongoing TechEdSat mission series has demonstrated the feasibility of a passive deorbiting technique utilizing its drag modulation technology, the Exo-Brake. Despite the successes of the Exo-Brake, the system lacks a capability for the precision attitude control necessary for uni-directional antenna usage. Leveraging on current cold gas thruster technologies, the SOAR Sat I CubeSat system utilizes a reaction control system in conjunction with the Exo-Brake to deliver precision attitude control and improved targeting capabilities during a passive deorbit. The following work provides an in depth explanation of the analysis, design, and implementation of the attitude determination and control system (ADCS) utilized on SOAR Sat I. An analysis of the design and integration of vital supporting subsystems is included. System validation and experimental data is evaluated to support the effi- cacy of the system.

A CUBESAT-BASED ALTERNATIVE FOR THE JUNO MISSION TO JUPITER. P. H. Stakem1, R. Santos Valente Da Costa2, A. Rezende3, A. Ravazzi3, and V. Chandrasenan4, 1Johns Hopkins University/Capitol Technolo- gy University, 2Universidada Federal do Rio Grande do Sul/Johns Hopkins University/Capitol Technology Universi- ty, 3Federal University of Uberlandia/Capitol Technology University, 4University of Maryland-College Park.

Introduction: This paper discusses the design of a strawman InterPlanetary Cubesat Mission based on the parameters of the ongoing Juno Mission to Jupiter. That mission put a large spacecraft into Jupiter orbit. The approach presented here has a quantity of Cu- besats as the primary payload. There is a large “Moth- ership” which enters Jovian orbit, and dispenses various Cubesats, acting as a store-and-forward communications relay back to Earth. The number of Cubesats is determined by the outlines (size, mass, power) of the mothership. We baselined the comparable numbers for the Juno spacecraft. With this scenario, we can include 333 3U Cubesats, with a large number of different instruments and sensors. How the various spacecraft interact on this mission will be outlined in the paper. In addition, we examine all technical issues of the mission, and address their Technology Readiness Levels. We include intras-swarm communications, and communications back to Earth. The individual Cubesats can organize into a “swarm of convenience,” to address simultaneous observation, and can form a compute cluster. A key concept is to use a sharable database across all unit. Contain Electronic Data Sheets. The individual Cubesats will be checked before being deployed, and failed or malfunctioning units can be discarded. In addition, we apply the concept of Rad-Hard software to counter the more intense radiation environment expected. We are also examining a similar approach to the exploration of the gas giants, Uranus and Neptune, based on the same concept. We believe this approach will enhance the science payback from these missions, and introduce large levels of redundancy and flexibility, while reducing costs. OVERVIEW OF ESA STUDIES ON INTERPLANETARY CUBESAT MISSION CONCEPTS. R. Walker1, 1ESA ESTEC, Noordwijk, The Netherlands.

Introduction: The presentation provides a summary of the recent studies that have been performed in the ESA General Studies Programme on interplanetary CubeSat mission and system concepts. A number of studies have involved mother-daughter system architectures where the CubeSats are carried to a target destination such as lunar orbit or to a Near-Earth Object (NEO) on a larger spacecraft and deployed at the target in order to fulfil their mission. In this case, the technical challenges of propulsion, long-range communication and deep space environment survivability are highly alleviated, since the host spacecraft provides resources and accommodation during cruise and communications to Earth ground stations in conjunction with Inter-Satellite links locally with the deployed Cu- beSats. Additionally, ESA has also recently performed a study in the ESTEC Concurrent Design Facility on a stand-alone deep space CubeSat system capable of rendezvous and characterisation of NEOs or transfer to Sun-Earth L5 lagrange point for space weather measurements, based on piggyback launch opportunities to near Earth escape (e.g. astronomy missions to L2, lunar exploration missions, outer planet missions with Venus flyby). Such a system relies upon a number of miniaturised technologies currently in development, and once launched is expected to reduce the entry-level cost of deep space exploration by an order of magnitude. The presentation will outline the mission and system designs, as well as the innovative technologies enabling their realisation in the future.

Flight Experiment of Nano-satellite “EGG” for Deployment Demonstration of Membrane Aeroshell

Kazuhiko Yamada1, Takahiro Moriyoshi2, Kazushige Matsumaru3, Hiroki Kanemaru2, Takahiro Araya4, Koji- ro Suzuki5, Osamu Imamura3, Daisuke Akita6, Yasunori Nagata7, Yasuhiro Shoji8, Yusuke Takahashi9, Ya- sumasa Watanabe5, Takashi Abe1 and MAAC group

1JAXA(3-1-1 yoshinodai chuo-ku Sagamihara, kanagawa Japan, [email protected]), 2Tokyo Univer- sity of Agriculture and Technology, 3Nihon University, 4Tokyo University of Science, 5The University of Tokyo, 6Tokyo Institute of Technology, 7Okayama University, 8 Osaka University, 9 Hokkaido University.

Introduction: layer of inflatable ring. The polyimide file bonding by The deployable and flexible aeroshell is promising as silicon rubber adhesion is used for gas tight layer of a future atmospheric-entry system for space transporta- inflatable ring. The total mass of aeroshell is 456g. tions and planetary explanations. The atmospheric en- After the EGG is deployed from ISS by J-SSOD (JEM try vehicle with a large and light flexible aeroshell Small Satellite Orbital Deployer), four solar panels make its ballistic coefficient low and can reduce the surrounding the main body are deployed and inflation aerodynamic heating during the atmospheric entry. Our gas is injected into the inflatable ring to sustain the group has been researched and developed on this tech- aeroshell. The EGG has two iridium SBD modules and nology to apply it to actual missions since 2000, focus- will communicate a ground station via iridium satellite ing on a flare-type thin membrane aeroshell supported network anywhere in LEO. This function may enables by a single inflatable ring. Two balloon drop tests were data acquisition even in re-entry phase. While the EGG carried out in 2004 and 2009 and a re-entry demonstra- is one of the demonstrators for the innovative re-entry tion from an altitude of 150km was carried out 2012 system using the deployable aeroshell, the EGG was using the S-310 sounding rocket[1]. As a next step of designed to burn out during the re-entry in this mission development, a flight demonstration in low earth orbit from the viewpoint of the ground safety against falling was planned utilizing an opportunity to deploy a nano- objects. satellite from ISS. We developed the nano-satellite with the deployable aeroshell called as EGG (re-Entry Quick Report of Flight Experiment: satellite with Gossamer aeroshell and Gps/iridium). EGG was deployed from ISS on January 16, 2017 The objectives of EGG are to deploy the inflatable (JST) and the soundness of the satellite was confirmed aeroshell under the high vacuum and micro-gravity by the telemetry data via Iridium network. And the condition, the function of positioning and communi- aeroshell deployment demonstration was successfully cating system using a commercial satellite network. carried out in February 11 in 2017. According to the The EGG was carried to ISS by HIIB and HTV in De- telemetry data, it is confirmed that the pressure in in- cember of 2016 and was deployed from ISS to LEO in flatable ring increases and was remained with expected January of 2017. The conceptual figure of EGG mis- leak rate. The Figure 3 is the picture of deployed aero- sion sequence is shown in Fig.1. After the deployment shell captured by onboard camera installed in main from ISS, the function check and initial setting is car- body. This picture also shows the successful deploy- ried out at first. After that, the membrane aeroshell is ment of the membrane aeroshell and the solar array deployed. The orbital decay occurs due to aerodynamic panels. In March of 2017, EGG flies on the Low Earth force acting on the deployed aeroshell. Finally, the Orbit with an altitude from 390km and 400km. The EGG reenters into atmosphere and burns out. EGG is gradually decreasing its flight altitude due to the aerodynamic force and will finally reenter the at- EGG Flight Model: mosphere to burn out. In presentation, the development The flight model of EGG is shown in Fig. 2. The left and verification test for the flight of EGG and quick figure is a launch configuration of EGG which is 3U- results including re-entry phase are reported. size (10cm x 10cm x 30cm, 4kg) with the packed aeroshell. The right figure shows the EGG with deployed Reference flexible aeroshell with a diameter of 80cm. This aero- [1] Yamada, K., Nagata, Y., Abe, T., Suzuki, K., Imamu- shell has a ring-shape inflatable structure to deploy and ra, O., and Akita,D.: Suborbital Reentry Demonstration of sustain a thin membrane flare. This aeroshell is made Inflatable Flare-Type Thin-Membrane Aeroshell Using a of same materials used for actual future probe. The thin Sounding Rocket, AIAA Journal of Spacecraft and Rock- textile made of ZYLON is used for flare part and outer ets, January, Vol. 52, No. 1(2015) : pp. 275-284

Fig.1. Conceptial image of mission sequence of EGG

Fig.2. Flight model of EGG in packed and deployed configuration

Fig.3. Deployed aeroshell image captured by onboard camera in LEO. (Left: Camera position and line of sight, Right : Captured image)

Semi-Hard Landing and Shock Absorption Mechanism of OMOTENASHI Project. J. Kikuchi1, T. Hashimoto2, T. Yamada3, M. Otsuki4 and T. Ikenaga5, 1Japan Aerospace Exploration Agency (JAXA) (E-mail : kiku- [email protected], Address : 3-1-1 Yoshinodai, Chuo, Sagamihara, Kanagawa, 252-5210, Japan), 2345Japan Aero- space Exploration Agency (JAXA)

Introduction: OMOTENASHI (Outstanding surface. At this point, the spacecraft starts the 3Hz ro- MOon exploration TEchnologies demonstrated by tation to get the spin stabilization. After spin stabilized, NAno Semi-Hard Impactor) will be the world’s small- two airbag for a shock absorption are inflated on both est moon lander. It will be launched by NASA’s SLS sides of OM. 5, by igniting RM with Laser Diode, RM (Space Launch System) with the Orion spaceship in and SP are decelerated and separated from OM as del- 2018. In the near future, industry, academia, and even ta-V2. 6. by RM burning for 20 seconds, RM along individuals will be able to easily participate in space with SP will decelerate to ±30 m/s. 7. SP is separated exploration. The technology demonstration of from RM which will simultaneously start the free fall OMOTENASHI will contribute to the realization of to the moon surface. such a world. It will also takes measurements of the As the payload of this spacecraft, a radiation meas- radiation environment of Earth and moon region. The urement monitor is mounted on the OM to transmit the observation will be helpful for future exploration. result of the radiation environment between the earth and the moon. On the other hand, the accelerometer is Mission Overview: OMOTENASHI spacecraft is mounted on SP to get the landing signal as well as a 14kg, 6U size cubesat which is composed of 3 mod- shock absorption capability of the airbag. ules as followings; Orbiting Module(OM):7kg, Rocket Motor(RM):6kg, Surface Probe(SP):1kg Mechanical Design: The most critical part for the The mission sequence is described as Figure.1. Ba- mission success is the deceleration by the rocket motor. sically, two burns (delta-V1 and delta-V2) are required The deceleration failure can be considered for the fol- to land on the moon surface. 1-3, after the separation lowing reasons; 1. Attitude error of the rocket motor, 2. from SLS, the spacecraft trajectory is controlled by the Ignition / separation timing deviation. Therefore, sev- gas jet then maneuvered to enter the lunar impact orbit eral unique mechanism are designed in this mission. as delta-V1. 4, the spacecraft velocity is estimated as For the reduction of motor vibration, the separation rail 2500m/s when approaching to 1km from the moon is investigated by interfereing a leaf spring and a teflon

Figure.1 OMOTENASHI Mission Overview which is a low friction rate between OM and RM. To Shock Absorption Mechanism: The prototype of reduce the timing deviation, the rocket motor is ignited the airbag system which is selected in this mission is by a laser diode via optical fiber and lens. This opera- described as Figure.3. This airbag is inflated to a barrel tional timiming of the laser diode is estimated as 30ms. shape using nitrogen gas within 24 kPa. This internal On the other hand, the separation mechanism is de- pressure is adjusted by the relief valve anytime. This signed to use a disk spring and a pinpuller which is airbag system consists of the airbag surface membrane NEA (operational time:35ms). Currently, these me- Zylon(external) including Polyimide film(internal). chanical designs are developed and experimented by The zylon has strong strength and heat resistance. On using a vacuum chamber. the other hand, the polyimide film can keep the higher airtightness of inside. It is difficult to evaluate the im- Landing Attitude Analysis: Attitude analysis of pact absorption performance of the airbag on the the rocket motor is very important for the mission suc- ground. Therefore, crash experiment was carried out in cess as well. Therefore, a dynamic analysis simulation February of 2017 as Figure.4. In this experiment, the (Figure.2) is carried out including the following errors; airbag is fixed collided by a high speed moving a motor vibration, a nutation, and an ignition / separa- truck(30m/s). The experiment result is described as tion timing deviation. The feasibility of this mission is shown in Figure.5. Based on these results, the system investigated from this analysis result. In addition, the design is promoted on top of being reflected for the motion behavior of SP after landing on the moon sur- higher feasibility. face is also analyzed. In this analysis model, the lunar regolith is modeled as multi-particle to simulate a high- References: [1] Sandy, C, Cadogan, D, Grahne, er accurate of the landing behavior. The parameter of M, “Development and Evaluation of the Mars Path- the lunar regolith, such as filling and friction rate, can finder Inflatable Airbag Landing System”, Paper IAF- be changed freely. 98-I.6.02

Figure.2 Attitude Analysis of Deceleration Figure.4 Crash Experiment of Airbag

Figure.3 Prototype of Airbag Figure.5 Acceleration Result of Crash Experiment

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Introduction: Every mission that has successfully configuration architectures during trade-study-level landed a payload on Mars has utilized heritage decel- investigations. The first step in the high-level method- eration technologies from the 1960’s and 1970’s Vi- ology examines the potential far-field flight envelope king era. Utilizing only Viking heritage deceleration of each piece of tumbling debris after ejection. The technologies presently available, it is estimated that a cumulative far-field flight envelope of all pieces of 1.1 mT payload landed at 0 km elevation represents the ejected debris is termed the “debris field envelope”. upper limit of current Entry, Descent, and Landing The debris field envelope is compared to the post- (EDL) capabilities [1]. The succession from the current ejection flight trajectory of the primary descent vehicle state of the art along NASA’s goal of extending and under propulsive deceleration from SRP. An initial sustaining human presence in our solar system will downrange offset distance between the primary vehicle require landing large robotic (~10 mT) and human and the point at which the debris begins to tumble is class payloads (~40-80 mT) on Mars with landed accu- calculated such that the descent vehicle trajectory will racies on the order of meters. Supersonic Retropropul- not pass through the debris field envelope. sion (SRP) is one promising candidate supersonic de- The second step of the high-level methodology uti- celeration technology currently under heavy develop- lizes the calculated offset distance and a three-step ment by both NASA and SpaceX to enable higher sequential approximation methodology (TSSAM) to mass Mars missions. prescribe a near-field “transit” trajectory for each piece To enable the use of SRP, an entry vehicle will of ejected debris. The first step of the TSSAM assumes likely need to perform a supersonic vehicle reconfigu- no aerodynamic forces are acting on any piece of de- ration during descent to the Martian surface to expose bris and determines the initial separation impulse force SRP rocket nozzles into the oncoming atmospheric that would be required to propel each piece of ejected flow. The change between the hypersonic entry vehicle debris away from the primary descent vehicle by a configuration and the SRP-ready vehicle configuration distance equal to calculated offset distance. In the se- will require the supersonic ejection of the vehicle aero- cond step of the TSSAM, the transit trajectory for each shell. Once ejected, the discarded aeroshell becomes an piece of debris is optimized to minimize the required intact, solid-mass piece of debris traveling in the same initial separation impulse forces as well as the continu- direction as the primary vehicle. This debris poses po- ous aerodynamic moment control torques necessary to tential catastrophic recontact risks to the primary de- stabilize the debris during the transit trajectory. The scent vehicle. Mitigating these debris recontact risks is optimization of the transit trajectory utilizes isolated a significant hurtle to the development of SRP as a aerodynamics for all pieces of debris. mission-ready technology. In the final step of the TSSAM, interference aero- Supersonic vehicle reconfigurations have never dynamics are calculated along the optimized transit been performed and there exists no published research trajectory from the previous step. The interference aer- in this field. However, supersonic ascent vehicle re- odynamics are not used to modify the transit trajectory. configurations are routinely performed and are well Rather, they are used to obtain a better approximation documented in the literature. The most well-known of the aerodynamic forces and moments experienced example of an ascent supersonic vehicle reconfigura- by the debris as it traverses the transit trajectory. Re- tion comes from the Space Shuttle ejecting its spent sponse Surface Methodology (RSM) is used to create solid rocket boosters on its ascent into orbit. While models of the difference between isolated and interfer- fundamental differences exist between supersonic de- ence aerodynamic coefficients along this transit trajec- scent and ascent vehicle reconfigurations, the two dis- tory. By modeling this difference, investigators may ciplines share some of the same challenges. These sim- determine the point at which interference aerodynamic ilarities allow for ascent analysis methodologies to be coefficients converge to their isolated values. Because adapted and modified for use with descent analyses. interference aerodynamics are exponentially more ex- This poster presents current efforts to develop the pensive to compute than isolated aerodynamics, signif- first supersonic descent vehicle reconfiguration rapid icant resource savings can be realized by only compu- analysis methodology. An overview is presented of a ting interference data where absolutely necessary. high-level, rapid analysis methodology that provides This step is one of the key features of the separa- mission designers the capability to assess the initial tion analysis methodology that enables its use for rap- feasibility of numerous candidate descent vehicle reid, high-level analysis. The substantial time and expense required to calculate interference aerodynamics precludes the creation of a full aerodynamic database due to the high dimensionality of the coupled dynamic motion inherent to supersonic vehicle reconfigurations. By calculating expensive interference aerodynamics only along the transit trajectory, we reduce the time and expense required to perform vehicle reconfiguration analysis while still gaining valuable insight that can be used to competitively compare multiple candidate vehicle reconfiguration architectures. This methodology quantitatively approximates required debris ejection impulses and debris transit stability control moments for a descent vehicle reconfiguration architecture. Results can be rapidly computed for several candidate vehicle reconfiguration architectures. The methodology output metrics can be used to quantitatively determine the fittest candidate(s) for further low-level, detailed analysis. [1] Dwyer Cianciolo, A.M., et al., (2010) Entry, Descent and Landing Systems Analysis Study: Phase 1 Report.

TAKING ON THE MANTLE OF VITAL WITH A NEW TENSGRITY ACTUATED TESSERA LANDER. K. Schroeder1, J. Bayandor1, J. Samareh2, 1CRashwothiness for Aerospace Structures and Hybrids (CRASH) Lab, Virginia Tech, Blacksburg, VA 24061 USA, 2NASA Langley Research Center, Hampton, VA 23666 USA.

Abstract: The Decadal Survey has stated, “There is a The unique multifunctional infrastructure is a critical future role for additional VISE-like missions to critical feature that sets TANDEM apart from the a variety of important sites, such as Tessera terrain.” [1] current state of the art but, on top of the concept’s The concept developed for the Decadal Survey to functionality, it has displayed a significant mass saving explore this region was the VITaL mission [2]. compared to other available concepts. The predicted However, the study showed, first, that the entry loads mass of the mission is reduced by 38% (approximately were too high for some of the onboard scientific 800 kg) after switching from the ADEPT-VITaL [4] instruments, and second, that the Tessera region can be design to the TANDEM design. Of this mass savings, “viewed as largely inaccessible for landed science due only about 100 kg was contributed by the decrease in to their known roughness.” [1] heatshield diameter (in order to maintain the same The VITaL lander used a variation of the ballistic coefficient as ADEPT). The change of lander heritage design from the Venera landers. However, for design resulted in a predicted mass reduction of 190 kg. landing in the Tessera regions additional mass is The remaining 510 kg of mass reduction is a direct required to maintain a low center of gravity and prevent result of TANDEM’s most unique attribute, namely the the lander from tipping over. Despite the very low integration of the landing and locomotion systems into center of gravity of the VITaL lander, it was decided the entry vehicle design. that a Venera-class design does not adequately remove Thus, TANDEM’s multifunctional the risk associated with landing in the rough terrain [1]. infrastructure presents a robust and lightweight system TANDEM [3] provides the same capabilities to address some of the Decadal Surveys most pressing that were proposed for the VITaL mission while questions for Venus. eliminating the identified risks associated with entry Bibliography: [1] Squyres, S., and et al., Vision and and landing by utilizing a multifunctional tensegrity Voyages for Planetary Science in the Decade 2013- infrastructure. Before entry, the tensegrity structure is 2022, National Academies Press, 2012. [2] Gilmore, used to deploy low ballistic coefficient heat shield, thus M., and Glaze, L., “Venus Intrepid Tessera Lander,” reducing the high entry loads. On touchdown, the 2010. [3] Schroeder, K., Bayandor, J., and Samareh, J., tensegrity structure provides omnidirectional impact “Venus Lander Vehicle Concept Development Part 2: protection at high velocities. Figure 1 shows the TANDEM, Tension Adjustable Network for Deploying TANDEM vehicle impacting a flat rigid surface at its Entry Membrane,” NASA TM Under Review, 2016. [4] terminal velocity (~25m/s). This capability removes the Smith, B., Venkatapathy, E., Wercinski, P., Yount, B., risk inherent to landing on the uneven and steep landing Prabhu, D., Gage, P., Glaze, L., and Baker, C., “Venus environments of the Tessera terrain. Furthermore, the In Situ Explorer Mission Design Using a Mechanically multifunctional infrastructure provides the additional Deployed Aerodynamic Decelerator,” 2013 IEEE benefit of surface locomotion at little to no extra mass Aerospace Conference, IEEE, 2013, pp. 1–18. penalty.

Figure 1 Von Mises stress on compression members for the impact at terminal velocity ANALYSIS, TEST AND SIMULATION OF LANDING SYSTEM TOUCHDOWN DYNAMICS S. Schröder1 and L.Witte2, 1DLR - German Aerospace Center (Robert-Hooke-Str. 7, 28359 Bremen, Germany, [email protected]), 2DLR - German Aerospace (Robert-Hooke-Str. 7, 28359 Bremen, Germany, [email protected])

Introduction: Existing studies for future exploration missions pose demanding requirements towards access by vehicles to scientifically interesting sites on surfaces of extraterrestrial bodies with lower gravity than on earth. This applies to the touchdown of landing vehicles as well as to the surface operations of roving vehicles. In order to support the design, development and operation of those vehicles on an experimental basis, the DLR Institute of Space Systems has developed the Landing and Mobility Test Facility (LAMA). The rationale of this facility is to provide a test bed to study full scale vehicle-soilinteractions, like tip-over stability of landing vehicles or terrain accommodation of rovers, in a reduced gravity environment by weight offloading with an industrial robot system, but same inertia mass. The advantage of the weight reduction is the similarity of lander body’s geometry, mass distribution and local gravity comparing to the real environment. In addition, the department of Landing & Explora- tion Technology of the DLR Institute of Space Sys- tems is also focusing on the development and verification of analytical methods for the investigation of the touchdown dynamics of landing system and its capabilities for embedding into the landing site assessment. This poster outlines the test facility, simulation and analysis tools developed by the department and their use for recent landing missions. Comparison of Crater-Detection Algorithms for Terrain-Relative Navigation. A.S. Moreno Gonzalez1 , I. El- Hajj1, J.W.F. Mes2, M. Henkel1 , R.S.D Autar1, R.A. Klavers1 and S. Woicke3, 1Bachelor student, Delft University of Technology, Delft, The Netherlands, 2Bachelor student, Leiden University, Leiden, The Netherlands, 3 PhD candidate, Delft University of Technology, Klyverweg 1, 2629HS Delft, [email protected]

Introduction: Precise landings on other bodies re- Algorithm 4 (A4): Simmilar to A1, craters are de- quire more than just dead reckoning using an inertial tected by edge detection. In this algotiyhm by multiple measurement unit on-board of the lander. If an inertial use of the Canny edge detection method with small and localisation of the lander with respect to a planetary large standard deviations. These edges are then paired surface is desired, so called crater detection (CD) and up based on parameters such as their length, orienta- crater-matching algorithms might be a valuable asset to tion and distance from one another, to form crater can- find the inertial position of the landing vehicle. This didates. And ellipse fitting function based on the di- would enable landing close to an inertially defined rect-least-squares method is then applied to the candi- landing site, which could for example, be a surface dates This method is based on [1]. asset of a previous mission. With the desire to reduce Algorithm 5 (A5): Craters are extracted by search- the landing ellipse size for future mission, more precise ing the image for lit areas. By using a predefined knowledge of the inertial state of the lander is required threshold on the grayscale values of the pixels, the for many of the next-generation missions. input image is filtered, resulting in an image with only Background: Within the scope of an eight-week the lit areas of in the image. By treating the lit pixels as student project at Delft University of Technology a points they are clustered together indicating the crater group of six bachelor students assessed, implemented positions in the input image. By finding the farthest and compared six different techniques to extract craters points in a certain cluster, the diameter can be derived, from lander imagery. These extracted craters could which is also used to derive the centerpoint of the then be used for vehicle localisation using so-called crater. This algorithm is based on [5] and [6] terrain relative navigation (TRN). Algorithm 6 (A6): CD is performed by identifiying Algorithms: The followings algorithms were im- contrast areas caused by global illumination, using the plemented. MSER blob detection technique. Matches of areas be- Algorithm 1 (A1): Craters are detected by The Edge longing to the same crater are found based on their & Image Gradient Determination (E&IG-D). The relative size and distance as well as their deviation method is based on [1]. E&IG-D is an edge based CD from the global illumination angle. Finally, an ellipse algorithm supplemented with image gradients and ge- can be fitted to these pairs. The algorithm is based on ometric tests to classify and pair edges. On each of the work presented in [7]. these pairs Direct Least Squares Fitting of Ellipses [2] Figure 1 shows the outputs of all six algorithms for is applied, these ellipses are then evaluated using their the same, simple, lunar analogue surface image. It can elongation and fitting error, if passed, the ellipse is said be seen that all algorithms are able to sucesfully detect to be 'a crater detection’. most craters. Please note, that A3 does not calculate the Algorithm 2 (A2): Craters are detected by pairing crater radius and only identifieds crater midpoints. bright and dark regions of craters of similar size and those which are in the proximity of one other. To check the pairs, a binary classifier is used: support vector machine. The task of the classifier is to decrease the false detection rate. [3]. Algorithm 3 (A3): CD is performed using template matching. Template Matching with Center of Mass Averaging (TM-CMA) is a simplified variant of Con- tinuously Scalable Template Matching (CSTM) [4]. Both algorithms use the correlation between a template of a crater and the image. However, TM-CMA only uses a single template instead of a scaled template family and uses both areas of high and low correlation to Figure 1: Results of the different algorithms. determine the center of the crater. Although TM-CMA does not provide information on the diameter of the crater, it is much faster than CSTM. Trade-off process: To compare the different algo- H.Z. (2016) Crater detection via convolutional neural rithms trade criteria have been established. The follow- networks. 47th Lunar and Planetary Science Confer- ing criteria are assessed: ence.[7] B. Maass, H. Kruger, S. Theil. (2011), An 1) Detection and false detection rates: even tough edge-free, scale-, pose-and ¨ illumination-invariant TRN does not require a 100% detection rate, this is approach to crater detection for spacecraft navigation, still an important criterion. 2) Accuracy: as the refer- 7th International Symposium on Image and Signal ence maps usually have rather large resolutions, inac- Processing and Analysis (ISPA), IEEE pp. 603–608 curacies of just a few pixels can cause large errors. 3) Robustness: how is the algorithm’s response to non- perfect craters as well as surface and image noise? 4) Run-time: the algorithm should be fast, as it should run on-board. Conclusions: In general, it was found that all algorithms are capable of performing the task of extracting sufficient craters for localising the landing vehicle with respect to a surface map, Figure 2 shows a very simple example of such a localization using Algorithm 5. However, from the six implemented and tested algorithms it was found that Algorithm 3 was most robust, while Algorithm 5 and 6 had the highest detection rate. Algorithm 5 has the lowest false detection rates. The runtime is lowest when using Algorithm 3 and 6.

Figure 2: Simple localisation task based on algorithm 5

References: [1] Cheng, Y., Johnson, A.E., Matthies, L.H. and Olson, C.F., (2003). Optical landmark detection for spacecraft navigation. Proceedings of the 13th Annual AAS/AIAA Space Flight Mechanics Meeting. [2]Fitzgibbon, A., Pilu, M. and Fisher, R.B., (1999) Direct least square fitting of ellipses. [3] Pedrosa, M. M., Pina, P., Machado, M., Bandeira, L., & da Silva, E. A. (2015). Crater Detection in Multi-ring Basins of Mercury. Iberian Conference on Pattern Recognition and Image Analysis [4] Burl, M.C., Fowlkes, C. and Roden, J. (1999) Mining for image content. Systemics, cybernetics, and informatics/information systems: analysis and synthesis. [5]Tian H. Yu M., Cui H. (2013) A new approach based on crater detection and matching for visual navigation in planetary landing. Technical report. [6] Lu T. Ding W. Cohen J.P., Lo Mars 2020 Parachute Workmanship Wind Tunnel Testing. Gregorio Villar1, Christopher Tanner1, Ian Clark1, Allen Chen1, 1Jet Propulsion Laboratory, California Institute of Technology

Abstract – The Low Density Supersonic Decelera- tor project revealed shortcomings in the conservatism of our parachute analysis and test methods and in our understanding of supersonic parachute inflations. For this reason, the Mars 2020 project is currently pursuing two types of parachutes: a Mars Science Laboratory build-to-print and a strengthened parachute. This includes fabricating flight-lots of both parachute types and subjecting them to workmanship wind tunnel testing, high altitude supersonic sounding rocket testing, and eventually selecting a parachute for flight. The objective of the workmanship wind tunnel testing is to verify that the flight-lot of parachutes (build-to-print and strengthened) have been properly built, prior to the sounding rocket test campaign in the Fall of 2017. The workmanship wind tunnel testing will take place at the National Full Scale Aerodynamics Complex, located at NASA Ames Research Center, and will be very similar to the Mars Science Laboratory Parachute Develop- ment Test Program [1]. For each type of parachute (build-to-print and strengthened), one flight-lot unit must survive at least one mortar deployment followed by ten sleeve deployments. This paper will discuss the test objectives and infrastructure of the workmanship wind tunnel testing, which is currently scheduled in the Summer of 2017.

References:

[1] Douglas Adams. and Nicholas Onufer. Mars Science Laboratory Parachute Development Test Pro- gram, AIAA 2011-2508. OVERVIEW OF THE ADVANCED SUPERSONIC PARACHUTE INFLATION RESEARCH EXPERIMENTS (ASPIRE) PROJECT. I. G. Clark 1and C. O’Farrell1 and M. Adler1 1Jet Propulsion Laboratory, California Institute of Technology, 4800 Oak Grove Drive, Mail Stop 321-220, Pasadena, CA 91109, [email protected]

Introduction: Since the landing of the twin Viking and reconstruction and investigation plans. The cam- spacecraft in 1976 [1], every United States mission to paign’s role in Mars 2020’s risk-reduction plans will be the Martian surface has used a variant of Viking’s Disk- discussed, along with the project’s broader science Gap-Band (DGB) parachute to decelerate from low su- goals and the benefits and limitations of the sounding personic speeds to the low subsonic speeds required for rocket architecture. terminal descent. The DGB parachute was designed, de- References: veloped, and tested during series of development cam- [1] Cooley, C .G. and Lewis, J. G. “Viking 75 Pro- paigns undertaken by NASA in the 1960’s and 1970’s, ject: Viking Lander System Primary Mission Perfor- which included wind tunnel testing, low altitude drop mance Report,” NASA Contractor Report 145148, testing, and high-altitude supersonic parachute test pro- April 1977. [2] Moog, R. D. and Michel, F. C., “Balloon grams [2]. Following Viking’s success, subsequent Launched Viking Decelerator Test Program Summary NASA missions to Mars successfully deployed a DGB Report,” NASA Contractor Report 112288, March during their descent phase [3]. While these missions 1973. [3] Cruz, J. R. and and Lingard, J. S., “Aerody- conducted subscale development tests and subsonic namic Decelerators for Planetary Exploration: Past, Pre- low-altitude qualification tests, none of these DGBs sent, and Future,” AIAA 2006-6792. [4] Gallon, J. C., were qualified supersonically before their successful Clark, I. G, and Witkowski, A. , “Parachute Decelerator use on Mars. Instead, previous missions have relied on System Performance During the Low Den- sity Super- heritage data from the Viking qualification. sonic Decelerator Program’s First Supersonic Flight Recently, NASA’s Low-Density Supersonic Decel- Dynamics Test,” AIAA 2015-2130. [5] O’Farrell, C., erators (LDSD) project has conducted supersonic, high- Brandeau, E. J., Tanner, C. L., Muppidi, S., and Clark, altitude tests of two large Ringsail parachutes for plan- I. G., “Reconstructed Parachute System Performance etary exploration. During supersonic testing at Mach During the Second LDSD Super- sonic Flight Dynamics numbers above 2.0, both parachutes experienced cata- Test,” AIAA 2016-3242. strophic failure at loads well below those the parachute had been successfully tested to in subsonic low-altitude testing [4,5]. The findings of the LDSD project have highlighted the need for supersonic, high-altitude testing of parachutes of the scale and materials necessary for future missions to Mars, as well as the importance of understanding the relationship between flight performance and behavior during subsonic testing. The Advanced Supersonic Parachute Inflation Re- search and Experiments (ASPIRE) project was established as a risk-reduction activity for the Mars 2020 project in 2016. The project will study the deployment, inflation, and performance of 21.5-m DGBs in supersonic, low-density conditions. The parachutes will be delivered to targeted deployment conditions representative of flight at Mars by sounding rockets launched out of NASA’s Wallops Flight Facility (WFF) starting in the in the summer of 2017. Two test articles will be deployed: a full-scale version of the DGB used by the Mars Science Laboratory, and a full- scale strengthened version of this parachute which has the same geometry but differs in materials and construction. This presentation will provide an overview of the ASPIRE project. It will discuss the sounding rocket test architecture, operations, test articles, instrumentation, De-orbit, Descent, and Landing on Europa: Key Challenges and an Architecture. D. M. Kipp,1 E.D. Skulsky1, A.M. San Martin1 1Jet Propulsion Laboratory, California Institue of Technology, 4800 Oak Grove Dr., Pasadena, CA 91109

Introduction: Jupiter’s moon Europa is of intense scientific interest because of the vast quantities of salty liquid water which likely lay beneath its thin icy crust and the tantalizing prospect of finding life elsewhere in the solar system. The planned Europa Mission, which would perform remote science through multiple flybys of Europa, is under development and promises to yield unprecedented insight into this intriguing body. How- ever, there remains a strong desire in the scientific community to perform in situ Europa science through a landed mission. Europa presents unique challenges to a landing mission because of its hostile radiation environment and the lack of information about its terrain. As a complement to the flyby mission, a bold concept to land on Europa and perform in situ science is being studied. Such a mission would require significant technology development to overcome the inherent landing challenges. This paper provides a brief overview of the Europa Lander mission concept and describes the significant challenges associated with landing on Europa, the technologies required to overcome those challenges, and a strategy for deorbit, descent, and landing. Recent Developments in Supersonic Retropropulsion for Mars Entry, Descent and Landing

Karl T. Edquist, Ashley M. Korzun, and Alicia M. Dwyer Cianciolo NASA Langley Research Center *[email protected] Manish Mehta NASA Marshall Space Flight Center

NASA continues to develop technologies for human exploration of Mars in the coming decades. In particular, new supersonic deceleration technologies are being developed to replace traditional parachutes, which do not scale well for human payloads (10s of metric tons). The leading descent technology for a wide range of payload sizes is supersonic retropropulsion (SRP), which involves using propulsive deceleration at supersonic Mach numbers, rather than just during the subsonic terminal descent phase. This paper will summarize the recent developments, and highlight the challenges, of advancing SRP’s maturity in preparation for implementation on a Mars mission in the 2020s. The Entry, Descent and Landing (EDL) project (2010 to 2012) is a recent example of NASA’s focus on advancing the technical maturity of SRP. The focus of that project was to execute two sub-scale supersonic wind tunnel tests using a generic model configuration and cold-gas air for the engine gas. A parallel task compared computational models against the wind tunnel data, with favorable results shown. Since 2012, NASA has continued to investigate SRP as the chosen descent technology for a wide range of Mars payload sizes. The most recent significant advancement of SRP has come from outside of NASA, with the use of SRP for reusing commercial launch vehicles. Inside NASA, recent studies have included SRP as the primary descent technology for a range of entry systems that use different hypersonic deceleration methods. Among the many challenges to advance SRP for Mars entry systems are ground testing and computationally modeling complex and unsteady SRP fluid dynamics. The flight mechanics behavior and thermal protection system requirements of an entry system will be influenced by the complex interactions between the SRP engine plumes, freestream, and vehicle. Short-duration, hot-fire engine test capabilities that recently were demonstrated for the NASA SLS program could be useful for testing SRP models with a realistic plume gas that more closely matches flight conditions. Such testing would provide a database against which high-fidelity fluid dynamics codes can be compared for relevant Mars geometries and conditions. Building confidence in these codes will help reduce the risk of predicting the effects of SRP fluid dynamics on the stability and aerothermal requirements of proposed entry systems. Recent examples of fluid dynamics predictions by NASA will be shown. There also are many design choices that result from using SRP on a Mars entry system. A few of those design parameters include engine location, whether or not to separate the hypersonic decelerator before SRP, and how to implement descent sensors in the presence of SRP engine plumes. Possible approaches to those challenges will be discussed for multiple hypersonic decelerator technologies. Finally, since the SRP engines will be used for terminal descent and landing, plume/surface interactions and vehicle landing stability are important considerations. The paper will discuss the status of this topic and approaches for modeling and testing the plume/surface interaction phenomena, as well as other recommended risk reduction activities prior to Mars missions using SRP.

MINIMUM FUEL THREE DIMENSIONAL POWERED DESCENT GUIDANCE FOR PLANETARY LANDING. Joel Benito, Jet Propulsion Laboratory, California Institute of Technology, 4800 Oak Grover Dr., M/S 198-326, Pasadena, CA 91109, USA. [email protected].

Introduction: The state of the art in powered de- ity. Also, during the parachute phase the vehicle is rap- scent guidance for planetary landing is represented by idly oriented vertically, and the vehicle is able to per- the Mars Science Laboratory [1], which used a polyno- form ground sensing. There is a limit, however, on how mial guidance approach [2]. Polynomial guidance was heavy a vehicle using parachutes can be, since at high originally formulated in the 60s and was used onboard ballistic coefficient the vehicle never reaches acceptable the Apollo lunar module for lunar landing [3]. The parachute deployment conditions. Vehicles using super- Apollo lunar guidance faced common problems for sonic retropropulsion (SRP) instead of a supersonic par- powered descent: a far downrange touchdown point, the achute enable larger masses but must accommodate a need to control the vehicle’s attitude to orient sensors larger range of initial conditions during the powered de- and allow a clear view of the ground to the pilot. The scent, using a limited amount of fuel. At SRP ignition Apollo lunar landing guidance planned the powered tra- the flight path is typically quite horizontal, which means jectory using a polynomial assumption for the accelera- that landing sensors may not be oriented to the ground tion. Polynomial guidance, although flexible and com- and that the touchdown point is a few kilometers down- putationally lightweight, cannot easily accommodate range while the altitude above ground may be less than trajectory constraints and disregards fuel optimality. a kilometer. Thus, in addition to the difficulty of ground Until now, Mars missions have used supersonic par- sensing at high speeds, there is the difficulty of estimat- achutes as the main means for terminal deceleration ing the position and velocity relative to the landing prior to powered descent. Parachutes are convenient be- point. In summary, SRP powered descent must accom- cause of their low mass and high deceleration power, modate a larger set of initial conditions and uncertain- which serves as a clean divider between entry and de- ties. Landing heavy payloads will require optimizing the scent: soon after parachute deployment the vehicle fuel usage, especially when pinpoint landing is required. reaches a velocity close to terminal velocity, which is a Optimal powered descent algorithms have been pre- function of vehicle weight, parachute drag area and at- sented, including a very promising numerical algorithm mospheric density, but independent of the initial veloc- based on convex optimization named Guidance for Fuel Optimal Large Diverts (G-FOLD) [4].

Figure 1: Sample powered descent trajectory for Mars landing. Note how the trajectory bends to the right looking to get close to the desired landing site. In order to cover the gap between the fully analytical phase, during which higher resolution images of the polynomial guidance approach and the fully numerical ground are obtained to pinpoint the final landing loca- approaches like GFOLD, a quasi-analytical approach tion. Then it may be possible to perform hazard detec- has been developed extending the work in [5]. The al- tion (rocks, slopes and potholes) followed by a hazard gorithm is based on a fixed attitude assumption during avoidance maneuver, leading the vehicle to a clean land- the descent, which is proven to be fuel optimal under the ing spot. Finally, the final touchdown maneuver takes right conditions. This characteristic enables an analyti- places, which enforces the touchdown vehicle limits. cal formulation, which requires some iterative compu- References: tations to obtain a solution. [1] Miguel San Martin, Gavin F. Mendeck, Paul B. Approach: The solution proposed, which is derived Brugarolas, et. al., “In-flight experience of the Mars Sci- from the inspection of the optimal trajectory resulting ence Laboratory Guidance, Navigation, and Control from focusing only on velocity, provides a minimum system for Entry, Descent, and Landing,” CEAS Space fuel solution. The algorithm provides the flexibility of Journal, June 2015, Volume 7, Issue 2, pp 119-142. targeting any combination of horizontal velocity, verti- [2] Singh, G., San Martin, A.M. and Wong, E.C., cal velocity and heading angle, unlike more rigid ap- ”Guidance and Control for Powered Descent and Land- proaches like gravity turn. Path constraints cannot be ac- ing on Mars,” IEEE paper 1548, Aerospace Conference, counted for, however, given the proposed formulation, 2007. the freedom in the resulting trajectory is small, and [3] A.R. Klumpp, “Apollo Lunar Descent Guid- therefore all the usual suspects in trajectory constraints ance,” Automatica, Vol 10, pp. 133-146, 1974. are de facto accounted for: attitude rates (none, because [4] Behcet Acikmese and Scott R. Ploen. "Convex the maneuver uses a fixed pitch), glideslope (deter- Programming Approach to Powered Descent Guidance mined by the boundary constraints, and in general well for Mars Landing", Journal of Guidance, Control, and behaved), guarantee (the solution is guaranteed to exist, Dynamics, Vol. 30, No. 5 (2007). provided the vehicle has enough fuel allocated). Given [5] Joel Benito, Erich Brandeau, Evgeniy Sklyan- these characteristics, it beats approaches like polyno- skiy and Steve Sell, ”Powered Descent Guidance Strat- mial guidance, which can provide badly behaved solu- egy and Algorithms for Mars Landing Using Supersonic tions, depending on the boundary conditions applied. A Retropropulsion,” IEEE Aerospace Conference, Big fixed attitude approach does not provide a solution for Sky, Montana, 2017. the end-to-end trajectory; additional guidance is necessary to orient the vehicle to the right attitude complete the landing. Some options for this end-to-end approach are outlined (see [5]). The algorithms are illustrated with trajectories corresponding to a high mass Mars landing; see. Fig. 1 for an example. This example shows the algorithm approach. Firstly, set of reachable landing sites using the proposed formulation is obtained (pink circle). Within this circle the algorithm searches for the location closest to the desired landing site. Also within this reachable set, the algorithm searches for the closest location within the reachable set that has an acceptable final horizontal velocity given some minimum and maximum bounds. The selected landing site is the one that satisfies the final velocity constraints and is the closest to the desired landing site. The final position error is minimized as a best effort; the focus of the algorithm is fuel usage minimization. The approach presented does not guide the vehicle to the ground, but to a point above it [5]. From that point on, at a relatively well known state and with a ground- relative navigation solution, the vehicle has options to reach the ground which are not outlined in this paper. Depending on the throttling capability of the vehicle, it may be possible to go to a constant descent velocity ! " # $!% & '() "#*# + , -# % ! " # - ! " #

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"# & 6 # . 03 # + .+=% . + #! + 03%. #+ 7 && ## # +# . ! ! " # /31 ! . &# ++ !& + ! ++#.!% "# +& #.# % + "# + . 03 +7&.% & 8-9-45 % + & # . % #% + &# # ## % # # ## &/**1! #.**&"% .7"# ** +# : 4 58 = 0/> ?1 # # +% + ! / 1+# +% ;@AB% #&. &# # + (C%? 450 < = + 2#** #"# + 2 . A - A ! % # # #+ ; ) 8" /> >1 + #.4 5 ! "#+ . # # #7% #& 4 5% + A" #+ % + ! + + + % &/"91 "# A Combined Algorithm for Final Descent Hazard-Relative Navigation and Hazard Detection. S. Woicke and E. Mooij2, 1 PhD candidate, Delft University of Technology, Kluyverweg 1, 2629HS Delft, [email protected], 2 Assistant Professor, Delft University of Technology, Kluyverweg 1, 2629HS Delft.

Introduction: Over the last decade many landings IMU-only lading. One should note that a couple of have been sucesfully performed on celestial bodies. outliers are present, which can be attributed to prob- However, with next-generation landers aiming for lems in the feature tracking and not the algorithm it- more complex landing scenarios, there is a great need self. Overall, in 97% of all cases the algorithm leads to to improve landing accuracy as well as to allow landers more accurate vehicle localisation opposed to an IMU- to sense and avoid landing hazards autonomously. only system. These abilities will enable landers to land in more complex terrain - or even unknown terrain - and will decrease the risk of a landing failure. To date, neither a full hazard-detection and avoidance (HDA) system, nor a terrain-relative navigation (TRN) system have been implemented and flown on any mission. Still, many developments were done in this field. Even though HDA produces outputs that might be used as input to the TRN system, combining or connecting both systems was not attempted yet. The current work tries to achieve this: using the output of HDA for TRN and using the TRN results for improved HDA. The resulting algorithm is capable of performing hazard relative navigation (HRN) of the landing vehicle. Algorithm: The algorithm focuses on lander local- Figure 1: TRN performance as compared to IMU- isation during the phase when hazard detection is pos- only localisation sible. This is at low atllitudes < 1km. At these altitudes A more detailed discussion of this algorithm no reference maps exist to perform inertial referencing can be found in [2]. of the HRN measurements. Therefore, the HRN algo- Hardware-in-the-loop testing: As software-in- rithm can only limit the relative errors. However, at the-loop testing was performed successfully, the next the moment hazards are identified, it is important to step is to perform testing of this algorithm in a more limit the accumulated error relative to the hazard posi- realistic hardware test. tions and the selected safe landing site, inertial locali- Conclusions: It was shown that it is feasible to sation is not necessary anymore. connect the HDA and TRN system to improve landing The developed algorithm uses a simmulataneous accuracy, and thus ensure safe landings. The algorithm localisatinon and mapping (SLAM)-like approach to was successfully tested using computer simulations. estimate the lander position using measurements ob- Hardware-in-the-loop testing is forseen for the near tained from stereo surface maps. These maps are used future. Moreover, this system provides the possibility for hazard detection and have thus a dual use; the ste- of combining multiple hazard maps into larger landing- reo algorithm is discussed in detail in [1]. Due to the region maps, which can serve as valuable input to sur- budgetary advantages of camera-based systems, cam- face mobility. eras were chosen as hazard-mapping sensor, however, References: different sensors could be employed too. [1] Woicke S., and Mooij, (2016) E. A stereo- One of the key features of the algorithm is that vision hazard-detection algorithm to increase plane- some surface points retrieved from the hazard maps are tary lander autonomy, Acta Astronautica 122 42-62. added to the state to improve the knowledge of the [2] Woicke S., and Mooij, (2017) E. Terrain Relative location of these points. As a final step these surface Navigation for Planetary Landing using Stereo Vision reference points can be used to assemble a large land- Measurements obtained from Hazard Mapping, Pro- ing-region map from all registered hazard maps. This ceedings of EURO GNC conference. map can, for example, serve as input to later rover operations. Figure 1 shows the results for a descent scenery, where the TRN localisation accuracy is compared to a

Mars 2020 Terrain Relation Navigation Performance During Landing. S. Mohan1, P. Brugarolas1, D. Way2, N. Trawny1, A. Stehura1, S. Dutta2, J. Montogmery1, A. Johnson1, S. Aaron1, A. Chen1 1NASA Jet Propulsion Labora- tory, Calilfornia Institute of Technology (4800 Oak Grove Drive, Pasadena CA 91109), 2NASA Langley Research Center (8 Lindbergh Way, Hampton, VA 23681).

Abstract: mate significantly exceeding required performance. The Terrain Relative Navigation (TRN) system is Current best estimates are on the order of 30m, com- an enabling Entry, Descent, and Landing pared to the 60m requirement. (EDL) technology baselined in the Mars 2020 mission [1]. TRN provides real-time, autonomous, terrain- relative position determination and generates a landing [1] Allen Chen et al. (2015) 2015 Update: Mars target based on a priori knowledge of hazards. TRN is 2020 Entry, Descent, and Landing System Overview, composed of the Lander Vision System (LVS) [2] and IPPW12 Presentation #2104. the Safe Target Selection (STS) algorithm [3]. The [2] Aaron Stehura et al. (2015) The Future of Land- LVS generates a map-relative localization solution ing: Terrain Relative Navigation From Prototype to by fusing measurements from a visible-wavelength Mars 2020, IPPW12 Presentation #3104. camera and an inertial measurement unit (IMU) using [3] Paul Brugarolas et al. (2015) On-Board Ter- the Map Relative Localization (MRL) algo- rain Relative Guidance-Target Selection for the Mars rithm operating on a high-performance compute ele- 2020 Mission, IPPW12 Presentation #3105. ment. Updated state knowledge is provided to the [4] Andrew Johnson et al. Design and Analysis of spacecraft navigation filter, which uses the Map Relative Localization for Access to Hazardous STS algorithm to direct a divert maneuver away Landing Sites on Mars, from known hazards within an onboard map. Mars 2020 has developed a high fidelity physics based simulation of LVS, which enables simulation over a wide variety of operational conditions. The LVS simulation (LVSS) is used to assess LVS performance over the entire operational envelope, which includes traditional descent conditions such as altitude, attitude, and velocity, but also includes aspects such as illumination, terrain relief, dust conditions, camera to IMU stability, and a range of camera errors. LVSS will be tuned based on field test results, and can be used to bound LVS performance. LVSS is used in conjunction with the high fidelity EDL simulation, Program to Optimize Simulated Trajecto- ries II (POST2), to determine the overall performance of TRN for Mars 2020. While LVSS provides LVS sensor level performance, POST2 provides a performance assessment for STS. The STS algorithm and corresponding map, which are flight products, have been integrated into POST2. The use of POST2 allows for assessment of how well STS performs, but also how the implementation of TRN affects the M2020 EDL performance with respect to the heritage MSL EDL. This paper presents a description of LVSS and POST2 and a description of the error sources modeled in each venue as it relates to TRN. The paper presents the overall TRN error budget structure and details which errors are derived from which simulation. Fi- nally, the paper presents design results from the TRN critical design review. Results show current best esti- THE MARS 2020 LANDER VISION SYSTEM: ARCHITECTURE AND I&T RESULTS. J. Montgomery,1 H. Ansari1, J. Chang1, Y. Cheng1, S. Schroeder1, N. Trawny1, J. Zheng1, A. Johnson1, 1Jet Propulsion Laboratory, California Institute of Technology, 4800 Oak Grove Drive, Pasadena, CA 91109 ([email protected]).

Human and robotic planetary lander missions re- Acknowledgments: This work was carried out at quire accurate surface relative position knowledge to the Jet Propulsion Laboratory, California Institute of land near science targets or next to pre-deployed as- Technology, under a contract with the National Aero- sets. These accurate position estimates can be obtained nautics and Space Administration. in real-time by matching sensor data collected during References: [1] Golombek M., et al. (2017) 3rd descent to an on-board map. The Lander Vision Sys- Landing Site Workshop for the 2020 Mars Rover mis- tem (LVS) being developed for the Mars 2020 mission sion, URL http://marsnext.jpl.nasa.gov/workshops generates landmark matches in descent imagery and /wkshp_2017_02.cfm. [2] Mustard, J.F., et al., (2014) combines these with inertial data to estimate vehicle Report of the Mars 2020 Science Definition Team position, velocity and attitude. The Mars 2020 EDL URL: http://mepag.jpl.nasa.gov/ system will use this position estimate to repurpose its reports/MEP/Mars_2020_SDT_Report_Final.pdf. [3] existing powered descent divert to avoid 100m scale Johnson, A. et al. (2015) “Real-Time Terrain Relative hazards identified in the landing ellipse map prior to Navigation Test Results from a Relevant Environment landing. This enables the selection of landing sites that for Mars Landing,” AIAA Guidance, Navigation, and have scientifically interesting terrain relief and were Control Conference. [4] Trawny N. et al., (2016) not selectable in the original implementation of MSL “Flight testing of terrain-relative navigation and large- EDL [1]. Mars scientists see great value in adding this divert guidance on a VTVL rocket,” AIAA Space Con- capability to the Mars 2020 lander mission [2]. ference. [5] Johnson A. (2016) “Design and Analysis LVS technology development at JPL has been on- of Map Relative Localization for Access to Hazardous going for more than a decade. A real-time flight-like Landing Sites on Mars,” AIAA Guidance, Navigation, prototype was flown in a helicopter field test conduct- and Control Conference. ed in February and March 2014 [3]. Position knowledge errors of less than 40m were demonstrated over a wide variety of terrains, illumination conditions, and attitude dynamics. The LVS prototype was then reengineered into a smaller self-contained package and integrated with a terrestrial vertical take off and landing rocket to demonstrate a closed-loop Mars EDL scenario [4]. Follow-on work focused on LVS algorithm improvements and an analysis, using MSL and other data sources, that showed the LVS could tolerate large changes in vertical motion [5]. The Mars 2020 flight design of the LVS occurred in parallel with this field testing and performance analysis and culminated in a preliminary design completed in November 2015. The LVS technology was baselined on the Mars 2020 mission in January 2016. Detailed hardware and software development has been on-going since this time. This presentation will describe the current LVS design architecture and initial integration and test results. It will touch on all aspects of the design including the Vision Compute Element (VCE) processor, the LVS Camera (LCAM), the Map Relative Localization algorithms and firmware, the map used for landmark matching, the flight software (VCEFSW) implementation, system interfaces (Figure 1, bottom) and LVS processing phases (Figure 1, top). Figure 1: The LVS system block diagram (bottom) and LVS processing phases (top). Mars 2020 Robust Rock Detection and Analysis Method. R. Otero, E. Almeida, B. Rothrock, M. Trautman, A. Huertas, M. Golombek, Jet Propulsion Laboratory, California Institute of Technology, 4800 Oak Grove Drive, Pas- adena, CA, 91109.

The hazard maps being generated for the M2020 mission are the most important maps to ever be created for a planetary mission. Not only are they being used to statistically select the proper placement of the ellipse, typical for most missions, these maps will be carried on board for landing-day guidance around identified hazards. This level of importance requires a new level of hazard and location accuracy. This work will detail a new method to more robustly and accurately determine rock fields than our current practice. The current state of the art in Mars landing site analysis characterizes rock-based hazards using cast shadows identified in HiRISE images. The shadow detection technique was detailed in earlier literature [1] and has been utilized for several flight missions to Mars.[2] The issue with the current technique is that the consistency of the rock hazard result is highly dependent on the ability of the operator to successfully find a set of parameters that balance false positive detections to false negative detections. The set of identified rocks is strongly dependent on the parameters selected. Figure Figure 1: Variability of Result to Parameters 1 demonstrates an example of variability for a rock analysis result using two different parameter sets. References: The new approach replaces the work of an opera- [1] A. Huertas, et al. (2006) IEEE Aerospace tor, to select a single set of parameters, with a sweep of [2] M. P. Golombek, et al. (2008) JGR, Vol 113, different parameters applied over the image. Each E00A09. parameter set is considered as a weak classifier (>50% accurate) for what is/isn’t a rock within the region of interest. A machine learning bootstrap aggregating technique is utilized to learn weights that should be applied to these weak classifiers to form an ensemble classifier with high rock detection accuracy. The results from this approach running on independent images over the same area are compared. Ad- ditionally, these results will be compared against the current best practice which involves parameter selection via a human operator. The usual practice of comparing statistical rock abundance results over the area will be performed as well as a more robust rock-to- rock comparison over the area of overlap. PILOT - PRECISE AND INTELLIGENT LANDING USING ON-BOARD TECHNOLOGIES T. Diedrich1, K. Hornbostel1, U. Soppa1, T. Jahnke1, J. Bolz1, B. Houdou2, D. de Rosa2 1Airbus Defence and Space (mailto: [email protected]), 2European Space Agency (mailto: beren- [email protected])

Introduction: PILOT (Precise and Intelligent Land- ing using On-board Technologies) is a suite of advanced landing technologies (hardware and software), enabling global access to the lunar surface, precisely and safely. It realizes the functions of visual navigation and hazard avoidance in an autonomous manner onboard a lander spacecraft. The PILOT subsystems consist of several sensors, i.e. landing cameras, a 3D-imaging LIDAR and a landing processing unit with high performance avionics.

Schedule: A first flight opportunity announced in 2021 (tbc), has been identified in the frame of the ESA- ROSCOSMOS cooperation on Lunar exploration. The PILOT subsystem will be integrated as part of the Guidance Navigation and Control of the Russian Luna- Resource-1 lander, enhancing its landing performance. It will allow reaching landing areas in the South Polar region of the Moon with an increased probability of success and to gain the experience and capability required to feed forward to future exploration missions. In addition, the possibility to embark a landing Camera Optical Unit as a “passenger” onboard the earlier Rus- sian-led Luna-Glob Lander mission has been identified in cooperation with ROSCOSMOS. The main objective of this precursor is to reduce the risk of precision landing for the Luna-Resource Lander by validating the landing camera and perform an imaging campaign in real external and platform environment.

Acknowledgemnt: The industrial team led by Airbus Defence and Space in Bremen, Germany integrates key contributions from partners all across Europe and Can- ada, in particular Omnisys (S), Airbus DS (F,G), OIP (B), NEPTEC (UK), Keopsys (F), GMV (RO,E), NGC (C) and TSD (I). The currently Phase B+ work has been established under ESA contract No. 4000116441/16/NL/SH (technical officer B. Houdou). DESIGN, SIMULATION ANALYSIS AND TESTS OF A RAM-AIR PARACHUTE RECOVERY SYSTEM. Zhuo Wu1, Zhen Yang1 and Qi Wang1, 1Beijing Institude of Space Mechanics and Electricity (P.O.Box 5142-269, Beijing, China,100094, [email protected]).

Abstract: Ram-air Parachute is a kind of parachute which can gild. With navigation and control equipment, Ram-air parachute can change the direction of flight to achieve autonomous and precise landing which is a great advantage over other kinds of parachutes in recovery missions. However, because of these special characteristics, it is difficult to develop a ram-air parachute recovery system. Simulation analysis based on computational fluid dynamics (CFD) can predict aerodynamic characteristics of ram-air parachute to identify design of ram-air parachute. Truck tow tests and drop tests lead to the design validation of ram-air parachute and successful performance results. This paper introduces a recovery system which makes a ram-air parachute as the main parachute to perform autonomous, soft and precise landing. Based on CFD simulation analysis, we got aerodynamic characteristics of the ram-air parachute, and then optimized the design. In truck tow tests and drop tests, the ram- air parachute was made to cooperate with other equipment, and the recovery system successfully completed the mission with high precision landing.

Structural and Configuration Design of a Novel Hypersonic Inflatable Aerodynamic Decelerator for Mars Entry, Descent, and Landing. L. Li1, N. Skolnik2, H. Kamezawa3, G. Rossman4, B. Sforzo5, and R. Braun6, [email protected], [email protected], [email protected], [email protected], [email protected], [email protected].

Introduction: Entry, descent, and landing (EDL) is especially challenging on Mars because the atmosphere is too thin to provide substantial deceleration, but thick enough to generate significant heating during the reentry phase. As a result, innovative ideas are required to enable future high-mass Mars landing missions. One such promising approach is to use an inflatable aerodynamic decelerator (IAD). Compared with traditional rigid aeroshells, IADs are made of lightweight, flexible materials that can be folded into a smaller volume in the rocket payload fairing and inflated prior to atmospheric entry. Such IADs are able to reduce the ballistic coefficient and peak heating, providing an opportunity to land at higher surface elevations on Mars. Currently, NASA Langley Research Center is investigating the development of Hypersonic Inflatable Aerodynamic Decelerators (HIADs) to enable future large robotic and human exploration missions. Much of the previous work performed on HIADs has focused on symmetric shapes that fly through the atmosphere with ballistic trajectories or trajectories with low lift- to-drag ratios accomplished via CG offset. The present investigation assesses the configuration design and structural feasibility of a novel asymmetric HIAD concept that can develop lift, is extensible to aeroshell diameters of 15 to 20 meters, and possesses an approximately smooth outer mold line to avoid localized heating. After determining the outer diameter and cone angle based on aerodynamic requirements, the number of tori and HIAD configuration were determined through a multiobjective minimization process which sought to minimize the mass of the decelerator and the complexity. Two different analyses were used to estimate the structural performed on the optimal configuration – high-fidelity finite element analysis in addition to an analytic approach used for validation of the results. Both analyses showed good agreement. Geometric modeling was performed on the final configuration to allow for visualization and multiple techniques were explored to allow for attachment and constraint of the tori. Overall, the analysis and design performed in this study showed the novel, asymmetric HIAD configuration to be feasible for future Mars missions.

PROTOTYPING NEW HIAD TECHNOLOGIES TO ENHANCE PERFORMANCE OF ENTRY, DESCENT, AND LANDING (EDL) SYSTEMS FOR HUMAN AND ROBOTIC MISSIONS TO MARS. B. Libben1, J. Williams2, S. Kosvick3, A. Scott4, 5A. Calomino, 1Masters Candidate, School of Aeronautics and Astro- nautics, Purdue University, 701 W. Stadium Ave., West Lafayette, IN, 47907-2045, email: [email protected], 2Masters Student, School of Aeronautics and Astronautics, Purdue University, 701 W. Stadium Ave., West Lafa- yette, IN, 47907-2045, email: [email protected] 3Undergraduate Student, School of Aeronautics and As- tronautics, University of Illinois, [email protected], 4Undergraduate Student, School of Aeronautics and Astro- nautics, University of Illinois, [email protected], 5NASA Langley Research Center, Hampton, VA, 23681, email: [email protected].

Introduction: The Entry, Descent, and Landing this end, a HIAD with dynamic morphing control is (EDL) phase of a Mars mission is one of the main chal- proposed. lenges facing a human-class mission to the “Red Plan- Study Goals: This study investigates concept gen- et.” With the density a fraction of Earth’s atmosphere, eration and the implementation of hardware to provide the Martian atmosphere is dense enough to create con- robust guidance of a Hypersonic Inflatable Aerody- cerns for heating, but does not provide sufficient drag namic Decelerator (HIAD) to enable precision landings to decelerate a high mass entry vehicle using traditional after atmospheric entry. It aims to propose a variety of aeroshells [1]. The recent 1-ton Curiosity mission ap- suggestions for control hardware, pursue a selected proaches the current size limit on rigid aeroshells that design to prove a simple, low budget, and robust solu- can be used on Mars with current launch vehicle vol- tion exists, and then report the lessons learned to aid ume constraints. To meet NASA's Human Journey to future investments in the field. Multiple concepts for a Mars mission, significant research must be invested controllable HIAD are generated to meet the require- into developing robust entry technology that will ena- ments necessary for precision landing based on trajec- ble manned missions to land safely near pre-deployed tory simulation, and then downselected to the final assets on the surface, consistent with currently pro- proposed concept based on a weighted decision proposed architectures. Compared to heritage rigid entry cess. A 1-meter diameter “demonstration of concept” systems, deployable decelerators allow for a larger prototype is then built using relevant materials to eval- drag area to slow down the lander, while possessing the uate the overall simplicity of the design from both a ability to fit into available launch vehicles when manufacturing and performance standpoint. A step-by- stowed. This new paradigm of entry vehicle design step production guide is generated, and suggestions and utilizing volume and mass efficient concepts is re- future considerations are discussed to further the de- quired to overcome volume constraints. In the EDL velopment of these systems. Systems Analysis (EDL-SA) study, flexible inflatable Design Process: To select a design that will meet decelerators were identified as an enabling technology the requirements of a low cost, robust, and easy-to- for landing large payloads on Mars [1]. manufacture system, an iterative process of brainstorm- Another challenge facing crewed Mars missions is ing sessions and prototype construction is implement- the requirement for pin-point landing. The atmospheric ed. Since the team had not worked on HIAD construc- variabilities during the hypersonic regime require any tion in the past, this was important to understand the trajectories to consider the landing error ellipse in- materials used in the system, and how feasible different flight. State-of-the-art technology involves using a rigid designs are from a manufacturing and implementation aeroshell with a center-of-gravity (CG) shift to gener- aspect. ate lift that provides control authority. This allows for Once the requirements are established, an initial guidance and control to be performed during the hyper- brainstorming session produces a wide range of possi- sonic phase, considerably decreasing the landing error ble concepts with unique features to allow for hybrid ellipse. solutions in the future. Construction of a single, circu- The recent Inflatable Reentry Vehicle Experiment lar toroid follows, which allows the team to better un- (IRVE) shows that a dynamic lifting Hypersonic Inflat- derstand the material behavior in deflated and inflated able Aerodynamic Decelerator (HIAD) is a feasible states, and to become familiar with the torus construc- design, and represents an upcoming game-changing tion process. This experience is used to determine, dur- technology [2]. Replacing the shifting CG lift control ing a second brainstorming session, which of the previ- IRVE III implemented with a shape-morphing aero- ously generated concepts are feasible to manufacture, shell would provide the modulating lift necessary to and then iterate on them to create both hybrid and decrease landing error with a faster reaction time. To unique concepts seen in Fig. 1.

Figure 2: Top and side views of the HEATS concept. Cables are used to actuate a toroid on top of the nominally symmetric stack to create an asymmetrical pressure distribution on the foreshell.

The T5’ is controlled using a configuration of four stepper motors positioned at every 90 degrees on the central payload. For the purposes of easier production and manipulating motor position during testing, these motors were mounted on an aluminum plate on top of Figure 1: Final generation of concepts, (a) Two unidi- the payload for the prototype. The four-stepper motor rectional elliptical toroids for isolated crossrange and assembly allows for individual control over each cardi- downrange control, (b) Cable system to morph outer nal direction, allowing T5’ to rotate over a full 360- toroid which is segmented into 4 bladders to allow for degree arc for precise control during flight. This design increased freedom, (c) Volume changing piston to was chosen mainly because it is scalable, lightweight, morph HIAD shape, and (d) Linear actuator system and effective in producing a significant shift in L/D. that push small inflatable tori with their own TPS into This design could be built and tested using a 1-to-20 the flow as control tabs. scale model of a full-sized 20-meter diameter HIAD article. The small timeframe for the design and produc- Multiple design reviews with a panel of experts in tion of the prototype was also a deciding factor. The the materials, construction, aerodynamic performance, simplicity of HEATS allowed the team to quickly de- and inflation system fields help drive the downselect sign and assemble a working prototype displayed in process, by inputting the major concerns each disci- Fig. 3. pline has in an entry system. After the second genera-

tion of concepts are finalized, the team designed and manufactured an asymmetric toroid, a concept with no previous construction research, to test the feasibility and schedule implications of including it in a lifting- body HIAD concept. Final Design: The selected prototype design is the Hypersonic Entry by Active Toroid Shifting (HEATS) concept seen in Fig. 2. The main concept behind HEATS is placing an identical torus to the symmetrical stack’s largest toroid, the T5 toroid in Fig. 2, on top of the main stack, denoted as the T5’ toroid. The T5’ is shifted along the horizontal plane on top of the stack which causes a protrusion into the flow. This protru- Figure 3: Final 1-m base diameter prototype of sion acts as a control surface, a common theme seen HEATS. Torii were inflated to 15 psi and actuated in among proposals for dynamic HIAD lift control. By all cardinal directions (90 deg separation) and one 45 moving this torus a distance of four inches into the deg pull to analyze system response. flow, HEATS can theoretically achieve a lift-to-drag ratio of 0.4, which was calculated through analysis based on Newtonian Flow Theory. References: [1] Cianciolo A. M. D. et al. (2010) NASA/TM-2010-2166720, “Entry, Descent and Land- ing Systems Analysis: Phase 1 Report.” [2] Smith B. P. et al. (2010) 2010 IEEE Aerospace Conference, 1-18, “A Historical Review of Inflatable Aerodynamic De- celerator Technology Development.” PLANNED ASSEMBLY, INTEGRATION, AND TESTING OF A 6M HIAD ORBITAL ENTRY VEHICLE. R. A. Dillman1, J. M. DiNonno1, S. J. Hughes1, R. J. Bodkin1, F. M. Cheatwood1, R. K. Johnson1, and K. M. Som- ervill1 1NASA Langley Research Center, Hampton, VA 23681, USA; [email protected], [email protected], [email protected], [email protected], [email protected], [email protected], and [email protected].

Introduction: The next planned flight test of a Hyper- as pneumatic load testing of the HIAD inflatable structure, sonic Inflatable Aerodynamic Decelerator (HIAD) is the and spin-balancing of the full assembly to minimize disturb- orbital velocity entry of a 6m diameter HIAD, intended to ance forces so the spin stabilized vehicle will maintain prop- verify the predicted aerodynamic and aero-thermodynamic er attitude from release through atmospheric entry. The environments as well as the fight performance of the HIAD increased size and mass of the planned HIAD, and the asso- aeroshell system, including the response of the HIAD flexi- ciated increased size and mass of the inflation system and ble thermal protection system and inflatable structure to the other components, will require larger handling and test fix- flight environments. Previous sounding rocket tests of 3m tures than previous flights and the use of powered equipment HIAD aeroshells have verified their ability to deploy in for integration of the major hardware sections, positioning in space from the packed configuration, maintain the inflated the necessary test facilities, transportation, and integration at shape through atmospheric entry and splashdown, and aero- the launch site. This paper will discuss the expected assem- dynamically maintain a stable orientation through all phases bly, integration, and test steps, and the support hardware of flight from hypersonic to subsonic. However, the sound- associated with those tasks. ing rocket tests exposed the HIADs to significantly lower entry heating than is expected in operational applications. Additional ground testing of 3-6m HIAD assemblies has demonstrated their structural and aerodynamic performance but could not fully duplicate the expected flight environments. This presentation will describe the proposed orbital entry test, and discuss in detail the assembly, integration, and ground testing planned to qualify the 6m HIAD for flight. Current plans call for launching a reentry vehicle incorporating a 6m HIAD as an Atlas V secondary payload. After release of the primary payload from the Centaur upper stage, and deployment of any other secondary payloads, the packed HIAD will inflate to its full 6m diameter while the reentry vehicle is still attached to the Centaur. At the appropriate point along its LEO trajectory, the Centaur will reorient, spin up, perform a deorbit maneuver, and release the HIAD flight experiment. Afterwards, the Centaur will perform a divert maneuver and break-up/burn-up in the atmosphere as usual. The HIAD entry trajectory is currently targeting a peak deceleration of 10G’s and should produce a peak entry heating slightly over 50W/cm2. These conditions are representative of HIAD conditions for direct entry at Mars as well as delivery of other payloads from LEO to Earth. The HIAD flight subsystems (inflation, data, power, etc) are intended to fit inside separate layers of the 1.3m diameter, 1.2m tall centerbody structure. Individual components will undergo the customary environmental qualification testing as required, with additional environmental and functional tests at the subsystem level after integration inside their centerbody layer, and further testing after integration of the full centerbody stack and the 6m HIAD. Planned tests include vibration to Atlas V launch levels and thermal exposure in vacuum chambers, as well as some less typical checks such FABRICATION OF THE HIAD LARGE-SCALE DEMONSTRATION ASSEMBLY AND UPCOMING MISSION APPLICATIONS. G. T. Swanson1, R. K. Johnson2, S. J. Hughes2, J. M. DiNonno2, F. M. Cheatwood2, 1AMA Incorporated, NASA Ames Research Center, Moffett Field, CA 94035 USA, [email protected], 2NASA Langley Research Center

Abstract: Over a decade of work has been con- near-term and future 10-15m HIAD applications will ducted in the development of NASA’s Hypersonic In- also be discussed. flatable Aerodynamic Decelerator (HIAD) technology. This effort has included multiple ground test campaigns and flight tests culminating in the HIAD project’s second generation (Gen-2) deployable aeroshell system and associated analytical tools. NASA’s HIAD project team has developed, fabricated, and tested inflatable structures (IS) integrated with flexible thermal protection system (F-TPS), ranging in diameters from 3-6m, with cone angles of 60 and 70 deg. In 2015, United Launch Alliance (ULA) announced that they will use a HIAD (10-12m) as part of their Sen- sible, Modular, Autonomous Return Technology (SMART) for their upcoming Vulcan rocket. ULA ex- pects SMART reusability, coupled with other advance- ments for Vulcan, will substantially reduce the cost of Figure 1. Large-Scale (top) and Subscale (bottom) access to space. The first booster engine recovery via HIAD Inflatable Structure Tori During Fabrication HIAD is scheduled for 2024. To meet this near-term need, as well as future NASA applications, the HIAD team is investigating taking the technology to the 10- 15m diameter scale. In the last year, many significant development and fabrication efforts have been accomplished, culminating in the construction of a large-scale inflatable structure demonstration assembly. This assembly incorporated the first three tori for a 12m Mars Human-Scale Path- finder HIAD conceptual design that was constructed with the current state of the art material set. Numerous design trades and torus fabrication demonstrations pre- ceded this effort. In 2016, three large-scale tori (0.61m cross-section) and six subscale tori (0.25m cross-section) were manufactured to demonstrate fabrication techniques using the newest candidate material sets. These tori were tested to evaluate durability and load capacity. This work led to the selection of the inflatable structure’s third generation (Gen-3) structural liner. In late 2016, the three tori required for the large-scale demonstration assembly were fabricated, and then integrated in early 2017. The design includes provisions to add the remaining four tori necessary to complete the assembly of the 12m Human-Scale Pathfinder HIAD in the event future project funding becomes available. This presentation will discuss the HIAD large-scale demonstration assembly design and fabrication performed in the last year including the precursor tori development and the partial-stack fabrication. Potential Gas Generators, the Challenges of Supporting Human-Scale HIADs (Hypersonic Inflatable Aerodynamic Decelerators) Richard J. Bodkin1, Walter E. Bruce2, F.M. Cheatwood3, Robert A. Dillman4 ,Johm M. DiNonno5 Philip M. Franklin6, Stephen J. Hughes7 Melvin H. Lucy8, Lisa K McCollum9 1NASA Langley Research Center, MS 432 Hampton, VA 23681, USA; Email: [email protected] 2NASA Langley Research Center, MS 431 Hampton, VA 23681, USA; Email: [email protected] 3NASA Langley Research Center, MS 489 Hampton, VA 23681, USA; Email: [email protected] 4 NASA Langley Research Center, MS 489 Hampton, VA 23681, USA; Email: [email protected] 5NASA Langley Research Center, MS 432 Hampton, VA 23681, USA; Email: [email protected] 6 NASA Marshall Space Flight Center, ER50 Huntsville, AL 35812, USA, USA; Email: [email protected] 7NASA Langley Research Center, MS 432 Hampton, VA 23681, USA; Email: [email protected] 8NASA Langley Research Center, MS 488 Hampton, VA 23681, USA; Email: [email protected] 9NASA Marshall Space Flight Center, ER52 Huntsville, AL 35812, USA; Email: [email protected]

Introduction: Hypersonic Inflatable Aerodynamic Decelerators (HIADs) are progressing from the small experimental scale to larger operational scale vehicles. One major technical challenge facing these larger vehicles is the need to upgrade the inflation system from the current blow-down approach, shown in Figure 1, to a more efficient system using gas generators, shown in Figure 2. Over the past year the HIAD-2 project has been researching the state-of-the-art for the gas generator market and beginning preliminary designs for gas generator inflation systems .

While no one technology has been selected several have been ruled out due to operational contraints and the limits of the inflation technology. Available technologies will be discussed in depth, as will possible methods for developing gas gererator systems that meet system requirement for total gas production required in allotted time, gas tempertatures within the limits of capability of HIAD material systems, gas compatibility Figure 1: IRVE-3 Inflation System Schematic with HIAD material systems, applicability to zero-g operation, tolerance to launch and cruise environments, reliability, operational safety and mass efficiancy.

Risk reduction and risk mitigation will also be discussed along with other HIAD requirments on the inflation system.

Figure 2: Gas Generator Inflation System Schematic HIAD AEROSHELL DEFORMATION TO GENERATE LIFT FOR USE IN DIRECT FORCE CONTROL STRATEGIES. S. J. Hughes 1, A. C. Slagle2, R. W. Powel3, A. M. Korzun4, F. M. Cheatwood4, 1NASA Langley Research Center, 1 North Dryden, M/S 432, Hampton, VA 23681 USA, [email protected], 2National Institute of Aerospace, 3Analytical Mechanics Associates Incorporated M/S 489, Hampton, VA 23681, 4NASA Langley Research Center, M/S 489, Hampton, VA 23681

Abstract: Hypersonic Inflatable Aerodynamic De- celerator (HIAD) technology has been identified as a potentially enabling for human-scale Mars entry systems. Further, entry simulations performed for NASA’s Entry, Descent, and Landing Architecture Study (EDLAS) indicate use of a direct force generation guid- Figure 2: Deformation of shoulder torus ance control strategy can save significant quantities of fuel in comparison to traditional bank angle control of a fixed lift direction. The flexible nature of the HIAD aeroshell provides additional opportunities for lift generation appart from the traditional aerodynamic control surfaces such as a tab or flap. Also, small deformations of the global shape, or pulling the perimeter of a portion of the aeroshell inward reducing drag area on one side of the aeroshell should produce more benign heating augmentation than employing a tab or a flap. Studies are being conducted for both global deformation of the entire HIAD aeroshell assembly and local deformation of single aeroshell elements to evaluate effectiveness of Figure 3: Deformation of HIAD test assembly both approaches for lift generation. (See figures 1 and 2.) In addition, testing is being performed utilizing existing HIAD project inflatable structure test assemblies to determine the forces required to generate the desired shape change and to correlate the deflected shape with the analytical model predictions. (See figure 3.) With the knowledge of the force required to generate the shape change, an estimate of additional system mass of the actuation system, power required to drive the deformation, and feasible actuation rates can be made. The lift force and actuation rate information generated can then be fed back into the entry simulation to predict the flight system performance of a deformed HIAD utilizing direct force control.

Figure 1: Generic deformed HIAD geometry AERODYNAMIC CAPABILITY OF FLAPS FOR PLANETARY ENTRY VEHICLES. J. Sepulveda1 and Z. R. Putnam1, 1University of Illinois at Urbana-Champaign, 104 S. Wright St., Urbana, IL 61801.

Independently articulated aerodynamic flaps may provide benefits during entry, descent, and landing by modifying the vehicle’s angle of attack and sideslip angle, therefore also modifying the lift-to-drag ratio and lift direction. The ability of flaps to directly control the vehicle’s lift vector eliminates the need for a center of gravity offset and may provide a relatively constant vehicle attitude. A near-constant attitude may enable the use of relative navigation sensors and regional- scale science investigations during the hypersonic portion of entry. The Configuration Based Aerodynamics tool was used to predict the trim angle of attack, trim lift-to-drag ratio, lift coefficient, and drag coefficient for variations in the number of flaps, individual flap configurations, and deployment angles. Results for one, two, and four flaps are presented for the hypersonic and supersonic regimes with the angle of attack ranging from -4 degrees to 20 degrees. These results demonstrate the effects of each flap, or set of flaps, on the aerodynamic performance of an entry vehicle and will inform the development of entry vehicle configurations and guidance, navigation, and control systems. NONLINEAR ADAPTIVE CONTROL LAW DESIGN WITH NONPARAMETRIC BAYESIAN MODELS FOR ADAPTING TO ATMOSPHERE AND WIND UNCERTAINTIES IN MARS ENTRY. N. Kemal Ure and Gokhan Inalhan, Aerospace Research Center, Istanbul Technical University, Maslak, Istanbul, 34469, Istanbul, Turkey ([email protected], [email protected])

An integral part of future Mars exploration mis- In this talk, we demonstrate that an NPBM based sions is achieving precise landing at prespecified loca- nonlinear control law can be used to design an entry tions [1]. Landing accuracy depends on several factors, control system, where the NPBM part plays a key role such as aerodynamic characteristics of the landing ve- in adapting for complex and stochastic atmosphere and hicle, atmospheric perturbations and the set of actua- wind dynamics. We will present simulation results that tors/sensors on the vehicle. Among these different fac- compare the tracking accuracy with various existing tors, control laws implemented on the vehicle and their approaches and show the improvements offered by the ability to operate under highly uncertain atmospheric NPBM based design. and wind conditions play a key role in determination of References: the landing precision [2]. In this talk, inspired by the [1] Braun, Robert D., and Robert M. Manning. "Mars recent success of nonparametric Bayesian models, we exploration entry, descent, and landing challenges." Journal apply a new family of nonlinear adaptive control laws of spacecraft and rockets 44.2 (2007): 310-323. to the design entry control laws, which demonstrates a [2] Steinfeldt, Bradley A., et al. "Guidance, navigation, significant improvement against existing control de- and control system performance trades for Mars pinpoint sign approaches, in terms of tolerating atmospheric landing." Journal of Spacecraft and Rockets 47.1 (2010): uncertainties and minimizing landing dispersion. 188-198. The problem of designing control laws that can [3] Li, Shuang, and Xiuqiang Jiang. "Review and pro- compensate for uncertain atmospheric and wind condi- spect of guidance and control for Mars atmospheric en- tions have been extensively studied in the last years try." Progress in Aerospace Sciences 69 (2014): 40-57. [3]. A variety of different approaches were proposed, [4] Marwaha, Monika, and John Valasek. ranging from adaptive control [4] to sliding mode con- "Fault tolerant control allocation for Mars entry vehicle trol [5]. Although these methods demonstrated suc- using adaptive control." International Journal of Adaptive cessful entry and landing performance in simulations, Control and Signal Processing 25.2 (2011): 95-113. in most of these works atmosphere and wind uncertain- [5] Li, S., and Y. M. Peng. "Neural network-based slid- ty models are limited to simple parametric uncertain- ing mode variable structure control for Mars en- ties and/or additive Gaussian noise assumptions, which try." Proceedings of the Institution of Mechanical Engineers, can fail to capture highly nonlinear and stochastic dy- Part G: Journal of Aerospace Engineering 226.11 (2012): namics of the atmosphere and wind uncertainties. Fail- 1373-1386. ure to capture these complex dynamics might result in [6] Quadrelli, Marco B., et al. "Guidance, navigation, accumulating tracking error over the time, which de- and control technology assessment for future planetary sci- grades the entry and landing performance. It is recog- ence missions." Journal of Guidance, Control, and Dynam- nized that more complex models of these external dis- ics 38.7 (2015): 1165-1186. turbances are needed for improving the control per- [7] Fox, Emily B. Bayesian nonparametric learning of formance [6]. complex dynamical phenomena. Diss. Massachusetts Insti- Modeling complex stochastic systems is a major tute of Technology, 2009. objective for machine learning/artificial intelligence. [8] Campbell, Trevor, et al. "Planning under uncertainty Recently, nonparametric Bayesian models (NPBMs) using nonparametric bayesian models." AIAA Guidance, showed a great promise in modeling complex dynam- Navigation, and Control Conference (GNC). 2012. ical phenomena [7]. Unlike probability distributions [9] Chowdhary, Girish, et al. "Bayesian nonparametric with fixed parameters, NPBMs offer a higher level of adaptive control using gaussian processes." IEEE transac- flexibility by allowing learning algorithms to infer tions on neural networks and learning systems 26.3 (2015): such parameters from the data. The use of NPBMs have been explored succesfully in various aerospace guidance and control problems [8]. In particular, Chowdhary et al. demonstrated that Gaussian Process- es (a particular form of nonparametric Bayesian model) can be utilized to develop nonlinear adaptive control laws that enable precise control of stochastic nonlinear systems [8]. EXPERIMENTAL DEMONSTRATION OF MAGNETOHYDRODYNAMIC ENERGY GENERATION IN CONDITIONS AND CONFIGURATIONS RELEVANT TO PLANETARY ENTRY. H. K. Ali1*, J.E. Polk2, and R. D. Braun3, 1Graduate Research Assistant, [email protected], 2Principal Engineer, NASA Jet Propulsion Laboratory, California Institute of Technology, [email protected], 3Dean, College of Engineering and Ap- plied Sciences, University of Colorado Boulder, [email protected], *Space Systems Design Lab, Daniel Guggenheim School of Aerospace Engineering, Georgia Institute of Technology, 270 Ferst Drive, Atlanta, GA 30331.

Introduction: Many concepts for future missions The number of experimental investigations of MHD to Mars, such as potential Mars Sample Return and interaction for conditions relevant to planetary entry is eventual human exploration, could require much higher limited. There are a few experimental investigations da- masses than have ever been landed on Mars. Previous ting back to the late 1950s and 1960s, mainly concerned Mars missions have relied primarily on Viking era tech- with the effects MHD interaction has on shock standoff nology for entry descent and landing [1].The limit of ditance.[5][6][7][8] More recent studies describe vari- this technology is being reached, with the Mars Science ous experimental campaigns aimed at investigating drag Laboratory (MSL) landing system in 2012 illustrating enhancement, heat-flux mitigation, artificial ionization, the difficulty in high mass Martian landings. and energy generation through MHD interaction.[9]

[10][11][12]From all of these experiments, it is apparent To achieve humanity’s goals for Mars exploration, that creating and testing MHD interaction in conditions significant technology development is required. Mass relevant to planetary entry is a challenging task requir- reducing technologies are particularly critical in this ef- ing specialized equipment. fort. Not only does a larger mass require more fuel to launch, but it also carries significantly more kinetic en- Results: The principal contribution made in this ergy that must be reduced to near zero if the vehicle is area is the design and execution of an experimental cam- to land safely. For all previous Mars missions, this ki- paign to demonstrate MHD energy generation for a non- netic energy has been primarily dissipated through aer- channel type MHD energy generation on a simulated odynamic drag during the hypersonic phase, which is blunt-body reentry vehicle. A gas source is accelerated limited by the size of the entry vehicle. In addition, Pre- to supersonic speed, artificially ionized to create a su- vious Mars missions have shown that the majority of the personic plasma, and allowed to pass over a perma- vehicle’s kinetic energy is dissipated during the hyper- nently magnetized ceramic model with electrodes for sonic entry phase, about 92% in the case of Mars Path- current collection. This experimental design has been finder [2]. demonstrated with a sample configuration, and appears

to show a positive effect for current through the MHD During this hypersonic entry phase, there exists a generator model when the supersonic plasma discharge highly heated, ionized flow around the entry vehicle. is present. This entry plasma is inherently conductive, and may thus be influenced by electromagnetic fields, facilitating magnetohydrodynamic flow interaction. This flow interaction can be used to provide a Lorentz force in addition to aerodynamic drag, thereby enhancing deceleration capability without necessarily increasing the size and mass of the entry vehicle forebody.

Magnetohydrodyanmic flow interaction for high speed aerospace applications has been studied since the Figure 1. Experimental demonstration in operation. dawn of the space race, with early theoretical studies dating back to the late fifties and early sixties [3][4]. At that time, practical implementation was limited due to the difficulties in generating high strength magnetic Acknowledgments: This work is funded through a fields and electrical energy storage. Since that time, NASA Space Technology Research Fellowship, grant however, dramatic advances in superconductivity and number NNX13AL82H. In addition, the authors would electrical energy storage have been made, warranting like to thank Dr. Robert Moses of NASA Langley Re- additional study. search center for his assistance in the completion of this work. References: “System Development for Mars Entry in Situ [1] R. D. Braun and R. M. Manning, “Mars Resource Utilization,” in Proceedings of the Exploration Entry , Descent and Landing 8th International Planetary Probe Workshop, Challenges,” in IEEE Aerospace Conference, 2011, no. 1, pp. 1–10. 2006. [2] D. A. Spencer, R. C. Blanchard, R. D. Braun, P. H. Kallemeyn, and S. W. Thurman, “Mars Pathfinder Entry, Descent, and Landing Reconstruction,” J. Spacecr. Rockets, vol. 36, no. 3, pp. 357–366, 1999. [3] W. B. Bush, “A Note on Magnetohydrodynamic-Hypersonic Flow Past a Blunt Body,” Journal of the Aerospace Sciences, vol. 26, no. 8. pp. 536–537, 1959. [4] P. O. Jarvinen, “On The Use of Magnetohydrodynamics During High Speed Re-Entry,” Everett, Massachusetts, 1964. [5] R. W. Ziemer and W. B. Bush, “Magnetic Field Effects on Bow Shock Stand-Off Distance,” Phys. Rev. Lett., vol. 1, no. 2, pp. 58–59, 1958. [6] E. Locke, H. E. Petschek, and P. H. Rose, “Experiments with Magnetohydrodynamically Supported Shock Layers,” Everett Massachusetts, 1964. [7] G. R. Seemann and A. B. Cambel, “Observations Concerning Magnetoaerodynamic Drag and Shock Standoff Distance,” Proc. Natl. Acad. Sci. U. S. A., vol. 55, no. 3, pp. 457–465, 1966. [8] S. Kranc, M. C. Yuen, and A. B. Cambel, “Experimental Investigation of Magnetoaerodynamic Flow Around Blunt Bodies,” Evanston, IL, 1969. [9] M. Kawamura, A. Matsuda, H. Katsurayama, H. Otsu, D. Konigorski, S. Sato, and T. Abe, “Experiment on Drag Enhancement for a Blunt Body with Electrodynamic Heat Shield,” J. Spacecr. Rockets, vol. 46, no. 6, pp. 1171– 1177, 2009. [10] A. Gülhan, B. Esser, U. Koch, F. Siebe, J. Riehmer, D. Giordano, and D. Konigorski, “Experimental Verification of Heat-Flux Mitigation by Electromagnetic Fields in Partially-Ionized-Argon Flows,” J. Spacecr. Rockets, vol. 46, no. 2, pp. 274–283, 2009. [11] D. J. Drake, S. Popović, L. Vušković, D. J. Drake, S. Popovi, and L. Vuškovi, “Characterization of a supersonic microwave discharge in Ar / H 2 / Air mixtures Characterization of a supersonic microwave discharge in Ar / H 2 / Air,” J. Appl. Phys., vol. 63305, no. 2008, 2008. [12] S. Popović, R. W. Moses, and L. Vuskovic, ADEPT SOUNDING ROCKET ONE (SR-1) FLIGHT EXPERIMENT DESIGN SUMMARY P. F. Wercinski1 1NASA Ames Research Center, M/S 229-3 Moffett Field, CA [email protected]

Introduction: The Adaptable, Deployable Entry and backshell) entry vehicle configurtion. The flight ex- Placement Technology (ADEPT) architecture repre- periment will utilize many features intended for one sents a completely new approach for entry vehicle meter scale space flight missions such as the carbon- design utilizing a high performance carbon-fabric to fabric decelerator, two-stage spring system for de- serve as the primary drag surface of the mechanically ployment, and a payload volume capable of accom- deployed decelerator. The ADEPT project team is modating a 3U CubeSat. This paper will describe the advancing this decelerator technology via systems- overall concepts of operations (Figure 1) of the SR-1 level testing at the one meter diameter (nano-ADEPT) flight experiment from launch to recovery and a sum- scale. The ADEPT SR-1 flight experiment, a 0.7m mary of key design features and associated driving diameter deployed mechanical decelerator, will utilize requirements. the NASA Flight Opportunities Program sounding rocket platform provided by UP Aerospace to launch SR-1 to an apogee over 100 km and achieve re-entry conditions with a peak velocity of over Mach 3. Launch is scheduled for late 2017. The SR-1 flight experiment will demonstrate most of the primary end- to-end mission stages including: launch in a stowed configuration, separation and deployment in zero-g, exo-atmospheric conditions, and passive ballistic re- entry of a 70 degree half-angle cone geometry. The ADEPT SR-1 will demonstrate the supersonic through transonic aerodynamic stability of the unique ADEPT axisymmetric, blunt body shape with an open-back (no

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'"#""(&! &#((#"-'(!&!"'#"#(!#'( ""(' '#$ "(&-,$ #&(#"!''#"' )'#(&'"''#& "(&-"*&#"!"('&#)"('('""#(&$ (('#"(#"' #"'%)"( - ""&'!)'(& -#"")!& !# '((#(" * (#"(#$&#* $((#+&",$"'** (#"("() -"(&-"*&' - #-'(! '$&('"* #$(("*&'(-#"() -

'("# #-!()&(#"'($(#+&(*"() # '$&(+ & '&#!(+#'$&('#)""&# ('( (()'(##("(#&( ' )"'&"!$#&("('($(#+&(",($'& '" ('(&( &#!("(&"(#" $((#"-&,$((#&'( "# #-"''* (#-!#"'(&("(%)'(#" #!!)"(#"'"'""'"'!#"'(&("&"''#&& '&#! ( Free Flight Ground Testing of ADEPT in Advance of the Sounding Rocket One Flight Experiment. B. P. Smith1 and S. Dutta2, 1NASA Ames Research Center (M/S 229-1, Moffett Field, CA, [email protected]), 2NASA Langley Resarch Center ([email protected]).

Abstract: The Adaptable Deployable Entry and Placement Technology (ADEPT) project will be conducting the first flight test of ADEPT, titled Sounding Rocket One (SR-1), in just two months [1]. The need for this flight test stems from the fact that ADEPT’s supersonic dynamic stability has not yet been characterized. The SR-1 flight test will provide critical data describing the flight mechanics of ADEPT in ballistic supersonic flight. These data will feed decision making on future ADEPT mission designs. This presentation will describe the SR-1 scientific data products, possible flight test outcomes, and the implications of those outcomes on future ADEPT development. In addition, this presentation will describe free-flight ground testing performed in advance of the flight test. A subsonic flight dynamics test conducted at the Vertical Spin Figure 2. ADEPT ballistic range test articles. Tunnel (VST) located at NASA Langley Research Center provided subsonic flight dynamics data at high References: and low altitudes for multiple center of mass (CoM) [1] Wercinski et al., “ADEPT Sounding Rocket locations. Figure 1 shows a photograph of one of the One (SR-1) Flight Experiment Design Summary,” 50% scale test article flow in the VST. A ballistic IPPW14 (2017). range test at the Hypervelocity Free Flight Aerody- namics Facility (HFFAF) located at NASA Ames Re- search Center provided supersonic flight dynamics data at low supersonic Mach numbers. Figure 2 dis- plays the five test articles fired in the HFFAF. Execu- tion and outcomes of these tests will be discussed. Fi- nally, a hypothesized trajectory estimate for the SR-1 flight will be presented.

Figure 1. ADEPT Vertical Spin Tunnel (VST) test article (50% flight scale = 0.35 m diameter). The test article is 3D printed plastic with tungsten weights and CoM adjustment. ASSESSING NONEQUILIBRIUM FLOW CHEMISTRY WITH A CUBESAT-CLASS MISSION J. W. Wil- liams1 and Z. R. Putnam2, 1Graduate Student, University of Illinois at Urbana-Champaign ([email protected]), 2Assistant Professor, University of Illinois at Urbana-Champaign ([email protected]).

Planetary probes are sensitive to the high-enthalpy, nonequilibrium, chemically-reacting, rareified flow in the upper atmosphere. Limited flight data from this regime requires large design margins for thermal protection systems, which increase the mass and cost of missions. Characterization of nonequilibrium flow on the ground is difficult with current technology and previous missions have provided incomplete or flawed data. The Student Aerothermal Spectrometer Satellite of Illinois and Indiana (SASSI2) will 1) to better identify the aerothermal and thermochemical effects of high-enthalpy, nonequilibrium flow and 2) verify and validate the effectiveness of commercial-of-the-shelf science-grade instruments to reduce risk for more advanced follow-on missions. The SASSI2 payload consists of multiple spectrometers, pressure sensors, and temperature sensors to capture data on the number and composition of molecules behind a diffuse bow shock as the vehicle reenters the upper atmosphere. These data will be used to validate and improve current computational fluid dy- namicsmodels used in the continuum and transi- tional/free molecular regimes. Models representing current best estimates of the expected aerothermal environment will be used to calibrate the spectrometers. The mission profile and preliminary model development are discussed. ANALYTICAL MODELLING OF INTERNAL THERMODYNAMICS OF ABLATIVE THERMAL PROTECTION SYSTEM MATERIALS. S. Pavesi1, A.S. Pagan1 and G. Herdrich1, 1Institute of Space Systems, University of Stuttgart, Pfaffenwaldring 29, 70569 Stuttgart, Germany, Email: [email protected].

Abstract: During planetary re-entry a vehicle is terms, respectively. The conductive term is evaluated as subjected to severe heating conditions, especially in function of other terms, thus if the latter increase it case of planets with highly dense atmosphere. Ablative would change accordingly. Once the in-depth energy Thermal Protection Systems (TPS) represent the only terms have been estimated, boundary conditions are in- valiable option for aforementioned applications. There- vestigated. A link between the in-depth response and the fore, continuous material investigation and characteri- simplified surface energy solution is established zation are needed to improve overall TPS performance. through the conductive heat flux, evaluated as close as In this direction, predicting temperature and density his- possible to the receding surface. Consequently, the heat tories of an ablative material is of utmost importance. energy transported by heat conduction or by diffusion Evaluating the transient thermal response of an ablative currents carrying chemical enthalpy due to different material can be complex and mainly two approaches are chemical species gases, ݍሶ௪, is evaluated as function of available: conducting experimental campaigns on a spe- the remaing surface energy terms. The most relevant cific materials or developing numerical models to solve outgoing heat flux is the radiation away from the sur- the equations governing the main processes within the face, while the pyrolysis gas mass flow rate and the con- material. The first approach has the advantage of relying vective heat flux from the boundary layer gases are the on direct measurements of material proprieties of inter- higher incoming contributes. The evaluated diffusive est. Nevertheless, entry conditions' simulation is prone energy term is in accordance with the conventional lim- to a certain margin of uncertainties, it might be time its of catalysis effect on the reference heat flux produced consuming and expensive. Moreover, it is not able to by the generator in Plasma Wind Tunnel 1 (PWK1) and discern between all contributions of each process occur- it accounts for all the thermo-chemical processes due to ring within material and boundary layer, which are di- the interaction between the boundary layer and the ab- rectly connected to the final performance of the ablator. lating surface [1]. The latter approach allows for great savings in time, but it relies on statistical tools or finite difference methods, References: which are by definition approximations of the analysed [1] S. Pavesi (2017), Characterisation of Ablative phenomena. Therefore, a combination of both these ap- Thermal Protection System Materials through Analysis proaches has been adopted to better understand the ab- and Inverse Methods, Master Thesis. [2] J. Rieser lation process and the corresponding performance of the (2016), Experimentelle und Numerische Charakterisie- ablative material [1]. More specifically, data from tests rung des Ablativen Hitzeschutzmaterials ZURAM®, Ba- on ZURAM®, a lightweight ablative material developed chelor Thesis. [3] D. Bianchi (2008), Modeling of Abla- by the German Aerospace Centre in Stuttgart and char- tion Phenomena in Space Applications, Ph.D. Thesis. acterised under relevant re-entry conditions at the Insti- [4] W. Li, H. Huang, Y. Tian, and Z. Zhao (2015), Non- tute of Space Systems (IRS), have been processed and linear analysis o n thermal behavior of charring mate- used as input for solving the energy equations which rials with surface ablation, In: International Journal of characterise the ablation process [2]. Heat and Mass Transfer 84, pp. 245-252. To analyse the different processes in the boundary layer and within the ablative material, more specifically a charred ablator, a preliminary division of the ablator in different zones is performed based on the main physical-chemical phenomena occurring within each specific layer. Main assumptions include: one-dimensional thermal conduction, thermal equilibrium between the pyrolysis gases formed and the char layer, absence of secondary cracking of pyrolysis gases and constant cross-section area. Moreover, the char is assumed to be composed of carbon only and failure modes are not considered [3][4]. After solving the in-depth energy equation, the conductive energy and convective energy from pyrolysis gases result to be the higher incoming and dissipative energy DEFINING THE CRITICAL DEPTH OF IMPACT DAMAGE FOR THERMAL PROTECTION SYSTEMS. N. L. Skolnik1 and Z. R. Putnam1, 1University of Illinois at Urbana-Champaign.

The future of spaceflight includes both human and robotic missions. In both cases, the reliability of the thermal protection system (TPS) is imperative for the safety of the astronauts and the mission as a whole. The purpose of TPS is to prevent the plasmas, and heat from the plasma, created during entry from reaching the interior of the spacecraft. It is important to both know the amount of damage the TPS can withstand from micro meteoroid/orbital debris (MMOD) impacts as well as how much repair is necessary to ensure that the spacecraft can enter the atmosphere and land at its destination. Research shows that impact holes in TPS undergo widening when exposed to entry conditions. One option for improving the survivability of TPS is to integrate a self-healing mechanism into the material so that the widening effect seen in the arc jet experiments is reduced or delayed. The key to making a self-healing TPS efficient in both complexity and mass is to find the critical depth of MMOD penetration which causes the TPS to fail. This will occur when the heat from reentry reaches a critical level at the TPS bondline. The critical depth is a function of the TPS material, location on the heat shield, entry environment, and impact energy. Once the critical depth is found, a self-healing system could be integrated at that depth, so that only damage that would cause mission failure activates the system. This analysis defines the critical depth through use of experimental results and computational tools.

LOCAL TERRAFORMING OF ICY BODIES AS A MEDIUM FOR AERO-BRAKING SYSTEMS. J. V. Koch1 and R. M. Winglee2, 1Graduate Student, Aeronautics and Astronautics, University of Washington, 2Professor, Earth and Space Sciences, University of Washington.

Introduction: Entry, descent, and landing technologies for the class of space exploration missions to bodies with atmospheres traditionally exploit the ambient atmosphere of the body as a medium through which the spacecraft or probe can interact to transfer momentum and energy for a ‘soft’ landing. For bodies with no appreciable atmosphere, a significant engineering challenge exists to overcome the lack of passive methods to decelerate a spacecraft or probe. Proposed is a novel means for the creation of a local atmosphere for airless icy bodies through the use of a two stage payload- penetrator probe. The first stage is a hypervelocity penetrator that impacts the icy body. The second stage is an aero-braking-capable probe directed to pass through the ejecta plume from the penetrator. Both experimental and computational studies show that a controlled high-energy impact can direct and transfer energy and momentum to a probe via a collimated ejecta plume. In an effort to provide clarity to this unexplored class of missions, a modeling-based engineering approach is taken to provide a first-order estimation of some of the involved physical phenomena. Three sub-studies are presented: an examination and characterization of ice plumes, modeling plume- probe interaction, and the extension of plume modeling as the basis for conceptual mission design. The modeling efforts are centered about the arbitrary Largrangi- an-Eulerian (ALE) set of techniques for material ad- vection and fluid-structure interaction through extensive use of the hydrocode LS-DYNA. A database of fully-developed hypervelocity impacts and their associated plumes is created and used as inputs to a 1-D mathematical model for the interaction of a continuum- based plume and probe. A simplified parametric study based on the hypervelocity impact and staging of the probe-penetrator system is presented and discussed. Shown is that a tuned penetrator-probe craft has the potential to increase spacecraft payload mass fraction over conventional soft landing schemes. References: [1] Winglee, R. M., C. Truitt, and R. Shibata, Sam- ple Return from Airless Solar System Objects, Acta Astronautica, submitted 2017. [2] Gowen, R. A., et al., Penetrators for in situ subsurface investigations of Europa, Adv. Space Res., 48, 725, 2011. EHD and MHD actuators: experimental works to enhance spacecraft control and deceleration during atmospheric entries. S. Coumar1 and V. Lago1. 1Laboratoire ICARE, CNRS, UPR 3021, 1C Avenue de la Recherche Scienti_que, 45071, Orléans Cedex 2, France. Mail : [email protected]

Study context: During the reentry phase into a dense gas atmosphere, spacecraft flying at hypersonic velocities are exposed to severe conditions and the mission planning must delicately balance three requirements: deceleration, heating management and accuracy of the localization and velocity when landing. In addition, the phenomenon of "Blackout" appears, preventing all radio transmissions. To this regard, one of the promising technologies consists in using flow control methods with plasma (EHD) and MHD based devices. Work introduction: Experimental investigations, carried out at the Icare Laboratory by the FAST team, focuse on plasma and MHD flow control in supersonic and rarefied regime[1]. This study aims at analyzing Fig.1: Comparison of the shock angle increase rate how the Mach number as well as the ambient pressure for the three nozzles modify the repercussions of the plasma actuator (EHD) MHD actuating: This experimental work is the on the shock wave. It follows previous experiments subject of preliminary studies on the interaction of a performed in the wind tunnel Marhy (ex–SR3) with a magnetic field with supersonic plasma developed Mach 2 flow interacting with a sharp flat plate where around an obstacle [3]. Previous works showed the modifications induced by a plasma actuator have been effects due to the interaction of the magnetic field with observed. Experimental measurements showed that the the plasma. This study was carried out with air plasmas boundary layer thickness and the shock wave angle and argon plasmas and two geometries have been stud- increase when the discharge is ignited and the results ied: a blunt body and a flat plate. Results showed that were presented last year. the plasma flow field is modified in the presence of EHD actuating: The current work is realized with magnetic field; mainly affecting the vibrational tem- two nozzles generating Mach 4 flows but at two differ- peratures and strongly decreasing the electron density. ent static pressures: 8 Pa and 71 Pa. These nozzles are chosen to study independently the impact of the Mach number and the impact of the pressure on the shock wave modified by the EHD actuator. In the range of the discharge current considered in this experimental work it is observed that the shock wave angle increases with the discharge current of +15% for the Mach 2 flow [2] but the increase rate doubles to +28% for the Mach 4 flow at the same static pressure showing that the discharge effect is even more significant when boosting the flow speed. When studying the effect of the dis- a b charge on the Mach 4 flow at higher static pressure, it Fig.2: Argon plasma flow around the blunt body is observed that the topology of the plasma changes without magnetic field (a) and with magnetic field drastically and the increase in the shock wave angle (b). with the discharge current is of +21%. This leads us to [1] R. Joussot, S. Coumar and V. Lago (2015), Aero- believe that the plasma actuator can cover a large spec- spaceLab, n°10. trum of the entry process from at least 140 km to 67 [2] S. Coumar et al. (2016), International Journal of km of altitude (which are the studied altitudes) in the Numerical Methods for Heat & Fluid Flow, vol. 26. aim of decelerating the spacecraft. However, to use [3] V. Lago and E. Tinon (2012), 18th AIAA/3AF In- efficiently this actuator, it is necessary to understand ternational Space Planes and Hypersonic Systems and that the phyiscs behind a same phenomena is different Technologies Conference. according to the ambient pressure. Heatshields : techno push-up for preparation of the future exploration missions Y.Mignot1, V.Debout1, J.Edaine1, F.Rozières2

(1)Airbus Safran Launchers Rue du General Niox, 33166 Saint Médard en Jalles (France) Email: [email protected]

(2)Airbus Safran Launchers Route de Verneuil, 78130 Les Mureaux (France) Email: [email protected]

Abstract: On the 19th of March, 2016, 3 days after its jettisoning from the TGO orbiter, the Exomars 2016 Schiaparelli module started its Entry into Mars atmosphere after a 7-month-cruise.

Although a problem of saturation of the Inertial Measurement Unit was encountered shortly after parachute opening, causing an erroneous assessment of altitude by GNC during the Descent phase and impairing the mission below 3.7km, valuable flight data could be transmitted to Earth. On that basis, many subsystems, among which the Heat- shield, can today claim to have reached a TRL 9. The 2020 mission will obviously benefit from the lessons learned.

It is hoped that Europe will continue to give itself the means to fullfil its ambitions wrt Mars exploration beyond 2020 by lauching further probes with the mid-term objective of performing sample return missions. Despite R&T shrinking budget allocated to products with low RoI, a reasonable financial effort is needed within space companies to assess emerging technologies likely to be beneficial for interplanetary probes whether in terms of performance (mass, reliability…), or in terms of manufacturing cycle and cost reduction. In that repect, AIRBUS SAFRAN LAUNCHERS will present on-going in-house initiatives aiming indeed at preparing the future of heatshields. .

Heat-shield for Extreme Entry Environment Technology (HEEET) Development Status

Ethiraj Venkatapathy1, Don Ellerby1, Peter Gage2, Matt Gasch1, Cole Kazemba3, Milad Mahzari1, Keith Peterson4, Carl Poteet5, Mairead Stackpoole1, and Zion Young1,

The Heat shield for Extreme Entry Environ- The heat-shield assembly is shown in Figure 1. One of ment Technology (HEEET) Project is a NASA STMD the significant challenges reported in the last IPPW-13 and SMD co-funded effort. The goal is to develop and is the design, development and testing of the seam. mission infuse a new ablative Thermal Protection Sys- HEEET material development and the seam concept tem that can withstand extreme entry. It is targeted to development have utilized unique test capabilities support NASA’s high priority missions, as defined in available in the US. The various test facilities utilized the latest decadal survey, to destinations such as Venus in thermal testing along with the entry environment for and Saturn for in-situ robotic science missions. Entry Saturn and Venus missions are shown in Figure 2. into these planetary atmospheres results in extreme The HEEET project is currently in it’s fourth heating. The entry peak heat-flux and associated pres- year of the five year development. This continuing sure are estimated to be between one and two orders of presentation will cover both progress that has been magnitude higher than those experienced by Mars Sci- made in the HEEET project and also the challenges to ence Laboratory or Lunar return missions. In the re- be overcome in manufacturing, integration and full cent New Frontiers Announcement of Opportunity, scale (1 mscale ) testing. At this juncture in its devel- NASA is offering HEEET as a new NASA developed opment, a review and assessment of the requirements technology and guarantees delivery of it at TRL 6. and key performance parameters planned vs what we NASA will provide an increase in funding to the PI are able to achieve will illustrate the extent of the suc- managed mission cost (PIMMC) for investigations cess. utilizing HEEET up to $20M. In addition, NASA will limit the risk assessment to only their accommodation Reference: on the spacecraft and the mission environment [1]. This commitment by NASA requires HEEET development to be matured sufficiently to achieve mission (1) http://newfrontiers.larc.nasa.gov/announceme infusibility by FY’2018. In addition to New Frontiers- nts.html 4 mission proposals, ESA M5 mission proposal to Sat- urn has expressed interest in and has baselined HEEET as a required technology. The HEEET ablative TPS utilizes 3D weaving technology to manufacture a dual layer material architecture. The 3-D weaving allows for flat panels to be woven. The dual layer consists of a top layer designed to withstand the extreme external environment while the inner or insulating layer is designed to achieve low thermal conductivity, and it keeps the heat from conducting towards the structure underneath. Arc jet testing combined with material properties have been used to develop a thermal response model for HEEET. A 40% mass efficiency is achieved by the dual layer construct compared to carbon phenolic for a broad range of missions. The 3-D woven flat preforms are molded to Figure 1. HEEET panels and seams integrated achieve the shape as they are compliant and then resin on to a structure has to be demonstrated to with- infused and cured to form rigid panels. These panels stand > 5000+ W/cm2 peak heating and stagnation are then bonded on to the aeroshell structure. Gaps peak pressures greater than 6 atmospheres. exist between the panels and these gaps have to be filled with seams. The seam material then has to be bonded on to adjacent panels and also to the strcuture.

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Figure 2. The numerous facilities, both arc jet and laser heated (LHEML) where tests have been conducted to assess the performance of and understand failure mode if any of the dual layer, 3-D woven, ablative TPS material along with seam concepts. These tests include stagnation point coupons as well as testing flat panels in a wedge configuration to test at combined (heat-flux, pressure and shear).

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2

3 HIAD-TDM CONCEPT DEVELOPMENTS. J. M. DiNonno1, F. M. Cheatwood1, S. J. Hughes1, R. A. Dillman1, R. J. Bodkin1, R. K. Johnson1, K. M. Somervill1, B. R. Hollis1, M. C. Lindell1, C. H. Zumwalt1, J. G. Reed2 1NASA Langley Research Center, Hampton, VA 23681, USA, [email protected], F.M.Cheat- [email protected], [email protected], [email protected], [email protected], [email protected], [email protected], [email protected], [email protected], [email protected]; 2United Launch Alliance, Centennial, CO 80112, USA, [email protected]

Introduction: Hypersonic Inflatable Aerodynamic provides an updated overview of the HIAD-TDM ex- Decelerator (HIAD) is a versatile technology that is po- periment, identifying developments since those pre- tentially applicable to any destination with an atmos- sented at IPPW-13 [2]. phere. By providing significantly more drag area than would otherwise be available to an entry vehicle with a References: rigid heat shield that must fit within a launch vehicle [1] Reed, J. G., Ragab, M. R., et. al., Performance shroud, a HIAD can efficiently decelerate much more Efficient Launch Vehicle Recovery and Reuse, AIAA mass and reduce peak heat flux by decelerating more in 2016-5321. the upper reaches of the atmosphere as compared to cur- [2] DiNonno, J. M., Cheatwood, F. M., et. al., HIAD rent rigid aeroshell technology. While HIAD technol- on ULA (HULA) Orbital Reentry Flight Experiment ogy can enhance, and even enable, larger missions to Concept, IPPW-13, June 2016. higher elevations at Mars, it can also be applied at Earth, such as providing capability for International Space Sta- tion (ISS) down-mass, or even enabling return for free flying orbital micro-g manufacturing. Over the past several years, the HIAD team has investigated a number of potential commercial applications to recover spent launch vehicle assets for reuse, thus lowering overall cost of access to space. The United Launch Alliance (ULA) Sensible, Modular, Autonomous Return Tech- nology (SMART) concept utilizes a HIAD to enable recovery of the Vulcan launch vehicle booster main engines [1]. HIAD technology readiness must be matured beyond its current state to buy down risk for such future applications, and this maturation requires an orbital ve- licity entry flight experiment. Building upon the successful 3m diameter Inflatable Reentry Vehicle Experiment (IRVE) flights and extensive ground testing and development efforts at various scales, the HIAD project team at NASA Langley Re- search Center (LaRC) is formulating a scaled up orbital velocity atmospheric entry experiment. The concept is to conduct an entry flight experiment with a 6m diameter HIAD to significantly advance the current state of the art HIAD materials and construction techniques. The approach is to partner with ULA to conduct the experiment as a hosted secondary payload, taking advantage of available mass and volume capacity within the payload adapter on an Atlas V launch vehicle. The proposed flight experiment, previously called HIAD on ULA (HULA) and now referred to as HIAD Technol- ogy Demonstration Mission (HIAD-TDM), will demonstrate the HIAD aeroshell performance, and collect invaluable data in environments that are relevant to a variety of HIAD mission applications. This presentation EUROPEAN STUDIES TO ADVANCE DEVELOPMENT OF INFLATABLE AND DEPLOYABLE HYPERSONIC DECELERATORS. Samuel J. Overend1, John C. Underwood2, J. Stephen Lingard3, H. Ritter4, L. Ferracina5, E. Johnstone6, B. Esser7 1Vorticity Ltd, Chalgrove, Oxfordshire, OX44 7RW, United Kingdom, [email protected], +44 1865 413421 2Vorticity Ltd, Chalgrove, Oxfordshire, OX44 7RW, United Kingdom, [email protected], +44 1865 413420 3Vorticity Ltd, Chalgrove, Oxfordshire, OX44 7RW, United Kingdom, [email protected], +44 1865 413423 4European Space Agency, Noordwijk, The Netherlands 5ATG Europe B.V. on behalf of the European Space Agency, Noordwijk, Netherlands 6Fluid Gravity Engineering, Ltd, Emsworth, Hampshire, PO10 7DX, UK 7Deutsches Zentrum für Luft- und Raumfahrt (DLR), Linder Höhe, 51147 Köln, Deutchland

Introduction: This paper will discuss the out- qualification. One of the key technologies required is a comes of studies performed under the ESA Technolo- flexible thermal protection system (TPS). This must be gy Research Programme (TRP) in order to advance the packed within the fairing and then deployed or inflated development of inflatable and deployable aerodynamic once the spacecraft arrives at its destination. Testing decelerators for future European missions. Particular has been conducted in the DLR L2K arc heater facility focus will be given to material testing of candidate on European materials, to determine their suitability flexible thermal protection system (TPS) materials in for use in constructing a flexible TPS layup. These the DLR L2K arc jet heater facility. results show that Cerafib Tex99 is a candidate for a Background: With the recent NASA MSL and lightweight high-heatflux protection, and Insulair ESA ExoMars 2020 mission currently in development, NP650 Plus offers similar performance to Pyrogel XT- the limits of what can be achieved with conventional E. Microtherm Slimflex was also considered since it EDL systems are being reached due to size constraints appeared to offer very good performance, but it was imposed on the heatshield by the launch vehicle fair- difficult to construct the necessary small scale test ing. For current and planned launch vehicles, a maxi- samples from this material. mum diameter of about 4.5 m can be accommodated, References:: which places a hard limit on the mass which is compat- [1] Adler, M, et al., et al. DRAFT Entry, Descent and ible with a realistic ballistic coefficient. The ballistic Landing Roadmap - Technology Area 09. s.l. : NASA, 2010 coefficient of the entry vehicle directly affects the aerothermal heating and trajectory of the vehicle and consequently limits the landed mass and altitude of the landing site. Lifting entries, as employed on MSL, can improve the situation to a limited extent but NASA have estimated that the MSL EDL system could only be extended to a maximum landed mass of 1500 kg from the 900 kg achieved[1]. To enable future missions with higher landed mass requirements such as Sample Return or manned precursor flights, or for missions to higher altitude sites of scientific interest, a new EDL approach is required. Advantages of deployable and inflatable decelerators: Through deployment or inflation of additional drag area, the overall heatshield area is increased and the ballistic coefficient of the entry vehicle is reduced. This drag area requires structure to transmit the increased aerodynamic loads to the entry vehicle and that structure requires thermal protection. Work reported: Work was performed both on identifying candidate technologies and identifying tools and facilities which will be necessary for their DETAILED DESIGN OF EARTH ENTRY VEHICLE FOR COMET SURFACE SAMPLE RETURN. T. R. White1, R. W. Maddock2, C. D. Kazemba3, R. G. Winski4, D. S. Adams5, 1NASA Ames Research Center, Moffett Field, CA, 94035, [email protected] 2NASA Langley Research Center, Hampton, VA, 23681, [email protected] 3STC Corporation, NASA Ames Research Center, Moffett Field, CA, 94035, [email protected] 4Analytical Mechanics Associates, Inc., Hampton, VA, 23681, [email protected] 5Johns Hopkins University Applied Physics Laboratory, Laurel, MD, 20723, [email protected]

Introduction: The 2013 Decadal Survey for New Frontiers missions identifies several high-value science missions, including Comet Surface Sample Return (CSSR) [1]. A CSSR mission will advance the scientific community’s fundamental understanding of the origin of the solar system and the contribution of comets to the volatile inventory of the Earth. An entry capsule, or earth entry vehicle (EEV), is be required to protect the scientific payload from the extreme conditions of atmospheric entry, descent, and landing. The Decadal Survey Mission Concept Study [2], along with an APL 2007-2008 Comet Surface Sample Mission Study [3] details several of the driving requirements for a CSSR EEV; these include a payload volume and mass and inertial entry velocity of ~ 9 km/s. The mission concept study selected a Multi- Mission Earth Entry Vehicle (MMEEV) design concept derived from the Mars Sample Return (MSR) entry capsule design because of its increased reliability over a parachute-based vehicle [4], [5]. This presentation will explore detailed design of a CSSR-capable Earth Entry Vehicle, including trajectories, aeroheating predictions and associated thermal protection system masses, and onboard instrumentation for entry science. References: [1] National Research Council, “Visions and Voy- ages for Planetary Science in the Decade 2013-2022,” The National Academies Press (2011). [2] Veverka, J., Johnson, L., Reynolds, E., “Mission Concept Study: Comet Surface Sample Return (CSSR) Mission,” Pre- pared for the Planetary Science Decadal Survey (2011). [3] The Johns Hopkins University Applied Physics Laboratory, “Comet Surface Sample Return Mission Study,” Prepared for NASA’s Planetary Sci- ence Division (2008). [4] Maddock, R. W., “Multi- Mission Earth Entry Vehicle Design Trade Space and Concept Development Status,” 7th International Plane- tary Probe Workshop (2010). [5] Maddock, R., Hen- ning, A., Samareh, J., “Passive vs. Parachute System Architecture for Robotic Sample Return Vehicles,” IEEE Aerospace Conference (2016). Taking A PEAQ at Uncertain Atmospheres (Atmospheric Prediction for Entry & Aerocapture Qualification) B. Tackett1†, 1Graduate Student email: [email protected], School of Aeronautics and Astronautics, Purdue University, 701 W. Stadium Ave., West Lafayette, IN, 47907-2045.

Introduction: With the latest Planetary Decadal risks associated with aerocapture is the single point Survey, Visions and Voyages published for 2013 - nature of the maneuver which makes it very 2022 [1], there has been a substantial increase in susceptible to uncertainties in aerodynamics, attention to the outer planets and moons of the solar atmospheric properties and initial trajectory system. For example, as stated in the decadal survey, properties. If an error were to occur during the third-highest-priority flagship mission is a Uranus aerocapture, propellant would need to be used to fix Orbiter and Probe mission. Although Neptune was the resulting trajectory in the best case or the mission also investigated, Uranus was more favored for the may fail completely in the worst-case scenario where current decade for reasons associated with available there is not enough propellant onboard to correct for trajectories, flight times, and cost. Delving deeper into the error. As such, minimizing these risk areas the survey, aerocapture is shown to be a very important (aerodynamics, atmospheric properties and initial topic for exploration of the outer planets and will trajectory) will significantly increase the viability of likely be part of any mission to Uranus or Neptune. aerocapture for future missions. As an example, While aerocapture is a very attractive technology for NeptuneGRAM predicts mean density profiles for the outer planet exploration, it does pose certain risks due Neptunian atmosphere which shows an error of up to to a lack of knowledge into the atmospheres of other +239% and -70% in atmospheric density in the range planets and its effects on the aerocapture trajectory. At of aerocapture altitudes[2], [3]. the moment, this risk is far too large to send a mission to Uranus or Neptune which depends on aerocapture and we cannot simply test aerocapture in appropriate conditions without a significant investment into leader missions. If this is the case, then the only option is traditional, costly, propellant maneuvers to slow down and circularize the orbiter around the planet, right?

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The possibility of encountering these uncertainties, even with a likely statistical mean density, is completely unacceptable in terms of aerocapture. This can lead to a massive difference between expected orbital trajectory and the actual resulting trajectory or )" ! $%$%# even descending too far into the atmosphere where The objective of this design is to determine a aerocapture becomes impossible. solution to decrease the risk of aerocapture in mostly Road Map – Present, Future: Currently, the unknown atmospheres. This will help to make closest we have come to aerocapture is aerobraking aerocapture a more attractive technology particularly which involves performing similar maneuvers to with aerocapture being a technology incentive posed aerocapture, but at a much smaller scale over long by the Visions and Voyages Decadal Survey. While periods of time to decrease the orbital ellipse radii to a this design may apply to multiple cases and planets, circular or near circular orbit. Conversely, aerocapture this abstract will focus predominantly on Uranus and seeks to decrease hyperbolic velocities down to orbital Neptune for simplicity and applicable data. velocities using a single pass through the atmosphere Call to Action: The ability to confidently perform followed by propulsive maneuvers for positioning and aerocapture around a planet has been plagued by the correcting any errors that occur. However, the risk has risks associated with the maneuver. One of the primary still been shown to be much too high to use such an approach for a mission due to the cost associated with predict the atmospheric properties of the planet of failure[4]. In hopes of alleviating some of this risk, interest. guidance systems have been proposed which use lift to The When and the Where: An interesting thing reduce the error caused by atmospheric uncertainty. about the outer planets is how their atmospheres These guidance systems have been shown to overcome maintain temperature during an orbital period. For this error significantly, but are heavily dependent on example, for a multi-year span of time one of the poles the density of the atmosphere being known relatively of Uranus is pointed directly at the Sun and on the well to make maneuvers. If we want to use aerocapture opposite end of its orbit, the other pole is facing the in the future for outer planets, we need some way to sun for almost the same amount of time[5], [6]. analyze the density of the atmosphere of interest prior Also, since these outer planets are so far from the to depending on it for aerocapture. Sun, their internal heating has a significant Revealing the Truth: To analyze the atmospheric contribution to the temperature of the atmosphere and properties of a planet of interest, this design poses a leader probe that will measure data during flight and relay it back to the main vehicle. Ideally, this data would tell us everything we need to know from composition to general gas properties, but this would require many instruments to measure and record the data for use, not to mention a large or powerful antenna to relay it to the main vehicle. However, since all heating analysis will have been performed previously with no way to change the heatshield midflight, the predominant atmospheric characteristic )" ( $%&+' for aerocapture is density. Density uncertainty propagates through the aerodynamics of the vehicle therefore it’s atmospheric density. Even so, very little and quickly leads to large error in trajectory simulation work has been published or is being worked on and prediction. So, to measure density while satisfying regarding the consistency of outer planet atmospheres our objective of a low mass system, it appears to be over time, like whether they are relatively the same most beneficial to measure density indirectly using from year to year, day to day or even hour to hour. This sensed deceleration. This sensed deceleration can estimation on atmospheric consistency comes into quickly be used to reconstruct the trajectory as well as play with leader probes such as this because it the atmospheric density encountered by the leader determines how far in advance a probe can be sent and probe which can then be used to great effect by the still provide relatively accurate data back to the main main vehicle. vehicle. This has been an interesting point of Blueprints: In theory, the optimal leader probe investigation along with the corresponding orbital would consist of a scaled down version of the actual points associated with the most consistent atmospheric probe to mirror the expected flight properties properties over time. experienced during aerocapture. However, this would References: most likely be too expensive or too complex which [1] Planetary Science Decadal Survey Committee goes against the objective of this design. It would be (2013) Visions and Voyages for Planetary Science in beneficial to mirror the most important aerodynamic the Decade 2013-2022, 18-311. [2] M. K. Lockwood, and physical parameters associated with the main (2004) Neptune Aerocapture Systems Analysis, 1–16. aerocapture vehicle, i.e. Ballistic Coefficient, L/D, [3] Justus C. G. et al. (2004) Atmospheric Models for heat shield shape, aerodynamic center and center of Aeroentry and Aeroassist, NASA/CP-2004-213456, mass most importantly. With these properties, the 41–48. [4] Wercinski P. et al. (2002) Aerocapture leader probe could theoretically fly a similar trajectory Technology Development Needs for Outer Planet as the main vehicle to emulate the intended trajectory Exploration. NASA/TM-2002-211386 [5] NASA before the main vehicle arrives. Beyond the outer shell Planetary Science Division, Uranus - In Depth | of the leader probe, it will include an antenna, a power Planets - NASA Solar System Exploration. [6] supply, accelerometers and a small computing system Williams D. R. (2016) Uranus Fact Sheet. [7] Image to measure, record and relay the atmospheric from setterfield.org/Astronomy. information to the main vehicle which can be used to DEVELOPMENT STATUS OF A MARS ASCENT VEHICLE TECHNOLOGY DEMONSTRATION. Joel Benito and Robert Shotwell, Jet Propulsion Laboratory, California Institute of Technology, 4800 Oak Grover Dr., Pasadena, CA 91109, USA. [email protected], [email protected].

Introduction: A Mars Ascent Vehicle (MAV) is a References: conceptual launch vehicle that would put into Mars orbit [1] Shotwell, R. et. al., “Drivers, Developments and an Orbiting Sample (OS) containing rock, soil and at- Options under Consideration for a Mars Ascent Vehi- mosphere samples taken from the surface of Mars. A cle,” IEEE Aerospace Conference, 2016. MAV would be the first vehicle to launch from extrater- [2] Shotwell, R., et. al., “A Mars Ascent Vehicle for restrial soil in the presence of an atmosphere and with Potential Mars Sample Return,” IEEE Aerospace Con- substantially greater gravity than previous lunar ascent ference, 2017. vehicles. A MAV would be part of potential Mars Sam- ple Return, which could include a prior mission to collect the samples, and a posterior mission to rendezvous with the OS and bring it to Earth for analysis [1]. A terrestrial MAV technology demonstration in 2019 is under consideration to advance the Technology Readiness Level of some of the key systems that form the vehicle [2]. The most important technology to mature is hybrid propulsion, which provides the advantages of liquid propulsion (throttleability, re-starta- bility) without some of its drawbacks like thermal sensitivity; an important advantage for deep-space missions. The Guidance, Navigation and Control functions will be computed on a Sphinx CubeSat avionics system, modified for MAV; a lightweight alternative to traditional deep space alternatives. The vehicle and flight path control will be performed using a cold gas reaction control system and a liquid injection thrust vector control (LITVC) system. LITVC was selected due to its low mass and simplicity compared to a gimbaled nozzle. Demonstrating flight control using LITVC and validating the predicted performance is one of the main objectives of the flight demonstration. The flight demonstration will as well be used to validate and anchor the aerodynamic and aerothermal models, as well as the structural performance predictions. In order to fly an aerodynamic profile representative of the Mars trajectory, the MAV demo vehicle will be launched from a high altitude balloon, at an altitude around 30 km above sea level. Unlike on Mars, the MAV demo will not reach orbit, but will climb to vacuum conditions to perform model characterization maneuvers. Overall, this demonstration is planned to test and validate many of the techonologies proposed for a Mars application, and is an important step toward the development of a MAV concept for a potential Mars sample Figure 1. A Mars Ascent Vehicle concept. return mission.

MISSION AND GNC CHALLENGES FROM IXV TO SPACE RIDER. R. Haya-Ramos1 , A. Ayuso1, D. Bonetti2, M. Kerr2 and E. Di Sotto3 1SENER Ingeniería y Sistemas, S.A. (Severo Ochoa 4, PTM, 28760 Tres Cantos, SPAIN [email protected]), 2DEIMOS Space (Ronda de Poniente 19, 28760 Tres Cantos, SPAIN). 3GMV (Isaac Newton 11, PTM, 28760 Tres Cantos, SPAIN).

Introduction: The Intermediate eXperimental Ve- per stage of the launcher as service module. It will re- hicle (IXV) is an ESA re-entry lifting body demonstra- turn to Earth performing a guided lifting re-entry up to tor built to verify in-flight the performance of critical the triggering of the parachute system in subsonics. re-entry technologies needed for the return from Low Precision ground landing will be achieved by the de- Earth Orbit (LEO) and for Exploration. The IXV was ployment and operation of a guided parafoil during the launched on February the 11th, 2015, aboard Europe’s descent phase. Afterwards, the vehicle will be refur- Vega launcher. The IXV´s flight and successful recov- bished for the next flight. ery represents a major step forward with respect to The return module has the same external shape as previous European re-entry experience in terms of IXV. It is a 5 m long lifting body with a lift-to-drag large reentry systems. IXV demonstrated technologies ratio of 0.7 in hypersonics and it is actuated through like Thermal Protection System (TPS) and the Guid- the combination of two body elevons mounted at the ance, Navigation and Control (GNC). aft windward side of the vehicle and a Reaction Con- Based on the IXV success and assets, an applica- trol System (RCS) mounted in the rear part. tion programme called Space Rider has been proposed to develop an affordable and sustainable reusable Eu- ropean space transportation system to enable routine “access to” and “return from” space, operating in-orbit, de-orbiting, re-entering, landing on ground and being re-launched after limited refurbishment. Space Rider will perform in-orbit operation, experimentation and demonstration for applications like micro-gravity experimentation, orbit applications and In-Orbit Demon- stration and validation of technologies. These technologies suitable for demonstration inside Space Rider cover a wide spectrum: from Earth science to planetary Exploration. The re-entry module itself is a test bed for entry technologies as the IXV precursor was. The project is currently running Phase B1 and heading towards Phase B2 by end of the 2017. Figure 1: Space Rider re-entry module Shape The evolution from the IXV system to the Space Rider System presents multiple synergies, like the re- Mission Analysis: the main challenges for the mis- use of the same aerodynamic shape for re-entry but sion analysis and flight mechanics are related with the also clear differences like the 2 months duration in flexibility in terms of target orbit and landing site, the orbit that will challenge the design and qualification of longer mission duration, new mission phases (e.g. or- the new system. This paper will address the synergies bital, transonics…), the heavier vehicle and new safety and challenges in terms of Mission Analysis and the aspects (e.g. landing on ground close to populated are- Guidance, Navigation and Control system starting from as). Based on IXV qualified methodologies, an exten- the solid grounds of the IXV mission. Analysis and sion is required to ensure the design is end-to-end cap- results used to capture and consolidate the mission and turing the applicable requirements. system needs will be presented with focus on the re-use GNC: the increased vehicle functionality com- and application for Exploration. The following para- pared to IXV translates into an increase of the number graphs summarise the main aspects and drivers that of GNC functions needed to meet the mission and sys- will be addressed in the paper. tem objectives, like the AOCS or the rendezvous GNC. Mission and System overview: Space Rider will On the other side, the operational nature of the mission be launched from Kourou using Vega-C and injected stresses all subsystems in terms of reliability, verifica- into orbit up to the ISS inclinations and altitudes. It tion, validation and resusability and GNC is not an will stay in orbit a minimum of 2 months using the up- exception. During the atmospheric entry the IXV GNC solution and flight heritage will be exploited and tailored to the specific mission needs. The main difference comes from the extension of the flight to cross the transonics in order to deploy the parachute in subsonics. As a relevant new element, pinpoint landing is achieved with a guided parafoil GNC which will operate in order to steer the vehicle towards the landing aerodrome with an accuracy < 150 m. Exo-Brake Drag Modulation Flight Experiment Results M.Murbacha*, A. Guarneros-Lunaa, R. Alenaa, A. Dono Perezb, A. Tannerb, J. Whelessb, S. Smith, A.Salas, C. Priscalc J. Pleaterc, A. Dwyer Cianciolod, R. Powelld .a NASA, Ames Research Center, Moffett Field, CA, U.S. 94035, b Millennium Engineering & Integration Company, NASA Ames Research Center, Moffett Field, CA, U.S. 94035 cStinger Ghaffarian Technologies, Inc, NASA Ames Research Center, Moffett Field, CA, U.S. 94035 dNASA Langley Research Center, Hampton, VA, US 23681

Introduction: The first of a series of drag-modulation experiments using the Exo-Brake was jettisoned from the International Space Station (ISS) on March 6, 2017. The TechEdSat-5 nano-satellite deployed the Exo-Brake through an automated 30 minute timer at an altitude of 410km. Over the course of the following weeks, the trajectory was monitored via TLE (Two Line Element) sets and on-board GPS (Global Posi- tioning System). Local attitude determination was ascertained via a unique Wireless Sensor Module and two cameras. Daily values of the F10.7 and geomag- netic variables were included in the Thermosphere models, and various propagators were used to develop experimental targeting methodologies. Results of the targeting efforts, as well as subsequent experiments, are discussed. The drag-modulated Exo-Brake may offer a convenient de-orbit system for future Earth orbit sample return, TPS/hypersonic experiments, as well as planetary exploration applications.

Technology Investments in the NASA Entry Systems Modeling Project. M. D. Barnhardt1, M. J. Wright1, and M. Hughes2, 1NASA Ames Research Center, Moffett Field, CA 94035, 2NASA Langley Research Center, Hampton, VA 23666.

Introduction: The Entry Systems Modeling Heating from shock layer radiation for many mis- (ESM) technology development project, initiated in sions is a large contributor to thermal protection sys- 2012 under NASA’s Game Changing Development tem (TPS) material selection and sizing. Experimental (GCD) Program, is engaged in maturation of funda- facilities and computational capabilities are being used mental research developing aerosciences, materials, to develop and validate new models and databases. The and integrated systems products for entry, descent, and current focus is on understanding radiation for entries landing (EDL) technologies [1]. To date, the ESM pro- to Mars and Venus as well as high-speed return to ject has published over 200 papers in these areas, com- Earth. The most detailed spectroscopic databases in the prising the bulk of NASA’s research program for EDL world are being assembled from fundamental meas- modeling. This presentation will provide an overview urements obtained in the Electric Arc Shock Tube [9] of the project’s successes and challenges, and an as- (EAST) coupled with quantum calculations that give sessment of future investments in EDL modeling and unprecedented insight into non-equilibrium processes. simulation relevant to NASA’s missions. The new databases feed directly into production radia- Aerosciences: The Aerosciences technical area is tion codes, substantially improving their accuracy. One tasked with the development and validation of simula- such code, NEQAIR [10], was named NASA’s Soft- tion tools for hypersonic EDL, making revolutionary ware of the Year for 2015. At the same time, university improvements to the current state-of-the-art (SoA), partnerships are being leveraged to explore a new par- with a focus toward increased reliability, reduced un- adigm in aerothermodynamic modeling which aims to certainty, and enabling new technologies for reentry fully couple radiation and fluid dynamic processes. applications. The three core research elements within The tools developed are being used to better under- Aerosciences are Advanced Computational Fluid Dy- stand data returned from two radiometers flown on namics (CFD), Shock Layer Radiation, and EDL Mod- Exploration Flight Test 1 in December 2014 [11], in el Validation. partnership with the Orion Program, as well as the in- Advanced CFD is devoted to developing next- strumentation of the Mars Science Laboratory heat- generation computational capabilities in order to en- shield [12]. The project is also renovating the Low hance predictive accuracy of aerothermodynamics Density Shock Tube, which may be able to provide modeling software and to fill modeling gaps identified unique insight into non-equilibrium radiation and envi- by current flight projects and the broader technical ronments relevant for low ballistic coefficient entry community. Two codes, US3D [2] and FUN3D [3], systems, and is building international collaborations were developed and released NASA-wide in 2014 to with European, Asian and Australian radiation model- replace the widely-used Data Parallel Line Relaxation ing groups. [4] (DPLR) and Langley Aerothermodynamic Upwind EDL Model Validation is conducting focused Relaxation Algorithm [5] (LAURA) software packag- ground testing to enable validation of aerothermal es. US3D and FUN3D enable greater geometric com- models in the other core areas. The Large Energy Na- plexity and introduce many advances in physical and tional Shock expansion tunnel (LENS-XX), part of the numerical modeling to enhance accuracy. For non- CUBRC wind tunnel research complex, is being used continuum modeling, further investment was made in partnership with NASA’s Space Technology Re- toward development of a new Direct Simulation Monte search Grants (STRG) Program to conduct tests in high Carlo (DSMC) code called Multiphysics Algorithm enthalpy CO2 in order to understand previously ob- with Particles [6] (MAP), in addition to fundamental served discrepancies in predicted shock standoff and improvements to the current SoA in DSMC modeling. heating in reflected shock tunnels [13]. The High En- Current technology development in the Advanced CFD thalpy Tube (HET) facility at CalTech is being used in research area include: Free-flight CFD modeling [7] partnership with NASA Space Technology Research (as seen in Fig. 1) and magnetic suspension wind tun- Fellowships to provide second source data. Heating nel development for vehicle stability prediction; an augmentation due to distributed roughness was inves- adaptive mesh capability for US3D; advanced numeri- tigated by testing in the Langley Mach 6 wind tunnel cal algorithm development suitable for capturing com- [14] and Ames ballistic range [15]. The data are being plex, unsteady phenomena [8]; and modeling of large- used to update roughness correlations and turbulence scale parachute dynamics. models for blunt bodies in air and CO2. address the increasing scale and complexity of current and future NASA missions. Future directions of research will pursue development of a material analysis framework geared toward computational design of woven materials, a breakthrough that could greatly reduce material development costs and enable tailored solutions for entry vehicles. Advanced TPS Materials Concepts contains three product lines. The aim of the advanced ablator development effort is to establish, through the use of fibrous felt systems and high performance charring resins, robust conformal thermal protection systems for large area heat shield applications. When matured, advanced conformal TPS systems should greatly reduce manu- Fig. 1: Dynamic simulation of Low Density Su- facturing costs and improve thermostructural reliability personic Decelerator ballistic range test. of ablative TPS, while simultaneously maintaining or improving the thermal performance over the current Thermal Protection Materials: The Thermal SoA. The flexible thermal protection materials (FTPS) Protection Materials technical area focuses on develop- effort is developing high performance thermal insula- ing software and hardware that will make revolution- tors and refractory cloth systems that provide low ary improvements to the current SoA and provide ena- weight and volume thermal protection solutions for bling capabilities in thermal protection materials and inflatable decelerator designs exposed to extreme entry the high-fidelity models used in their design and analy- environment heating. ESM has extended the usable sis. The technology development effort has two ele- range of this material class to 75 W/cm2, making in- ments: Advanced Ablator Thermal Models and Ad- flatable decelerators a viable architecture for a variety vanced TPS Materials. of robotic and human-scale entry missions. Finally, the Advanced Ablator Thermal Models is focused on Convective Heating Improvement for Emergency Fire developing and validating a high-fidelity volumetric Shelters (CHIEFS) effort is using materials and con- ablation computational capability that can be used to cepts developed under FTPS to design and test ad- support materials development and uncertainty analy- vanced fire shelters for use by the United States Forest sis through a rigorous examination of the physical and Service and other agencies. An image of the CHIEFS chemical processes of ablation at a fundamental level. material layup is shown in Fig. 3. Early testing indi- The centerpieces of this effort are the Pyrolysis Abla- cates that FTPS derived shelters are providing major tion Toolkit in OpenFOAM [16] (PATO) and Porous improvements to the current SoA. Media Analysis [17] (PuMA) codes. PATO employs a high-fidelity macroscopic material model suitable for exploring advanced research applications relevant to thermal protection materials and for creating reference simulations to guide development of engineering- fidelity tools. PuMA utilizes particle-based methods in conjunction with high-resolution micro-tomography data to compute morphological, thermal, and transport properties of porous materials. An example of PuMA applied to carbon fiber microstructure is shown in Fig. 2. Model validation experiments are being conducted at several universities in partnership with Early Stage Innovation (ESI), as well as within NASA and at the Lawrence Berkeley National Laboratory Advanced Light Source (cyclotron) [18]. Ablation response has also been coupled with CFD, which has improved the predictive capability of NASA codes in this highly Fig. 2: PuMA-derived analysis of thermal con- nonlinear flight regime. The lessons learned in the de- duction through carbon fiber microstructure. velopment of the volumetric models are being used to develop a new engineering design tool, Icarus [19], to

Fig. 3: Image of CHIEFS material layup derived from flexible TPS material research. Fig. 4: Depiction of deployed Exo-Brake at-

tached to cubesat for controlled payload return. Integrated EDL Systems: The primary technical objective of the Integrated EDL systems technical area is to foster game-changing EDL or EDL-enabling con- References: cepts at the system level via either subscale flight or [1] Wright, M. J. et al. (2015) AIAA Paper 2015- ground testing. The primary products of this technical 1892. area are ground test articles and flight hardware. The [2] Candler, G. V. et al. (2015) AIAA Paper 2015- technical area currently consists of a single element: 1893. Exo-Brake. [3] Gnoffo, P. A. et al. (2013) AIAA Paper 2013- The Exo-Brake [20], depicted in Fig. 4, is a con- 2558. cept for propellentless deorbit from Low Earth Orbit [4] Wright, M. J., White, T., and Mangini, N. (LEO) via a controllable drag sail device. The concept (2009) NASA TM-2009-215388. has been previously demonstrated at the cubesat level. [5] Mazaheri, A. et al. (2010) NASA TM-216836. The current effort intends to demonstrate control of the [6] Liechty, D. S. (2015) JSR, 52, 1521-1529. landing ellipse through various methodologies, includ- [7] Brock, J. M., Stern, E. C., and Wilder, M. C. ing drag modulation, timed release, and predictor- (2017) AIAA Paper 2017-1437. corrector guidance techniques. ESM will fly two mis- [8] Gnoffo, P. A. (2015) AIAA Paper 2015-2756. sions: TechEdSat-5, currently on the International [9] Brandis, A. M. and Cruden, B. A. (2017) AIAA Space Station awaiting a February 2017 deployment, Paper 2017-1145. and TechEdSat-6 scheduled for late 2017. Using an [10] Brandis, A. M. and Cruden, B. A. (2016) iterative design and flight test cycle allows ESM to NASA ARC-E-DAA-TN19420. rapidly mature the technology for potential use in [11] Cruden, B. A. and Johnston, C. O. (2016) small satellite return from LEO, debris deorbit, and AIAA Paper 2016-4113 (ITAR). netlander missions to Mars or other planets. [12] Cruden, B. A. et al. (2016) JTHT, 30, 642-650.

[13] Hollis, B. R. et al. (2017) JTHT, January

online ed. [14] Hollis, B. R. (2014) AIAA Paper 2014-0238. [15] Wilder, M. C., Reda, D., and Prabhu, D. K. (2015) AIAA Paper 2015-1738. [16] Lachaud, J. and Mansour, N. N. (2014) JTHT, 28, 191-202. [17] Ferguson, J. C et al. (2016) Carbon, 96, 57-65. [18] Panerai, F. et al. (2017) Int. J. Heat Mass Transfer, 108, 801-811. [19] Schulz, J. C. et al. (2017) AIAA Paper 2017- 0667. [20] Dutta, S. et al. (2017) AIAA Paper 2017-0467. Development of the European Conformal Ablative-charring material as an extension of the ASTERM family in the frame of the ESA DECA TRP.

J. Bertrand1, G. Pinaud1, M. Desbordes1, JM. Bouilly1, J. Barcena2, B. Esser3, G. Verkinis4,

(1)Airbus Safran Launchers Rue du General Niox, 33166 Saint Médard en Jalles (France) Email: [email protected]

(2)Tecnalia Industry and Transport Division Mikeletegi Pasealekua 2, E-20009 Donostia-San Sebastián (Spain)

(3) Deutsches Zentrum für Luft- und Raumfahrt e.V. German Aerospace Center (DLR) Institute of Aerodynamics and Flow Technology, Supersonic and Hypersonic Technologies Linder Hoehe, 51147 Cologne (Germany)

(4) NCSR Democritos, Neapoleos str., 153 4, Agia Paraskevi, Attica (Greece).

Abstract: The state of the art thermal protection ESA's TRP, a conformal ablator material was manufac- material for high speed Earth Return Capsule is un- tured and deeply characterized. The objective of this doubtedly light weight ablator based on carbon fibre activity was to mature heatshield material technology impregnated by phenolic resin. The most famous ex- based on a conformal carbon felt, and thus to extend ample is the PICA material (Phenolic impregnated ASTERM family. Carbon Ablator) that flew on the faster man-built Star- The test plan that was performed not only includes dust Capsule. In Europe, such material development laboratory thermo-physical and thermo-mechanical started in 2004 with ASTERM family. ASTERM ther- characterization (mainly performed in Tecnalia labora- mal protection from Airbus Safran Launchers is the tory) but also arc jet plasma campaign both in DLR European reference low density carbon based ablative (stagnation point configuration in L2K) and Airbus material and has been specifically developed for the Safran Launchers (wedge configuration in SIMOUN) most severe Earth entry environment typically encoun- and facilities. This unique plasma test was also the op- tered during sample return missions. portunity to qualify advanced measurement sensors for The standard ASTERM version is using a highly the ablation characterization. porous rigid substrate impregnated with phenolic pol- A numerical thermal and ablative model of the C- ymer. The composite proved to be well suited for 1 m ASTERM material is built by a mix of elementary scale probe where monolithic heat-shield could be di- characterizations and plasma results cross checking. rectly milled from rough cylindrical preform. For larg- The model is mathematically closed by a set of as- er vehicle, heat-shield must be assembled from multi- sumptions and laws (when missing) taken from the ple pieces or tiles since the material shows a brittle standard ASTERM material. behaviour and low strain to failure. Finally, a scale 1 tiled static demonstrator made of Since few years, Airbus Safran Launchers has on standard and conformal ASTERM are bonded on a going activities to extend the ASTERM material family representative CFRP-composite honeycomb sandwich and to mature technology on conformal charring abla- structure to validate bonding and assembly process tor with the goal to tackle this standard version thermo- previously adjusted in laboratory. mechanical limitation. Instead of rigid felt, this new The final goal of the project is to reach TRL 4 on material uses a flexible carbon preform that could be C-ASTERM based thermal protection system. easily shaped by moulding process. After curing, it results in a rigid ablator with potential curved surface. Due to the properties of the felt, it is expected that C- ASTERM will reveal better compliance behaviour with regard to the sub-structure compared to the standard rigid ASTERM. In the framework of DECA (Development of a rigid conformal ablator for extreme heat flux applications) INVESTIGATIONS ON CO2 PLASMA JETS WITH THE ICOTOM RADIOMETERS OF THE EXOMARS DESCENT MODULE. M. Jacquot1, P. Boubert1, A. Bultel1, and J. Annaloro2, 1CORIA – University of Rouen, 76800 Saint-Etienne-du-Rouvray, France, e-mail: [email protected], 2 CNES – 18 avenue Edouard Belin, 31401 Toulouse, France, e-mail: [email protected].

Introduction: The back heat shield of the Exo- The radiative heat flux received by the ICOTOM mars descent module Schiaparelli, that entered the radiometers depends mainly on the dissociation rate of Mars atmosphere on October 19, 2016, was equipped CO2. That parameter controls the temperature of the with three COMARS+ packs located on a radial line. plasma as well as the CO and O densities, and then the Each COMARS+ pack contains a heat flux sensor, a radiative transfer from the plasma to the detectors. So, pressure gauge [1] and two CNES infrared radiometers the analysis of the ICOTOM signal includes the re- called ICOTOM [2]. Each ICOTOM radiometer col- building of that signal through the measurements of lects light from the back entry plasma in a 17.5° half- CO and O densities and temperatures. In order to per- angle cone. Each pair is composed of a radiometer form a mapping of the plasma (assumed to be axisym- operating in the range 4.17 – 5 µm (B1) and a radiome- metric), laser measurements (spontaneous Raman scat- ter operating in the range 2.6 – 3.36 µm (B2). Some of tering and two-photon laser-induced spectroscopy) the radiometers designed for the Exomars missions were conducted within the view cone of the ICOTOM were not used for the mission on board but in laborato- signal. ry wind tunnels in order to help to characterize the Results: The results obtained show a significant emission of CO2 plasmas and to better understand the increase of the ICOTOM signals with pressure espe- Schiaparelli measurements during the entry phase. cially in the lower part of the pressure range. The sig- Calibration: Calibration experiments were carried nals also increase with the global specific enthalpy but out on non-flying ICOTOM in order to use them in very slightly, following the opposite effect of a higher plasma wind tunnels and to study their measuring be- temperature and a higher CO2 dissociation. The ratio haviour with their own temperature. Those measure- B1/B2 strongly decreases with pressure (Fig 1.) while ments were crosschecked with manufacturer's calibra- it slightly decreases with global specific enthalpy. The tion. Manufacturer's data were also post-processed in signals and the ratio stabilize at higher pressures. order to obtain direct flux density information from the ICOTOM electrical signal collected from the ground- based plasmas and to make ready the reception of the in-flight measurements. The experiments carried out with a blackbody-like electrical furnace as infrared source confirmed that the ICOTOM signals decrease with their own temperature and that effect is especially significant for the B1 band. Large differences can then be observed on the output electrical signal depending on the housing temperature. Experiments with plasmas: Embedded in a subsonic inductively-coupled plasma wind tunnel, the Fig. 1 : B1/B2 signal ratio vs pressure radiometers were exposed to CO2 plasma at various for various global specific enthalpies. pressures and global specific enthalpies. The experi- Calculations: Some calculations were carried out ments carried out on the wind tunnel SOUPLIN [3] with CDSD4000 and HITEMP databases at equilibri- allowed to expose the ICOTOM to CO2 inductively- um in order to compare the measured ratio with refer- coupled plasmas. Pressures were included in the range ence conditions and black body rediation. In order to 1.2-12 kPa whereas global specific enthalpies were rebuild the ICOTOM signals, measurements were in- included in the range 5.5-13 MJ/kg. The radiometers cluded in a radiative transfer code. Final comparison were placed in a cooled holder, perpendicularly to the between calculations and experiences are on progress. plasma jet, on the border of it. The cooling system al- References: lowed to controlled the ICOTOM temperatures. In [1] Gülhan A., Thiele T. and Siebe F. (2013), 7th order to analyze, the radiation received by the EWTPSHT 7, Noordwijk. [2] Omaly, P., Hebert, P-J ICOTOM inside the plasma, complementary meas- (2014) IPPW11, Pasadena. [3] Boubert P., Bultel A., urements were carried out in order to characterize the plasma in terms of densities and temperatures. Chéron B. G. and Vervisch P. (2009), J. Tech. Phys., 50(3), 163-179.

RAMAN LASER SPECTROMETER (RLS) INSTRUMENT FOR 2020 EXOMARS MISSION A. G. Morala,1, F. Rullb, S.Mauricec, I. Hutchinsond, C.P. Canoraa, G. Lópezb, R. Canchala, P. Gallegoa, L. Seoanee, J.A.R. Prietoe, A. Santiagoe, P. Santamaríaa, M. Colomboa, T. Belenguera, G. Ramosa, Y.Parotc, R. Ingleyd, S. Woodwardf, W.Shulteg aInstituto Nacional de Técnica Aerospacial (INTA), Ctra. Ajalvir, Km 4, 28850 Torrejón de Ardoz, Spain. bUniversity of Valladolid (UVa) - Centro de Astrobiología (CAB). Av. Francisco valles, 8, Parque Tecnológico de Boecillo, Parcela 203, E-47151 Boecillo, Valladolid, Spain www.cab.inta.es. www.inta.es. cInstitut de Recherche en Astrophysique et Planétologie (IRAP), France. www.irap.omp.eu. dUniversity of Leices- ter, UK. www.le.ac.uk. eIngeniería y Sistemas (INSA), Spain. www.insa.es. fRutherford Appleton Laboratory (RAL), UK. www.stfc.ac.uk. g OHB System ENG, Germany. www.ohb-system.de

Introduction: The Raman Laser Spectrometer (RLS) is one of the Pasteur Payload instruments, with- Results: RLS expected main characteristics are as in the ESA’s Aurora Exploration Programme, Exo- follows: Mars 2020 mission. • Laser excitation wavelength: 532 nm ExoMars Rover would carry a drill and a suite of • Irradiance on sample: 0.6 - 1.2 kW/cm2 instruments dedicated to exobiology and geochemistry • Spectral range: 150-3800cm-1 research and its main Scientic objective is “Searching • Spectral resolution: between 6 cm-1 and 8 for evidence of past and present life on Mars". cm-1 Raman Spectroscopy is used to analyze the vibra- • Spectral accuracy: < 1 cm-1 tional modes of a substance either in the solid, liquid or • Spot size: 50 microns gas state. It relies on the inelastic scattering (Raman Scattering) of monochromatic light produced by atoms TRL8: After a wide campaign for evaluating In- and molecules. The radiation-matter interaction results strument performances by means of simulation tools in the energy of the exciting photons to be shifted up or and development of an instrument prototype, Instru- down. The shift in energy appears as a spectral distri- ment Structural and Thermal Model was successfully bution and therefore provides an unique fingerprint by delivered on February 2015. which the substances can be identified and structurally Since then, the RLS Engineering and Qualification analyzed. Model has been manufactured and is expected to be The RLS is being developed by an European Con- delivered by May 2017, after a full qualification testing sortium composed by Spanish, UK, French and Ger- campaign developed during Q3 & Q4 of 2016. man partners. It will perform Raman spectroscopy on A summary of main Instrument functionalities and crushed powdered samples, obtained from 2 meters performances obtained during the last months, achiev- depth under Mars surface, inside the Rover’s Analyti- ing high levels of scientific performances will be de- cal Laboratory Drawer. scribed. The Raman Laser Spectrometer Instrument: The RLS Instrument is composed by the following units: • SPU (Spectrometer Unit) • iOH: (Internal Optical Head) • ICEU (Instrument Control and Excitation Unit) Other instrument units are EH (Electrical Harness), OH (Optical Harness) and RLS Application SW On- Board.

Figure 1: RLS layout over Rover ALD (Analytical Laboratory Drawer) APPLICATION OF GENERAL CIRCULATION MODELS FOR THE PREDICTION OF ATMOSPHERIC CONDITIONS AT THE LANDING SITE: A planetWRF STUDY O. Temel, von Karman Institute for Fluid Dynamics, Royal Observatory of Belgium, J. van Beeck, von Karman Institute for Fluid Dynamics, O. Karatekin, Royal Observatory of Belgium, ([email protected]).

Introduction: Global circulation models (GCM) Schiaparelli landing site: The lander crashed at are being used not only to perform realistic predictions 2.1 degrees south and 6.2 degrees west. Thanks to the for the global-scale and synoptic scale atmospheric quasi-homogeneous terrain of Meridiani Planum re- transport mechanisms but also provide initial and gion, it is possible to stat that the GCM simulations can boundary conditions for mesoscale models. Even also provide accurate predictions of atmospheric condi- though the spatial resolution that can be reached with tions without performing additional nested mesoscale GCMs is limited in comparison to mesoscale models, or microscale simulations sustained from the inflow GCMs can also be used for the prediction of atmos- conditions of GCM. pheric conditions at the landing sites where the topog- The images taken from NASA orbiter has shown raphy can be considered as flat. In this present study, that the parachute is located at 1 km south of the im- the planetWRF [1], an extension of the Weather Re- pact point of Schiaparelli lander, indicating the fact search and Forecasting model which is being used for that wind direction was from north to south. research and operational purposes for the Earth’s atmosphere [2], has been used to predict the landing site conditions of the ExoMars – Schiaparelli lander which was planned to land on Meridiani Planum region during the dust storm season corresponding to the solar longitude of 244.7. Overview of the GCM model: planetWRF is actually a multiscale atmospheric model including the nesting option and generalized map projection formulation. The model uses the classical terrain-following coordinated on a C-grid and the baseline model, WRF, has been modified by adapting the transport equations on the non-conformal grid configuration. Model can run as a single-column model as well as two dimen- Figure 1: Mars Reconnaissance Orbiter Camera im- sional and three-dimensional GCM. Several different ages taken after the landing of Schriapelli EDM lander. physical processes are parameterized by using the sub- grid schemes: The results of the GCM simulations over the Me- x Multiphysics schemes for the parameterization ridiani Planum will be presented and the sensitivity of of water, CO2 cycles and the interaction of the models prediction to the model parameters will also dust transport. be included in the presentation. x A wideband longwave and shortwave radia-

tion parameterization schemes in addition to a References: [1] Richardson, Mark I. et al. Journal UV heating parameterization scheme [3]. of Geophysical Research: Planets 112.E9 (2007). [2 x Different dust scenarios including the ones Skamarock, William C., et al. National Center For Atmos- based on the TES observations or specific pheric Research Boulder Co Mesoscale and Microscale Me- dust scenarios like MGS dust-scenario. teorology Div, 2005.. [3] González‐Galindo, F., et al. x A thermodynamic equilibrium radiative trans- Journal of Geophysical Research: Planets 110 (2005). fer model for infrared emissions [4]. [4] López‐Valverde, M. A., and M. López‐Puertas. x Subsurface schemes to model the heat and Journal of Geophysical Research: Planets 99.E6 mass transfer within the Martian soil. (1994): 13093-13115. [5] Hong, Song-You, and Hua- x Surface layer and planetary boundary layer Lu Pan. Monthly weather review 124.10 (1996): 2322- schemes for the parameterization of turbulent 2339. [6 Mischna, Michael A., et al. Planetary and Space fluxes within the boundary layer [5]. Science 59.2 (2011): 227-237. x Extended multiphysics schemes to model the transport of passive scalars within the Martian atmosphere [6]. Preliminary Performance Analysis of the Entry, Descent and Landing of the Exomars 2020 Mission. ExoMars EDL team,1,2,3, 1ESA([email protected]), 2, LAV 3 TAS

Introduction: The 2020 mission of the ExoMars programme will deliver a European rover (ESA) and a Russian surface platform with a set of scientific instruments to the surface of Mars. The scientific objectives of the ExoMars programme are the search for traces of past and present life on Mars, the characterization of the water/geochemical environment as a function of depth in the shallow subsurface, the characterization of trace gases in the Martian atmosphere, the investigation of scientific issues in the frame of long-term stationary platform. The Exomars programme seeks to demonstrate the following: – SC entry, descent and landing on the surface of Mars; – Rover surface mobility; – Access to shallow subsurface for sampling; – Collection of samples, set up for analysis and analysis of samples; – Qualification of Russian ground space-tracking stations in cooperation with ESTRACK (ESA).

This paper intends to describe the EDL phase of the EXM 2020 mission and the preliminary results of its performance.

ExoMars Entry, Descent and Landing Demonstrator Schiaparelli: flight overview and mission results. T. Blancquaert1, O. Bayle1, L. Lorenzoni1, A.J. Ball1 and the ESA ExoMars EDM team, 1ExoMars Project, Directorate of of Human Spaceflight and Robotic Exploration, European Space Agency ([email protected], ESA ESTEC, Keplerlaan 1, PO Box 299 NL-2200 AG Noordwijk, The Netherlands).

The ExoMars EDM Schiaparelli was launched with the Trace Gas Orbiter on 14 March 2016, operated for checkout and timeline programming purposes during cruise, and separated as planned from TGO on 16 Oc- tober 2016 during approach to Mars. At the end of the coast phase, Schiaparelli awoke from hibernation and began transmitting as expected. The UHF carrier was monitored by ESA's Mars Express spacecraft in Mars orbit, and from Earth by the GMRT radio telescope in Pune, India. Data from the 'essential telemetry' UHF transmission was recorded by the ExoMars TGO and relayed to Earth. Analysis shows that Schiaparelli entered the Martian atmosphere successfully and on target, deployed its parachute at 12km above the surface, jettisoned the Front Shield at 7.8km, then separated from the back shell and continued transmitting until signal was lost about 40s before the expected touchdown. Although the planned 4 sols of surface operations (relay of the full EDL dataset, DeCa descent images and science measurements from the DREAMS surface payload) did not occur, the essential telemetry received has enabled reconstruction of Schiaparelli's flight profile and main events of EDL, and provided the COMARS+ and AMELIA experiment teams with data for analysis in line with their research objectives. Good progress has been made in investigating the Ex- oMars Schiaparelli anomaly. This presentation will serve as an introduction to the ExoMars session that includes more detailed presentations on particular aspects and the outlook for the ExoMars 2020 mission.

http://exploration.esa.int/ EXOMARS 2016, THE SCHIAPARELLI MISSION. EDL DEMONSTRATION RESULTS FROM REAL TIME TELEMETRY BEFORE UNFORTUNATE IMPACT

S.Portigliotti1, C.Cassi1, M.Montagna1, P.Martella1 , M.Faletra1, S.De Sanctis1, D.Granà1 O.Bayle2, T.Blancquaert2, L..Lorenzoni2 1Thales Alenia Space – Italia, Demain Exploration and Science Italy Str. Antica di Collegno 253, 10146 TORINO – ITALY, 2ESA-ESTEC, Wuropean Space Agency Researcj and Technology Centre, 2200 AG Noordwijk, The Netherlands

Abstract: Martian exploration, represented by the second Exo- ExoMars programme is part of a cooperation be- Mars mission scheduled to be launched in 2020. tween ESA and Roscosmos. It consists of two missions to Mars that aim to searching for signs of past The first part of the paper will focus on several as- and present life on Mars, investigating the Martian pects of the GNC design that were successfully proven atmosphere and surface and demonstrating the technol- in flight: the innovative logic invented to propagate the ogies needed to land on Mars and perform robotic ex- attitude solution across the 3 days period of hiberna- ploration. tion, the inertial navigation of the complete state vec- The first spacecraft composed of an Orbiter Mod- tor, the logics for detecting the Mars atmosphere inter- ule (the Trace Gas Orbiter, TGO) carrying an Entry face entry point and for the parachute deployment. descent and landing Demonstration Module (EDM) Also some other aspects, in particular the RDA pro- have been launched on 2016 March 14th by a Pro- cessing and the RCS capability to track an attitude proton/Breeze-M launcher. The EDM, called Schiaparelli file, later on compromised by the anomaly occurred at in honor of the Italian scientist Giovanni Schiaparelli, parachute inflation, are presented, to observe that the famous for his studies on Mars, was released by the GNC software has performed as designed and ex- TGO on October 16th 2016 and reached the Martian pected, albeit without any possibility to target the aim soil three days after. of the soft landing owing to the anomaly identified in The EDM acronym was used to identify the prima- the gyroscopic measurements that occurred at para- ry objective for Schiaparelli: EDL Demonstrator Mod- chute inflation. ule. Despite Schiaparelli has not been in the condition Exploitation of flight data for the trajectory recon- to achieve the soft landing, the availability of real time struction, the assessment of the on-board navigation telemetry data transmitted during EDL and open-loop solution and the exploitation of the implemented EDL recorded by the TGO during its own Mars Orbit Inser- sensors is described in the second part of the paper, as tion provided a unique capability for Europe to analyze preparatory work for more in-depth EDL science post the module status and performance as well as to under- flight analyses that are currently carried on by scientific stand the origin of the mishap. Combination of real- teams. The exploitation of the Front Shield and Back time telemetry to radiometric tracking from Pune Giant Shield instrumentation, with pressure sensors, thermal Metrewave Telecope, Mars Express and TGO Electra plugs and the advanced DLR COMARS+ and radiome- receiver allowed the inspection of EDL phases through ters present the unique opportunity to verify the ade- correlations with post-flight analyses currently limited quacy of the margins and assumptions for thermal pro- to Level-0 assessment. tection system design as well as correlations to flight This paper describes the main achievements for dynamics rebuilding for the verification of the aero- EDL demonstration standpoint, still achievable in spite thermodynamic databases assumptions. of the mishap and of the availability of telemetry lim- The study of any observed major or minor anomaly ited by the data volume and data rate constraints, as and clear definition of lessons learned from the “EDL well as by the plasma blackout phase. Demonstration” of Schiaparelli is considered as a key The GNC system aspects of the mission are worth to improve the probability of success for the oncoming to be highlighted for the relevant number of aspects 2020 mission completing the ExoMars programme. that where successfully demonstrated. Furthermore the output and the lessons learned of this mission are fundamental in view of the paramount milestone, in the EXOMARS 2016 POST FLIGHT MISSION ANALYSIS OF SCHIAPARELLI COASTING, ENTRY, DESCENT AND LANDING

D. Bonetti1 ([email protected]), G. De Zaiacomo1, G. Blanco Arnao1, I. Pontijas Fuentes1, S. Portigliotti2 ([email protected]), L. Lorenzoni3 ([email protected]), O. Bayle3

1 DEIMOS Space S.L.U., Ronda de Poniente 19, Tres Cantos, 28760, Spain 2 Thales Alenia Space Italia, Italy, 3 European Space Agency (ESA), The Netherlands,

Introduction: The ExoMars programme is pursued communicate with the TGO and with the Mars Recon- as part of a broad cooperation between ESA and naissance Orbiter (MRO), transmitting its real time on- Roscosmos with significant contribution from NASA. board telemetry (OBT). The data collected along the Two missions compose the ExoMars programme with Schiaparelli mission is extremely valuable in prepara- the launches in 2016 and 2020. The main programme tion to the 2020 mission. goals of the first mission are to search for evidence of For the ExoMars2016 Mission, Thales Alenia methane and other trace atmospheric gases that could Space Italia acted as prime contractor, leading the be signatures of active biological or geological pro- Spacecraft Composite development and verification cesses and to test key technologies in preparation for (including the system design and verification of the ESA's contribution to subsequent missions to Mars, EDM and key GNC/EDL technologies). DEIMOS which will deploy a landing Platform and a Rover for Space has been involved in the Exomars Programme searching sign of present and past life as well as map- (2016 and 2020 missions) since 2004 providing more ping surrounding environment. than 10 years of technical activities in the areas of End The ExoMars 2016 mission is led by ESA; it has to End (from launch to landing) Mission Engineering been successfully launched from Baikonur by the Rus- and GNC. In particular, for the 2016 mission DEIMOS sian launcher Proton-M on March, 14th 2016 and Space was responsible in Phase E of the Mission Anal- reached Mars on October, 19th 2016. The mission in- ysis of the Schiaparelli mission, from separation from cludes the Trace Gas Orbiter (TGO) and the Entry, the TGO to landing on Mars, covering the pre-flight Descent, and Landing Demonstrator Module (EDM, trajectory predictions and performance analysis [1] and named Schiaparelli), both supplied by ESA. the post flight analyses presented here. The ExoMars 2020 mission ("ExoMars Rover and For what concers flight predictions, DEIMOS Surface Platform Mission") includes a Carrier Module Space has been part of the core team being responsible and a Mars Rover developed by ESA, and a Descent of the Schiaparelli EDL mission analysis, from separa- Module including a Surface Platform developed by tion from the TGO to landing. After key interplanetary Roscosmos. The project is is currently in Phase C/D maneouvres designed and operated by ESA, EDL per- and it is scheduled to be launched by Proton in 2020. formance have been monitored covering multiple as- The 2016 TGO scientific mission aims at investi- pects: system margins identification through local en- gating atmospheric trace gases: it is currently in the try corridors analyses and 3DoF/6DoF end-to-end aerobraking phase that will place the spacecraft on its Monte Carlo campaigns, verification of nominal ESA target orbit for scientific operations, expected to begin trajectories and separation maneuver optimization for in March 2018 and to run for five years. On October landing site targeting, EDM aerodynamic database 16th 2016, after 7 months of interplanetary flight, inspection and Flying Qualities Analysis, and TGO- Schiaparelli separated from the TGO and 3 days later Schiaparelli geometric visibility analyses [1]. entered into the Martian atmosphere with a pre-defined For what concerns post flight, DEIMOS Space per- flight-path angle. with the objective to land at the tar- formed the activities presented in this paper, in support geted landing site at Meridiani Planum. and strict collaboration with Thales Alenia Space Ita- Following the coasting phase, the hypersonic entry lia. These activities aimed to assess the main flight phase of Schiaparelli was nominal, and the EDM de- performance from the mission analysis point of view, creased its velocity until its disk gap band parachute determine the behaviour of the Schiaparelli capsule was deployed. During the descent phase an anomaly with respect to the predicted one, and identify the tra- occurred, and the parachute cut was commanded earli- jectory flown during the end-to-end EDL mission. The er than expected, compromising the landing phase of objective was to compare and validate the mission de- the mission. During the EDL, Schiaparelli was able to sign methodology, and the associated performance, against flight data. The analysis covered different activities: trajectory reconstruction, entry aerodynamic and flying qualities analyses, EDL performance reconstruction and assessment, descent and landing data analysis and trajectory reconstruction. The results obtained in terms of most likely and reconstructed trajectories are compared with respect to the variability associated to the baseline mission design as predicted in phase E. A large part of the end-to-end Schiaparelli mission, from Launch to Descent, has been successfully validated with the 2016 flight. Overall, the post flight results contributed to the validation of key technologies and design tools, including the DEIMOS Space Plane- tary Entry Toolbox (PETBox) [2] for Mission Engi- neering and the related design methodology for At- mospheric Flight, that are now Flight Qualified for both missions on Earth (through the successful ESA IXV mission, in which DEIMOS Space was responsible of the Mission Analysis and co-leader of the IXV GNC as part of the core team [3]) and Mars.

References:

[1] Bonetti D. et al (2016) “ExoMars 2016 Schia- parelli Mission Analysis”, 67th IAC. [2] Bonetti D. et al (2016) “PETbox: Flight Quali- fied Tools for Atmospheric Flight”, 6th ICATT. [3] Bonetti D. et al (2015) “IXV Mission Analysis and Flight Mechanics: from design to postflight”, AIDAA 2015.

ATMOSPHERIC MARS ENTRY AND LANDING INVESTIGATIONS & ANALYSIS (AMELIA) BY THE EXOMARS 2016 SCHIAPARELLI MODULE

F. Ferri1, O. Karatekin2, A. Aboudan1, B. Van Hove2, G. Colombatti1, C. Bettanini1, S. Debei1, N. Gerbal2, S. Lewis3, F. Forget4, S. Asmar5 1Università degli Studi di Padova, Centro di Ateneo di Studi e Attività Spaziali “Giuseppe Colombo” (CISAS) ([email protected]) 2Royal Observatory of Belgium (ROB), Brussels, Belgium 3School of Physical Sciences, The Open University, Walton Hall, Milton Keynes MK7 6AA, UK. 4Laboratoire de Météorologie Dynamique, UPMC BP 99, 4 place Jussieu, 75005, Paris, France 5Jet Propulsion Laboratory, California Institute of Technology - NASA, Pasadena, CA, USA

On the 19th October 2016, Schiaparelli, the Entry and descent trajectory. These data contribute to explor- Demonstrator Module (EDM) of the ESA ExoMars ing a wide altitude range and a vertical resolution not Program entered into the martian atmosphere. Alt- fully covered by remote sensing observations and im- hough it did not complete a safe landing on Mars, it posing important constraints on validation of Mars’ transmitted data throughout its descent to the surface, atmosphere models. but the signal was lost about 1 minute before the ex- Within the AMELIA team, different approaches, pected touch-down on Mars’surface. algorithms, methods and data set are used for simula- The main objective of the Atmospheric Mars Entry tion and reconstruction of the Schiaparelli trajectory and Landing Investigations and Analysis (AMELIA) and attitude during the EDL phases in order to retrieve experiment was the assessment of the atmospheric and validate the most accurate atmospheric profile. science and landing site by exploiting the Entry De- A strong effort was also put into atmospheric mod- scent and Landing System (EDLS) sensors of Schiapa- elling and data assimilation in order to improve predic- relli beyond their designed role of monitoring and tions and weather forcasts, monitoring weather condi- evaluating the performance of the EDL technology tions and to assess the atmospheric context at entry, so demonstrator. AMELIA aimed at studying some of the as to forecast the environmental conditions that Schia- major properties of the martian atmosphere, such as parelli was going to face, but also in view of scientific density, pressure, temperature and wind, from an alti- analysis and interpretation of the AMELIA results. tude of about 130 km all the way down to the surface, and at characterizing the nature of the landing site by Despite the ultimate failure of Schiaparelli to land means of touch-down measurements. safely, sufficient EDL data was returned in order to During its descent to the surface of Mars, Schiapa- reconstruct the trajectory and attitude of the EDM and relli continuously transmitted telemetry that was re- retrieve atmospheric profiles over the altitude range ceived from the ExoMars TGO (Trace Gas Orbiter) from 121 km to 4 km above the surface. while the signal carrier was recorded by the Giant Me- We will report the results on the atmospheric re- tre-wave Radio Telescope (GMRT) in Pune (India) and construction in terms of the assessment of the atmos- by the ESA MarsExpress orbiter until the loss of sig- pheric science and put the experience and lessons nal. Due to reduced data rate only essential telemetry learned into perspective for the ExoMars 2020 mis- data were transmitted during EDL; the whole complete sion. EDL data set should have been transmitted after landing. Due to the crash landing, no data have been returned from the Mars relay orbiters during the planned subsequent passages. The radio signal and the flight data, although more limited than expected, are essential to investigate the anomaly that caused the crash landing and for the achievement of the AMELIA scientific objectives.

The measurements recorded during entry and descent have been used for the reconstruction of the EDM trajectory and attitude determination and for the retrieval of an atmospheric profile of parameters such as density, temperature and pressure along the entry DIRECT-TO-EARTH RADIO LINK FROM THE EXOMARS SCHIAPARELLI LANDER. S. W. Asmar1, S. Esterhuizen1, Y. Gupta2, K. De3, D. Firre4, C. D. Edwards, Jr.1, Ö. Karatekin5, and F. Ferri6, 1Jet Propulsion Laoratory, California Instiute of Technoclogy, USA, ([email protected]), 2Giant Metrewave Radio Telescope, India, 3California Institiute of Technology, USA, 4European Space Operations Center, Germany, 5Royal Observato- ry of Belgium, 6Università degli Studi di Padova, Centro di Ateneo di Studi e Attività Spaziali “Giuseppe Colom- bo,” Italy.

Introduction: The European Space Agency’s Ex- made available via Internet from GMRT to ESOC dur- oMars Trace Gas Orbiter (TGO) released the Schiapa- ing the landing event. The GMRT analysis provided relli Lander on 16 October 2016, three days after arriv- confirmation of the aliveness of the lander through ing at Mars. During the planned sequence of the entry, peak heating, and subsequent parachute deploy- probe’s Entry, Descent, and Landing (EDL), NASA- ment. At 14:57:08 UTC Earth Receive Time, the signal JPL’s Interplanetary Network Directorate (IND) ar- was unexpectedly lost. This epoch corresponds to a ranged for and carried out the reception and recording Spacecraft Event Time (SCET) of 14:47:21, well in of the UHF signal, intended only for proximity links to advance of the predicted touchdown time of 14:48:33 several Mars orbiters, at the the Giant Metrewave Ra- dio Telescope (GMRT) near Pune, India. This resulted 97 72+#* 7 in the availability of the only near real-time visitbility !%&' & 5)*78 at the project’s mission operations center. With a simple cooperaticve agreement, the leadership and staff of !%&'(&% )*+ , GMRT successfully configured the array to provide real-time detection of the lander carrier signal. During the latter stage of the sequence of EDL events, an !%&''&%$)* + + - anomaly resulted in the loss of the lander, which was 9& first observed and reported via the loss of the UHF 7 8)* >8 + ! signal at GMRT. !%&'.& 5)* 6 -A 7# > The data acquired in the Direct-to-Earth (DTE) effort are useful for reconstructing the EDL events, espe- Fig 1: Received Schiaparelli UHF carrier frequen- cially when combined with the carrier signal received cy at GMRT, up to loss of signal at 14:57:08 UTC by the TGO and Mars Express relay, and the telemetry (Earth Receive Time). from on-board sensors. In addition to understanding the probe’s state and trajectory, the DTE data especial- References: ly have a potential to contribute understanding Mars’s [1] Edwards, C, D, et al., “Relay Communications atmosphere in cooperation with the Atmospheric Mars Support to the ExoMars Schiaparelli Lander,” 2017 Entry and Landing Investigations & Analysis IEEE Aerospace Conference. (AMELIA) team. AMELIA aims include studying the major properties of the Martian atmosphere, such as density, pressure, temperature and wind. GMRT Doppler Profile: The GMRT is an array of thirty 45-m antennas operated by the National Cen- tre for Radio Astrophysics. As reported in [1], a sub- array of 12 of the antennas was utilized during the EDL event, representing an effective collecting area equivalent to a single 156-m diameter antenna. For the separation event three days earlier, a larger sub-array of 16 antennas was utilized. The predicted signal level was too low to allow recovery of the Schiaparelli data stream; instead, the objective was by design detecting the lander’s residual carrier signal, which was predicted to be just a few dB above the array’s noise floor. Fig. 1 shows the observed Schiaparelli carrier frequency, after removal of an a priori model based on the predicted lander trajectory. This information was PRELIMINARY ANALYSIS OF ENTRY AND DESCENT RADIO COMMUNICATIONS OF EXOMARS 2016 SCHIAPARELLI. Ö. Karatekin1, N. Gerbal1, B. Van Hove1, F. Ferri2, S. Asmar3 , D. Firre4 , M. Denis4 1Royal Observatory of Belgium, Ringlaan 3, Brussels/Uccle 1180, Belgium, ([email protected]), 2Università degli Studi di Padova, Centro di Ateneo di Studi e Attività Spaziali “Giuseppe Colombo” (CISAS), Italy, 3Jet Propulsion Laboratory, California Institute of Technology, California, USA. 4 European Space Agency (ESA).

Schiaparelli, the Entry Demonstrator Module [2] S. Asmar et. al., Direct-to-Earth radio link from (EDM) of ESA’s ExoMars 2016 mission entered the the Exomars Sciparelli Lander, 14th International Mars atmosphere at 14:42 GMT on 19 October 2016. Planetary Probe Workshop, 10-11 June 2017, The All ensuing communications during entry and descent Hague, The Netherlands. were transmitted by UHF (ultra-high frequency) radio onboard EDM using patch antennas. The back shell antenna was used for the transmission of real-time essential flight data during entry, in case of failure to land [1]. The UHF radio signal used for proximity relay communications was also recorded by the Giant Me- trewave Radio Telescope (GMRT) located in Pune, India. The GMRT was able to track Schiaparelli in real-time prior to Mars atmospheric interface, as well as during its atmopsheric entry and descent. The radio contact was lost with EDM shortly before the expected touchdown [2]. The Schiaparelli radio signal was recorded by ESA spacecrafts, Mars Express (MEX) and Trace Gas Or- biter (TGO), in close proximity. The MEX Melacom communication system monitored and recorded carrier signals from Schiaparelli in open-loop mode. In addition, TGO was able to relay successfully the essential flight data transmitted during entry and descent.The MEX and TGO doppler information and essential data were extracted from the relayed data by ESOC, following the receival of data on ground and transfer to the ESOC mission control centre in Darmstadt. The communications during EDM entry were interrupted by a plasma blackout, which lasted about 1 minute. The radio communications are proved to be critical to provided essential information about the EDM state and the events before the lost of the signal. In this study we present a preliminary analysis of radio communications of ExoMars 2016 Schiaparelli during its entry and descent through the Martian atmosphere. Doppler shifts and power levels received by radio receivers on Earth and by Mars relay orbiters will be analyzed to provide information on the EDM state and trajectory, as well as on the Mars atmosphere.

References: [1] T. Blancquaert et al., ExoMars Entry, Descent and Landing Demonstrator Schiaparelli: flight overview and mission results. 14th International Planetary Probe Workshop, 10-11 June 2017, The Hague, The Netherlands.

AMELIA RECONSTRUCTION OF EXOMARS 2016 SCHIAPARELLI MODULE TRAJECTORY AND ATMOSPHERIC PROFILES BY MEANS OF INERTIAL AND RADAR ALTIMETER DATA. A. Abou- dan1, G. Colombatti1, C.Bettanini1, B. Van Hove2, O. Karatekin2, F.Ferri1, S. Debei1, 1CISAS G. Colombo, Universi- ty of Padova, via Venezia 15, 35131, Padova (PD), Italy, contact email: [email protected] 2Royal Observa- tory of Belgium (ROB), Brussels, Belgium.

Introduction: On 19th October 2016 Schiaparelli, (J2), the aerodynamic acceleration has been derived the Entry Demonstrator Module (EDM) of the Exo- removing the gravitational term from the inertial accel- Mars 2016 mission encountered the martian atmos- eration and finally transformed in the EDM reference phere and begun its entry and descent toward the sur- frame. face. The module was equipped with several sensors to The initial conditions for the integration have been implement guidance tasks, to characterize the aero- derived considering both GNC TM data and the post- thermo-dynamical performance of the module during separation assessment of the EDM trajectory perfomed entry and to support scientific investigations at the by Thales Alenia Space Italy (TAS-I). surface. For each data gap, the last valid TM sample before Schiaparelli transmitted a sub-set of the data ac- the gap and the first TM sample after the gap have quired on-board in real-time during the mission. Re- been used to compute the variation in position (Δp), ceived TeleMetry (TM) allowed the reconstruction of velocity (Δv) and attitude (Δq, where q is the attitude the mission events although the module failed the last quaternion). Then, Δp, Δv and Δq have been used in part of the descent and crashed on the martian surface. the trajectory integration to compute the EDM state The AMELIA experiment is aimed at performing after each gap. the analysis of Schiaparelli’s mission data for scientific First order covariance analysis has been perfomed purposes. In particular this work presents the trajectory to assess the sensitivity of the trajectory to both the reconstructed using the on-board data (acceleration, initial state and the on-board measurement errors. This angular rates and radar data) and the derivation of the corresponds to the Extended Kalman Filter propaga- atmospheric density, pressure and temperature profiles tion step and made possible the assimilation of any along the traversed path. other available data by implementing the proper update Available data: The Schiaparelli Guidance, Navi- step. In particular the RDA measurements have been gation and Control (GNC) subsystem was designed to used to fix the EDM altitude. Then the final trajectory compute the estimate of module position, velocity and has been derived by back-propagation (instead of sta- attitude at 100 Hz using the on-board Inertial Meas- tistical smoothing). urement Unit (IMU) data. The inertial velocity and The atmospheric density has been derived along the position in the real-time telemetry were subsampled at EDM trajectory by inversion of the drag equation, then 1 Hz while the inertial acceleration, the attitude qua- assuming hydrostatic equilibrium pressure and temper- ternions and the measured angular rates were at 10 Hz. ature have been computed by integration. Moreover, during the entry there are some gaps in Results: EDM reconstructed trajectory starts about the received data, in particular a 57.2 s gap due to 15 s before the interface point with the atmosphere at transmission loss because of plasma blackout. The 120.1 km altitude and ends about 2.8 km above the handling of these data gaps in the trajectory recon- surface, 5 s before the activation of the EDM retro- struction is described below. rockets; after this time inertial data is not usefull. The EDM was also equipped with a RaDar Altime- The latitude, longitude and elevation of Schiaparel- ter (RDA) to measure the distance from the surface li at the end of the descent are in good agreement with during the last part of the descent. In particular two reference EDM trajectory provided by TAS-I. Consid- RDA measurements have been used to improve the ering the uncertainty bounds (about 4 km 1σ) the tra- EDM trajectory estimates. jectory is consistent also with the impact position de- Trajectory and atmospheric profiles reconstruc- tected on HiRISE images of the landing site. tion: AMELIA team reconstructed the Schiaparelli Derived density is higher than the one predicted by trajectory by numerical integration of inertial accelera- models for dust storm season. As a consequence de- tion and angular rate data. rived temperature profile is consistent with atmospher- The first step in the reconstruction of the trajectory ic cold and standard climatology scenarios of the Mars was to compute the EDM acceleration in the body- Climate Database (MCD 5.2). fixed frame. This step involves all the GNC data: the position of the module was interpolated at 10 Hz and then used ot estimate the gravitational acceleration ACHIEVEMENTS OF THE COMARS+ INSTRUMENTATION PACKAGE DURING THE ENTRY FLIGHT PHASE OF THE EXOMARS SCHIAPARELLI LANDER

A. Gülhan, T. Thiele, F. Siebe, T. Schleutker, R. Kronen Supersonic and Hypersonic Technology Department of the Institute of Aerodynamics and Flow Technology, German Aerospace Research Center (DLR), Linder Hoehe, D-51147 Cologne, Germany

J. Annaloro, P. Omaly, P.J. Hebert Propulsion, Pyrotechnics and Aerothermodynamics Section, CNES Toulouse, 18 avenue Édouard Belin, 31401 Toulouse Cedex 9, France

The European Space Agency (ESA) performed the ExoMars mission in 2016. One of the main objectives of this mission was the demonstration of a successful Entry, Descent and Landing (EDL) on Mars. Scientific data gathered during the entry and descent phase is very valuable and could for example be used for an optimization of the heat shield. Due to the fact that the prediction of the aerothermal and radiative loads on the back cover using existing experimental and numerical tools still has big uncertainties, the design of the back cover heat shield was carried out with relatively high Figure 1: COMARS+ sensors integrated on back cover TPS. safety margins. In order to measure the aerothermal loads on the Several COMARS+ functional checks during the back cover of Schiaparelli, the EDM back cover in- cruise phase to Mars showed no anomalies. Due to the strumentation package COMARS+ consisting of three failed landing it was not possible to retrieve the com- aerothermal sensors, one broadband radiometer sensor plete data package of COMARS+ with a sampling rate and an electronic box was developed. The aerothermal of 10 Hz, because the main data should have been sensors combined four discrete sensors measuring stat- transmitted after touchdown. Using the data link be- ic pressure, total heat flux, temperature and radiative tween the Schiaparelli lander and the TGO some flight heat flux at two specific spectral bands. For the pres- data were transmitted with a sampling frequency of sure measurement a very small Pirani-type pressure 0.1Hz. Therefore COMARS+ sensor data at ten trajec- sensor was used and the total heat flux was measured tory points were sent, with one point before and nine by a commercial heat flux microsensor which also in- points shortly after the black-out phase. The measured cluded a surface temperature signal. The radiative heat pressure, total heat flux rate, radiative heat flux rate flux at two different spectral bands ([2.6; 3.36 ߤm] and and temperature provided very consistent data. Since [4.17; 5 ߤm]) was measured by two ICOTOM sensors the sensor housing temperatures during Mars entry contributed by CNES [2]. The infrared radiation in a were lower than the predicted values, an additional broadband spectral range was measured by the separate calibration check was carried out in the post-flight broadband radiometer sensor. The sensor signals were analysis using the flight spare models. recorded by the onboard data acquisition system using three analogue channels. Hence the overall 23 sensor References: and 8 housekeeping signals of the payload were ampli- [1] Gülhan, A., Thiele, T., Siebe, F., Koch, U., Kronen, fied to the needed input voltage range and multiplexed R., “COMARS+: A combined sensor package on the th to the three analogue acquisition channels.. The ambi- backcover of ExoMars EDM demonstrator”, 8 Euro- tious low mass and low power design ended at a total pean Workshop on Thermal Protection Systems and mass of 1.73 kg and a power consumption of 4.5 Watt Hot Structures, Noordwijk, The Netherlands, April for the complete COMARS+ payload. After passing all 2016. qualification and acceptance tests COMARS+ was in- [2] Omaly, P., Hebert, P-J, “ICOTOM: Integrated Nar- th tegrated into the back cover (Figure 1). row band infrared radiometer”, 11 International Planetary Probe Workshop, Pasadena, USA, June 2014. &+89@/ F+<+>/538 $#-26/?>5/< H62+8 /<<3 ,9?.+8 9697,+>>3 "9C+6 ,=/<@+>9@29,=/<@+>97/8> 90 >2/ 8=>3>?>/ 90 /<9.C8+73-= +8. 69A $/-289691C /<7+8 /<9 =:+-/"/=/+<-2/8>/<"96918//<7+8C %83@/<=3>G./163#>?.3.3!+.9@+/8><9.3>/8/9.3#>?.3/ >>3@3>G#:+D3+63K3?=/::/9697,9L##>+6C ! 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G. Bailet1, A. Bourgoing2, T. Magin3 And C. O. Laux4, 1PhD Candidate, CentraleSupélec, Laboratoire EM2C, CNRS UPR288, Université Paris Saclay, 92290 Châtenay-Malabry, France . [email protected], 2AIRBUS Safran Launchers, Route de Verneuil, 78133 Les Mureaux CEDEX, France, 3Associate Professor, Von Karman Institute for fluid dynamics, Aeronautics & Aerospace Dept, Chaussée De Waterloo 72, 1640 Rhode-St-Genèse, Belgium, 4Professor, CentraleSupéle, Laboratoire EM2C, CNRS UPR288, Université Paris Saclay, 92290 Châtenay-Malabry, France.

Introduction: Reentry radiation is one of the key ucts (such as CN, CO…) or even phenols, alcalines, phenomena to take in account while designing a high metallic ions. The poster will present the expected speed reentry vehicle (>10km/s). In the case of an INES results for the QARMAN mission (example in ablative Thermal Protective System (TPS), the physi- Fig. 1). co-chemical processes involved are complex and rarely Pushing the INES envelop: The INES payload fully characterized. effectively measure radiation without being impacted The QARMAN mission [1] (Flight 2017 Q3) gave by the pollution of an ablative TPS, even after several us the opportunity to develop a dedicated payload minutes of plasma exposure. Efforts at EM2C and von named INES (Imbedded Nano-platform-size Emission Karman Institute ground test facilities (ICP torches) Spectrometer) to study the radiation of the reentry allowed for successful testing of INES on a wide range plasma in the presence of an ablative heatshield [2]. of missions, from low heat flux (300 kW/m2) with The poster presents the different aspects of the pay- PTFE TPS material to high heat flux (> 8 MW/m2) and load design and the efforts developed to push the enve- for a Mars entry (2.6 MW/m2) . lope of the payload use. Conclusion: The INES payload will be rated INES for QARMAN: The QARMAN vehicle (34 TRL9 if the flight of QARMAN reentry vehicle by the cm long for 5 kg) will reenter the Earth’s atmosphere end of 2017 is successful. In the meantime, the qualifi- from an initial altitude of 400 km by the end of 2017. cation tests are planned to expend the use of INES and Located inside the payload bay, INES will take meas- propose it to the EDL community. urements of the radiation along the stagnation line. References: Spectra predictions: The idea behind the method [1] I. Sakraker, et al., Qarman: an Atmospheric En- is to determine the absorption and emission of a try Experiment on CubeSat Platform, 8th ESASV, Por- hot/ablative boundary layer when a TPS sample is tugal, March 2015. subjected to a high enthalpy flow (ICP Torch). [2] NASA-CR-170881, SRB TPS Materials Test By characterizing the radiation environment of a Results in an Archeated Nitrogen Environment, Lock- ground facility and comparing it with simulations (no heed Missiles and Space Co, 1979. ablation) and experiments (with a INES payload inside the TPS sample), it is possible to predict the radiation of missions with any TPS material with ablation prod-

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Figure 1: Calculated spectrum (61 km of altitude, 5,3 km/s, no ablation and cold wall, top) and overlayed in red, the prediction of QARMAN radiation taking in account the absorption/emission of the hot ablative boundary layer (bottom) MARS 2020 ENTRY, DESCENT, AND LANDING INSTRUMENTATION 2 (MEDLI2) SENSOR SUITE Helen H. Hwang,1 Henry S. Wright2, Christopher A. Kuhl2, Mark Schoenenberger2, Todd R. White3, Christopher D. Karlgaard4, Milad Mazhari3, Tomo Oishi5, Steven Pennington6, Dominic Trombetta2, and Jose Santos7, 1NASA Ames Research Center ([email protected]), 2NASA Langley Research Center, 3NASA Ames Research Center, 4Analytical Mechanics Associates, Inc., 5Jacobs Technology, Inc., 6Science Systems and Applications, Inc., 7Sierra Lobo, Inc.

Introduction: The Mars 2020 Entry, Descent, and [2] F. Cheatwood et al. (2014), NASA/TM 2014- Landing Instrumentation 2 (MEDLI2) sensor suite 218533 seeks to address the aerodynamic, aerothermodynamic, and thermal protection system (TPS) performance issues during atmospheric entry, descent, and landing of the Mars 2020 mission. [1] Based on the highly successful instrumentation suite that flew on Mars Science Laboratory (MEDLI)[2], the new sensor suite expands on the types of measurements and also seeks to answer questions not fully addressed by the previous mission. Sensor Package: MEDLI2 consists of 7 pressure transducers, 17 thermal plugs, 2 heat flux sensors, and one radiometer. The sensors are distributed across both the heatshield and backshell, unlike MEDLI (the first sensor suite), which was located solely on the heat- Figure 1. Supersonic pressure transducer developed shield. The sensors will measure supersonic pressure for MEDLI2. on the forebody, a pressure measurement on the aftbody, near-surface and in-depth temperatures in the heatshield and backshell TPS materials, direct total heat flux on the aftgbody, and direct radiative heating on the aftbody. Instrument Development: The supersonic pressure transducers, the direct heat flux sensors, and the radiometer all were tested during the development phase. The status of these sensors, including the piezo- resistive pressure sensors, will be presented. The current plans for qualification and calibration for all of the sensors will also be discussed. Post-Flight Data Analysis: Similar to MEDLI, Figure 2. Medtherm radiometer selected for MEDLI2. the estimated flight trajectory will be reconstructed from the data. The aerodynamic parameters that will be reconstructed will be the axial force coefficient, freestream Mach number, base pressure, atmospheric density, and winds. The aerothermal quantities that will be determined are the heatshield and backshell aeroheating, turbulence transition across the heatshield, and TPS in-depth performance of PICA. By directly measuring the radiative and total heat fluxes on the backshell, the convective portion of the heat flux will be estimated. The status of the current tools to perform the post- flight data analysis will be presented, along with plans for model improvements. References: [1] H. Hwang et al. (2016), 46th Thermophysics AIAA Conference, AIAA 2016-3536. Entry and Descent Gamma-Ray Measurements on a Venus Probe: A New Window on the Atmosphere Pro- file. David J. Lawrence1, Ralph D. Lorenz1, and Patrick N. Peplowski1 1Space Exploration Sector, Johns Hopkins Applied Physics Laboratory, Laurel, MD, 20723, USA. [email protected].

Introduction: During the hypersonic entry phase secondary particle production increases until it reaches of a planetary probe, the number of viable measure- the RP maximum at a pressure of 50-100 mbar (~150 g ments that can safely sample the atmosphere is limited. cm-2, which is ~15 – 20 km on Earth, or a density- Typically only accelerometers are operated, providing equivalent altitude of ~60 – 70 km on Venus). information that yields a density profile. An instrument Proof of Concept: As part of a GRS maturation sensitive to energetic particles can also be operated in project, two GRS sensors were flown on a stratospher- the safe interior of an entry vehicle, and provides com- ic balloon for nearly 22 days over Antarctica in De- plimentary atmosphere information by characterizing cember of 2016. The sensors were a hosted payload on the local charged particle environment. the STO-2 mission [3]. The two sensors were a 2.54 The energetic (>1 MeV) charged particle environ- cm diameter by 2.54-cm long CLYC sensor [4], which ment within a planetary atmosphere is the result of is sensitive to both gamma-rays and neutrons, and a 5 interactions between the space radiation environment cm diameter by 5 cm long CeBr3 sensor [5], which is and the atmosphere itself. Galactic cosmic rays primarily sensitive to gamma rays. CeBr3 has nearly (GCRs), primarily protons, have a local flux that is the same excellent energy resolution as LaBr3, but inversely related to the column density above that without the locally induced background, and is being point, and therefore GCR protons exhibit a column- considered for a Venus surface application. Only re- density dependent profile. As the GCR protons interact sults from the CeBr3 measurements are shown here. with the atmosphere and lose energy, secondary parti- The flight-integrated gamma-ray spectrum is cles (including protons, neutrons, muons, electrons, shown in Fig. 1. The ubiquitous 0.511 MeV (produced pions) are produced via intranuclear cascades. The by positron annihilation in the atmosphere and GRS- secondary flux increases with increasing column densi- adjacent materials) is clearly seen along with numerous ty, until it reaches a local maximum known as the Re- other lines produced by interactions between cosmic gener-Pfotzer (RP) maximum [1]. The altitude loca- ray primary and secondary particles and nearby mate- tion of the RP maximum is directly sensitive to the rials (e.g., aluminum housing and intentionally placed column density of the atmosphere as well as the energy graphite to induce carbon gamma-ray lines), and the spectra of the incoming GCR. Charged particle meas- Earth’s atmosphere. Count rates versus altitude for urements can be iteratvely compared to outputs of radiation transport models to infer the column density profile meas- Oxygen from ured during atmospheric entry. 0.511 MeV atmosphere Venus Application: As annihilation performed in the Venera/Vega missions, a passive Gamma Ray Spectrometer (GRS) can pro- Carbon from vide a measure of radioactive graphite elements in the Venus crust (Uranium, Thorium, Potassium). Neutron-activated measurements could explore other elements [2]. In either case, the Housing GRS can be operated passively materials during entry and descent. Ener- gy deposition in the GRS sensor is caused by both the primary GCRs as well as the secondary particles generated by GCRs. During entry and descent into a Fig. 1. Measured gamma-ray spectrum from the CeBr3 sensor for nearly 22 days at float thick atmosphere, the primary altitude (~40 km). The ubiquitous 0.511 MeV is clearly seen, along with other gamma-ray lines produced by neutron interactions in nearby materials and the Earth’s atmosphere. GCR signal decreases, but the both the 0.511 MeV and high-energy deposition (>6 pends only on the profile directly above the probe loca- MeV) by gamma rays and charged particles are shown tion. In effect, the measurement acts as a static pres- in Figures 2 and 3, respectively. The 0.511 MeV count sure sensor during the entry, and could improve the rate clearly shows the RP maximum at an altitude of accuracy of trajectory reconstruction. 17 km with a high-statistics count rate of >100 counts Application to Other Bodies: In principle, such a per second (cps). The high-energy count rate is at- measurement could yield useful insights on any plane- tributed to interactions between the CeBr3 and primary tary entry/descent to ~1 bar altitude (Venus, giant GCRs, and shows a local maximum of almost 150 cps planets, Titan), and gamma-ray instrumentation is be- at an altitude of 22 km. ing proposed by these authors on proposed New Fron- Scientific Applications: The cosmic-ray flux, like tiers missions to Venus and Titan, To our knowledge, the ultraviolet environment, is in principle of interest in such measurements have not been performed to date for astrobiology [6] in that it influences DNA damage on any planet other than Earth. The Venera 9 and 10 rates. While biota on Venus indeed seems possible, GRS instruments were only turned on at a pressure of the near-cloud-top environment near the RP maximum 7 bars (to measure the spacecraft background [9]), well is the most hospitable region in the Venusian atmos- below the cosmic ray interaction altitude. The Galileo phere. Furthermore, the atmospheric interaction which probe had an Energetic Particle Instrument to map the produces the secondary-particle flux that dominates the Jovian radiation belts, but this was switched off before signal is essentially the same which produces ion- atmospheric entry. In fact gamma-ray measurements electron pairs, predicted to have a maximum rate in the have been made in Venus orbit [10], although for the Venus atmosphere at an altitude of ~60km [7]. The purpose of triangulating the direction of astrophysical Huygens probe measured this maximum, coincidental- gamma-ray bursts, not for studying Venus itself. A ly also at an altitude of ~60km [8]. Attachment of lesson here is that over-parsimonious allocation of data electrons to aerosol particles is a pivotal factor in con- or power on entry probes may preclude unanticipated trolling cloud droplet aggregation physics, and thus the scientific insights. radiative and chemical environment of Venus' cloudy References: [1] Carlson, P. and A. A. Watson, atmosphere. Hist. Geo. Space. Sci., 5(2), 175, 2014; [2] Parsons, A. There is a small variation [7] of the ionization pro- M. et al., IEEE Nuc. Sci. Sympos., Conference Record file with solar activity (which in turn modifies the N46-2; [3] Walker, C. et al., in Ground-based and Air. cosmic ray flux) but accounting for this, the signal is Tele. III, SPIE, 7733, 2010; [4] Glodo, J. et al., J. essentially a function of the column mass above the Crystal Growth, 379, 73, 2013; [5] Quarati, F.G.A. et sensor. It is therefore a close proxy for atmospheric al., Nuc. Inst. Meth. A, 729, 596, 2013; [6] Dartnell, pressure. The GRS measurements, and more specifi- L.R . et al., Icarus, 257, 396–405 2015; [7] Nordheim, cally their variation with altitude, may serve as an im- T. et al., Icarus, 245, 80-86, 2015; [8] Fulchignoni, M. portant constraint on the entry/descent trajectory re- et al., Nature, 438, 785-791, 2005 [9] Surkov, Y. construction. For example, the accelerometer record (1977) 8th Lunar and Planetary Science. 2665-2689. indicates the density along the entry path, which de- [10] Lorenz, R. and Lawrence, D., Planetary and pends on the flight path angle as well as the atmos- Space Science 109/110, 129-134, 2015 phere profile, whereas the column mass (pressure) de-

RP maximum

Fig. 2. Net count rate of the 0.511 MeV gamma- Fig. 3. Count rate of energy deposition greater than ray line as a function of altitude. The RP maximum 6 MeV in the CeBr sensor versus altitude. is clearly seen at an altitude of ~17 km. 3 VISTA: a μ-THERMOGRAVIMETER to detect volatile compounds in space and in different planetary environments. E. Palomba1, F. Dirri1, A. Longobardo1, D. Biondi1, A.Boccaccini1, A. Galiano1, E. Zampetti2, B. Sag- gin3, D. Scaccabarozzi3. 1IAPS-INAF, Via Fosso del Cavaliere 100, 00133 Rome, Italy ([email protected], fabri- [email protected], [email protected]), 2IIA-CNR, via Salaria km 29,300 Monterotondo, Rome ([email protected]); 3Politecnico di Milano, Polo Territoriale di Lecco, Lecco, Italy ([email protected], [email protected])

Introduction: VISTA (Volatile In Situ Thermogra- - determination of composition of non-ice materials on vimetry Analyser) is a μ-Thermogravimeter device, icy satellite surfaces; developed by a consortium of Italian institutes. Ther- - characterization of organic species by measuring its mogravimetric analysis (TGA) is a technique used to enthalpy of sublimation, in order to monitor the evolu- monitor processes involving volatile compounds, such tion of atmospheric aerosol (Earth's environment). as deposition/sublimation and absorption/ desorption [1,2]. In planetary missions, μ-TGA would measure VISTA is composed by two different subsystems: water and organics desorption, whose presence is connected to habitability of the planet/satellite, as well as a. Sensor Head 1 (SH1), able to work at low temperau- to monitor outgassing contamination [3]. res (down to -200°C) and to detect the contaminant The instrument core is composed by a Piezoelectric molecules coming from outgassing processes in space Crystal Microbalance (PCM), equipped with built-in and to perform thermal cycles (up to 100°C); heater and built-in temperature sensor on crystal sur- b. Sensor Head 2 (SH2), able to performs TGA meas- face, and the related Proximity Electronics (PE). The urements (large temperature range, i.e. >200°C) with a crystal oscillation frequency linearly depends on the low power budget. mass deposited on its sensible area, according to the Sauerbrey equation [4]. PCM temperature can be in- The VISTA sensor heads are shown in Fig.1. SH1 is creased by means of integrated heater in order to allow currently at TRL 6 (Fig.1, Left), whereas a laboratory sublimation/desorption of the most volatile component breadboard of SH2 has been developed and is under of the analyzed sample. Mass and composition of the testing (TRL 4-5) (Fig.1, Right). volatile can be inferred by the frequency change and by desorption temperature, respectively. In order to characterize the volatile species, during the deposition and TGA thermal processes it is possible to infer some physical-chemical parameters, e.g. enthalpy of sublimation/evaporation ΔHsub,evap, entropy of sublimation/evaporation ΔSsub,evap and vapor pressures, Pvap (at a given temperature and pressure), typical for each compounds [5].

Planetary/space applications: VISTA instrument has been studied for space applications and proposed for Fig.1. VISTA instrument. Left: SH1 breadboard planetary in-situ missions [6], i.e. it has been selected (Quartz Crystal Microbalance). Right: SH2 breadboard in the payload of the ESA MarcoPoloR mission study (GaPO Crystal Microbalance). [7] (addressed to a primitive Near Earth Asteroid), and 4 has been studied, e.g. on Europa and Ganimede [8] and Technical characteristics: the main innovation of on Mars [9]. The main tasks achievable by VISTA and VISTA concerns the special design, i.e. the built-in depending to the planetary environment are: heater and a built-in temperature sensor. VISTA sensor heads have its own PE (including the frequency coun- - measurement of abundance of volatiles (e.g. water, ter and the temperature control system) while the Main organics) in planetary/asteroidal regolith and come- Electronic (ME) can be shared with other devices of tary-like activity (MarcoPolo-R mission); the scientific package, reducing the total instrument - measurement of dust and ice settling rate, water con- weight. VISTA has a very small mass, volume and tent in dust and humidity (Mars); power requirements and needs a quite small amount of - discrimination between water ice and clathrate hy- material for analysis, i.e. <1 mg. The technical charac- drates (basing on their different sublimation tempera- teristics of the sensor heads are summarised in Table 1. ture) on icy satellite surfaces;

Unit SH1 SH2 Sensor type Quartz Crys- GaPO4 Crys- tal Microbal- tal Microbal- ance ance Resonant Frequency(MHz) 10 5.8 Mass [g] 90 60 Volume [mm] 50x50x38 35x35x25 Power [W] 1 W (peak); 0.62 (peak); 0.12 (mean) 0.37 (mean) Data rate 30 bit/ meas- 30 bit/ measurement urement Operating Temperatures [K] < 180 < 550 TRL 6 4/5 Tab.1 VISTA technical characteristics.

Performance Tests: the isntrument has been validated with contamination, saturation tests and TGA cycles (SH1) and deposition tests to characterize organic compounds in space (SH2) [5].

References: [1] Zinzi, A. et al., (2011), Sensors and Actuators A, 172, 504-510; [2] Dirri, F. et al., (2016), AMT, Atmos. Meas. Tech., 9, 655-668; [3] Wood B.E. et al., (1997), AIAA 97-0841; [4] Sauerbrey, G., (1959), Z. Phys., 155, 206-222; [5] Dirri et al. (2017), 14th IPPW Abstract; [6] Palomba E. et al., (2016), OLEB Journ., 46, 2, 273-281; [7] Barucci, M. A. et al., (2011), Exp. Astronomy, DOI 10.1007/s10686-011- 9231-8; [8] Jones, G. et al. (2017), IPPW abstract; [9] Palomba, E. et al., (2011), EPSC-DPS Abstract, 87. MICRO-PROCESSOR QUALIFICATION FOR PLANETARY EXPLORATION. A. Martín-Ortega1, J. Man- zano1, N. Andrés1, I. Traseira1, J.R Mingo1, P. Manzano1, S. Sampedro1, M.J. Rívas1, M. Alvarez1, S. Martin1, I. Arruego1. 1INTA. National Institute of Aerospace Technique, Spain. [email protected]

Introduction: Planetary exploration strives for Single or Double precission. more capable processing technologies. In-situ pro- Desirable cessing capabilities enable the possibility of executing payloads for longer periods, reducing the size of data #"-# ( > 128 Kbytes needed to be downlinked to Earth. Moreover, further # ( > 8 Kbytes processing capabilities facilitates the decision making, reducing the iteration of payloads operating far away 2x UART from ground stations. This definitely increases the #!!)" - 2x SPI productivity of the science data retrieved from those in- ( #"&#(## ' 1x CAN struments. 1x Serial High Speed > 20Mbits/s In addition, mass and power consumption require- 8/10 bits Converter ments are even more restrictive than for orbital space ( > 16 Input/Outputs missions. This is true for the overall planetary mission, (& RTC and Timers but even more for payloads, which are not considered 5.0V or 3.3V critical for the survival of the whole mission. #"')!$( #" Reference: 10mW/MHz Taking all these requirements into account, it is No BGA or CC (CFP preferred) clearly required to obtain better micro-controllers, capa- Low susceptible to SEL (immune ble of operating in such environments, without being preferred) at Mars environment. careless of the intrinsically reliability requirements of TID > 10 Krad space missions. -130ºC to +70ºC (operational) Previous Studies: INTA (National Institute for Aerospace Technology) in Spain has been involved dur- Research Conclusions: After several months of ing the last two decades in developing miniaturized in- study and revision of the published literature, very few struments for space. More recently started developing candidates barely fulfill most of the imposed require- miniaturized instruments for Mars exploration. Mars ments. Space qualified manufacturers appear to fulfill missions add another complication to space develop- quite well the so called low and high capabilities micro- ments. This is the low temperatures (down to -130ºC) controllers. This is, there are old micro-controllers (8 bit levels, along with a high variability during each Sol in architectures) with very poor processing capabilities, the Martian surface (gradients in the order of ~170ºC). and new micro-controllers (or processors) with very In such conditions, INTA carried out a deep analysis high processing capabilities. However, these ones also of the available micro-controllers having in mind two have very high demanding power consumption, and are different sources. First source was to make a review of definitely not suitable for miniatured payloads. With re- all available Rad-Hard and Rad-Tolerant micro-control- spect to the COTS components, some appear to be valid lers. Once this path was exploited, INTA started revis- candidates, however, further testing was required. ing literature to find COTS that might been tested either Those components were in cryogenic and/or radiation conditions. The state of the - Texas Instruments MSP430 art of mid-range micro-controllers was produced. - Microchip dsPIC The following table shows the flowed down require- - Atmel SAMD Family D21. ments of the micro-controller based on previous experi- The Atmel one wasn’t available at the moment of the ence developing payload and control units: study. The other two had flight experience and some proprietary radiation testing. INTA, performed cryo- Table 1. Micro-controller initial requirements. genic testing of dsPIC30F in the past, and therefore, that component was selected for the next step in the way of qualifying a whole LOT for future exploration missions. Single or dual core Qualification Process: the qualification of the mi- & (()& crocontroller was divided in three mayor groups. Tem- &%)"+ > 16 MHz (80 MHz preferred) Hardware Multiplier and Divisor perature, Radiation and Reliability. A set of 200 micro- controllers has been purchased with the certificate of the Radiation Qualification: to fully qualify dsPIC in same manufacturing lot number. equivalent radiation environment, three different radiation Mars Temperature Qualification: both package and campaigns are defined. The behavior of the microcontroller w.r.t. Single Event Latch-Ups, produced either by HI- or P+ functional operation has been tested. The component is must be characterized. Also an estimation of the Functional validated at manufacturer level from -40ºC to +125ºC. Interrupts provoked by Single Event Upsets must be ob- A package qualification validation campaign of up to tained. Finally, the tolerance of the device to Total Ionizing one thousand cycles has been carried out. This cam- Dose must be also characterized. paign validates the soldering process and package of the - Heavy Ion Test: Heavy ion Single Event Latch- given microcontroller to survive in a Martian environ- up (SEL) test was carried out at the irradiation facility ment for more than a year (Mars year ≈ 687 Earth Days). at the Catholic University of Louvain-la-Neuve (UCL) on September 21-23rd, 2016. The test performed over During the campaign (1122 cycles), the microcontroller three different devices, was based on six different LET was dynamically tested and different temperatures. It from 3.3 to 62.5 MeV·cm2/mg. A deeper explanation also experimented cold power up at -120ºC. Post cy- of the test has been submitted to NSREC’17 by P. cling visual (Figure 1) and X-Ray (Figure 2) inspections Manzano et al. were performed. - Proton Test: Proton testing is prepared to be carried out during the Q2 of 2017 at Paul Scherrer Institut (PSI) in Switzerland. The test setup to be used will be the same as the one used at UCL. It is described in the next section. - Total Ionizing Dose: TID test were performed to other dsPIC device several years ago, by INTA at Radiation Physics Laboratory of Universidad Santiago de Compostela. Conclusions of that test showed that this family may perform normal operation up to 20 Krad. However, more precise testing

Figure 1. Visual and X-Ray inspection after 1122 thermal cy- over the given lot are foreseen also for the Q2 of cles. 2017.

Reliability Screening: a set of reliability test has Qualification Setup: in order to be capable of qual- been prepared following MIL-STD-883 methods for ifying the microcontroller properly, a complex dedi- plastic components screening. The following figure cated hardware and software setup has been. This setup summarize the whole process: aims to facilitate qualification, but it has been especially useful during the radiation qualification campaigns. Hardware Setup: the hardware setup is divided into three different units. A PC, equipped with a serial USB to RS422 interface. An electronic ground support equipment called EGSEPIC. This board is commanded by a microcontroller; Its purpose is to rapidly command and retrieve status of the Device Under Test (DUT) in real time while the DUT is being radiated. Finally, another board equipped only with the dsPIC30F, a CAN Bus driver and an oscillator. Figure 3 shows a descriptive diagram of the whole system.

Figure 2. Screening for dsPIC microcontroller. Figure 3. Hardware setup description diagram.

EGSEPIC board is built with a configurable hard- initialized to 0x00. Then, application software ware latch-up controller. This controller notifies soft- counts the bit flips of each section and sends the ware throughout an interrupt, if the configured threshold results through UART. has been trespassed. This method allows software to dis- - Input Capture / Output Compare Test: these two criminate between Single Event Transients and Single IO peripherals are very useful for motor con- Event Latch-ups. If a configurable amount of time is trolling and other closed loop control applica- maintained with a latch-up condition, then software per- tions. EGSEPIC generates a Pulse Modulated forms a power down cycle to reset the DUT. Apart from Waveform (PWM) that is immediately cloned the latch-up controller, EGSEPIC is provided with al- by DUT using Input Capture peripheral. Once most all possible interfaces of dsPIC DUT. This allows the frequency of the PWM has been resolved, us to test every peripheral inside the microcontroller an identical PWM is generated back from DUT while been tested with high energy particles hitting it. to EGSEPIC. Software Setup: three pieces of software have been - CAN Bus / Timer Test: during this test the CAN developed for committing in-deep analysis of the dsPIC Bus peripheral and one of the dsPIC timers shall vulnerabilities to high energy particles. The first one and be tested. Whenever EGSEPIC generates an possibly the simplest one is the application software specific CAN message, DUT starts a timeout running inside the DUT. After a power cycle is per- timer. When timer expires, DUT echos the re- formed to DUT, it remains at an idle state until it is com- ceived data back to EGSEPIC through the CAN manded to start performing a cyclic testing. Test to be Bus. executed are: - I2C / Register test: this test is thought to test - SPI / Analogue Test: during this test, internal both, I2C peripheral and internal processor reg- ADC and SPI peripherals shall be tested. An an- ister. Whenever EGSEPIC generates a specific alogue signal is generated from EGSEPIC message through I2C, DUT increments an inter- board. Then, DUT converts that analogue signal nal processor register. Then, it takes control of into a digital value that is transmitted back to I2C bus and sends the content of the register. EGSEPIC through SPI. - Discreet Input / Output Test: this test is meant - SRAM / UART Test: during this test SEU sen- to test the digital IO Ports of the dsPIC. Three sitivity of internal dsPIC SRAM and UART pe- inputs and three outputs are used. During the ripheral shall be tested. The test starts whenever test, DUT shall copy the values written by EGSEPIC writes two specific bytes though EGSEPIC from the inputs towards the outputs. UART. Then, DUT performs a memory self- The following diagram shows a tipical execu- test. A part of the SRAM memory has been pre- tion cycle: viously initialized to 0xFF. Other part has been

Figure 4. Software execution cycle.

Preliminary Test Result: the selected devices were mission application, in order to apply any hardening irradiated using six different LETs. All runs were per- measures if required. formed at room temperature and no changes were observed using a Pt1000 sensor. For each LET several runs were performed to reach the minimum numbers of SELs needed to have a good cross section. After the highest first LET and after the last programmed run, the program memory of each DUT was read in order to check it was not affected by the SEEs. No changes were observed in program memory of any of the three devices under test. CREME96 was used to SEL rate estimation and data for the three-tested device are depicted in the following table:

Table 2. SEL rate from Creme96

',* ,+ .+'/ 0 0.03 35.24 1 0.04 2 0.03 31.50

During the whole test, several SEFIs where detected and corrected through the application of a processor reset. No power cycle was required to recover from any conditions. Conclusions: Overall situation of the dsPIC30F microcontroller qualification campaign has been very promising so far. Intensive thermal testing was presented and fantastic results were obtained. Heavy Ion testing was conducted with better results than expected in a commercial device. SEL rate and MTBF for a worst-case Martian scenario, and a 100%- time operational payload is very promising. Proton testing has still to be performed to get a more realistic number. However, no much contribution of high energy protons is foreseen in such scenario. In-deep analysis of SEFI during dynamic test is still to be performed. This analysis shall be the input to each PLANETARY RADIO INTERFEROMETRY AND DOPPLER EXPERIMENT (PRIDE) FOR PLANETARY PROBES Vidhya Pallichadath1, Tatiana Bocanegra2, Giuseppe Cimò2, Dominic Dirkx1, Dmitry Duev3, Leonid Gurvits1,2, Guifré Molera Calvés4, Bert Vermeersen1

1. Faculty of Aerospace Engineering, Delft University of Technology, 2629 HS Delft, The Netherlands 2. Joint Institute for VLBI ERIC, PO Box 2, 7990 AA Dwingeloo, The Netherlands 3. California Institute of Technology, Pasadena, CA, USA 4. Finnish Geospatial Research Institute, Geodeetinrinne 2, FIN-02430 Masala, Finland

PRIDE is a multi-purpose, multi-disciplinary en- craft for other purposes, e.g. communication radio hancement of planetary missions science return, which lines. is able to provide ultra-precise estimates of spacecraft state vectors based on the phase-referenced VLBI (Very Long Baseline Interferometry) tracking and radial Doppler measurements. The Planetary Radio Inter- ferometry and Doppler Experiment (PRIDE) has been developed originally by the Joint Institute for VLBI ERIC (JIVE) for tracking ESA’s Huygens Probe during its descent in the atmosphere of Titan in 2005 and since then adopted for a number of planetary science missions. PRIDE is based on exploiting the technique of VLBI originally developed for other than planetary science applications. The essence of the technique is in interleaving observations of the spacecraft radio signal and signal of background natural celestial sources, usually quasars, enabling estimates of the lateral posi- Figure 1: PRIDE Deliverables tion of the spacecraft in the celestial reference frame and Doppler-shift of the spacecraft’s radio signal. References: These estimates can be applied to a wide range of [1] D. A. Duev. et al. (2012) Spacecraft VLBI and research fields including precise celestial mechanics of Doppler tracking: algorithms and implementation, planetary systems, geophysics and planetary dynamics Astronomy & Astrophysics, 541, A43 and measurements of interplanetary plasma properties. [2] G. Molera Calvés. et al. (2014) Observations PRIDE has been included as a part of the scientific and analysis of phase scintillation of spacecraft signal suite on a number of ESA missions. on the interplanetary plasma, Astronomy & Astro- We present some of the experimental results from achieved by our group over the past decade i.e, on the physics, 564, A4 Huygens Probe in the atmosphere of Titan and the latest results on ESA’s Venus EXpress (VEX) and Mars EXpress (MEX) missions. PRIDE was selected by ESA as one of the eleven experiments of the ESA’s L- class JUpiter ICy moons Explorer mission (JUICE) mission, scheduled for launch in 2022. It will address those of the prime objectives of the JUICE mission which require precise determination of the lateral position of spacecraft on the celestial sphere. This poster will also present some of the current & prospective PRIDE targets.The PRIDE approach described in this work proves its applicability to virtually any deep-space mission almost anywhere in the solar system. Ultimately, PRIDE is an affordable addition to the science output of planetary missions since basically it relies on the instrumentation available onboard space- International Planetary Probe Workshop 2017 Instrumentation and Reconstruction for the ASPIRE Supersonic Parachute Test Campaign Bryan Sonneveldt, Jeremy Hill, Andrew Owens, Clara O’Farrell, Ian Clark Jet Propulsion Laboratory, California Institute of Technology

Abstract – The Advanced Supersonic Parachute obtained from these flights, including the Inflation Research and Experiments (ASPIRE) imagery necessary for a first-ever three- test program is developing an infrastructure for dimensional reconstruction of a parachute the testing of parachute inflation at supersonic inflating at supersonic conditions. test conditions and dynamic pressures analogous to those encountered during Martian EDL. The sounding rocket test infrastructure and associated instrumentation will allow for evaluation of the deployment, inflation, and performance of a full scale, 21.5-m-diameter Disk-Gap-Band parachute of the kind planned for use on the Mars 2020 mission. Although the ASPIRE sounding rocket platform relies on existing launch vehicles to achieve the desired flight conditions, a new payload is being developed to carry the parachute and the required instrumentation. This presentation will describe the development of the state-of-the-art instrumentation suite and the plans for reconstructing the performance of the parachute system. The ASPIRE instrumentation suite is critical to obtaining the data necessary for reconstruction of the test conditions, determination of the parachute loading environment and aerodynamic performance, and obtaining high-resolution data of the parachute deployment and inflation events. This will be achieved using state-of-the-art cameras capturing images at 1000 frames per second and at a 4K resolution, situational video cameras, parachute bridle load pins, and an inertial measurement unit. The development of the payload and instrumentation suite introduced a number of challenges including preventing interference of the parachute deployment with the cameras, designing the parachute bridle interface that allows for accurate load reconstruction, and maintaining data recoverability after splashdown. This presentation will provide an overview of the instrumentation suite and mechanical accommodation design, the test condition reconstruction plans, and the data set to be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

Introduction: The Europa Lander Science Defini- are sublimated by the laser and ii) the tube is moved tion Team Report establishes “searching for evidence down by a pneumatic actuator and once in contact pen- of life in Europa” as higher level science goal and de- etrates in the ice 5 cm. The tube is pressurized and fines as first objective “to detect and characterize any heated to get a conditions where the water is stable at organic indication of past or presente life”. Beside that, this moment the sample is sucked by syringe (con- there are two additional goals: “asses the habitability trolled by a spring) to fill the sample deposit. From this of Europa and to characterize surface and subsur- deposit instruments are filled. face”... Figure 1 show the AWL mechanical configuration. Traces of extant or extinct life could be found po- It has the warm box (WB) to maintain the operative tentially at the surface and near-surface layers, which temperature and protect all the electronics for radia- might be incorporated through reworking (impact gar- tion. The WB by design will guarantee bio-cleanliness dening, mass wasting and internal dynamics) of mate- after integration. The WB will have an opening for the rial brought up from aqueous reservoirs by geological SSS, it wil be closed once at the end of the integration activity such as plumes spewed up into the exosphere: to maintain the biological cleanliness. An opening pro- those are the places to look for signatures. The organic tected with an EPA filter will allows the decompres- molecules are likely to be refractory, particularly if sion during landing. The external structure allows the hanging by the manipulator of the lander. they have been exposed to the surface for long periods of time. Therefore, the search for signs of life needs access to fresh endogenic materials, which should be coming from the aqueous habitable environment in the case of extant life, and must be performed with a specific instrumentation and in the appropriate layers: i) subsurface sampling for better protected samples that could be analysed in different physical states (solid/liquid). Analysis of samples in the aqueous phase will be obtained by melting near-surface ice samples, while chemical disequilibria will be simultaneously characterized during the search for biosignatures. AWL concept: The Astrobiology Wet Laboratory (AWL) is the name given to the small platform composed by: i) Shallow Surface Sampler incharges of making a 10 cm hole on the water ice surface and take liquid sample, ii) the Data Processing Unit; iii) Power Unit; iv) Communication Unit to establish physically the connections with the An external structure allows Figure 1. AWL concept. A radiation protection box where are located de sensor and a sampler. to deploy the AWL with the lander manipulator, v) Multi-Probe ArraySensor (MPAS) for immunoassay References: [1] based on lateal flow concept and vi) Multipareametric http://solarsystem.nasa.gov/docs/Europa_Lander_SDT probe (MPP), with a set chemFET sensor to determine _Report_2016.pdf [2] Ulamec S. et al. Access to glacial different physco-chemical parametes. and subglacial environments in the Solar System by melting For the Shallow Surface Sampler (SSS), we have probe technology. Rev. Environ Sci Biotechnol 6 (2007). [3] evaluated different alternatives for drilling (Ulamec Biele J. et al. In situ analysis of Europa ices by short-range 2007, Biele 2011, Weiss 2011, Sakurai 2016). Taking melting probes. Advance in Space Research 48 (2011). [4] into account the limitations on resources and trying to Weiss P et al. Thermal drill sampling system onboard high- reduce as much as possible the use of any mechanism velocity impactors for exploring the subsurface of Europa. the most promising option is a drilling system based on Advances in Space Research 48 (2011). [5]Sakurai T. et al. Studies of melting ice using CO laser for ice drilling. Cold laser. Sakurai (2016) has demonstrated the capabilities 2 Regions Science and technology 121 (2016). of this concept. At AWL, the water sampling is peformed in two steps: i) the first 5 cm of ice (altered by the radiation) ORGANIC COMPOUNDS CHARACTERIZATION IN SPACE: EXPERIMENTAL ACTIVITY OF VISTA INSTRUMENT. F. Dirri1, E. Palomba1, A. Longobardo1, D. Biondi1, A.Boccaccini1, E. Zampetti2, B. Saggin3, D. Scaccabarozzi3. 1IAPS-INAF, Via Fosso del Cavaliere 100, 00133 Rome, Italy ([email protected], erne- [email protected], [email protected]), 2IIA-CNR, via Salaria km 29,300 Monterotondo, Rome ([email protected]); 3Politecnico di Milano, Polo Territoriale di Lecco, Lecco, Italy ([email protected], [email protected])

Introduction: VISTA (Volatile In Situ Thermogra- vimetry Analyser) is a μ-Thermogravimeter device, composed by a Piezoelectric Crystal Microbalance (PCM), equipped with built-in heater and built-in temperature sensor and the related Proximity Electronics (PE) [1]. The TGA analysis is frequently used to monitor the deposition/sublimation and absorption/desorption processes of volatiles compounds in different environments: outgassing rates from degassing in space [2], dehydration and organics decomposition in minerals [3] and the fate of various materials in atmospheric environments [4]. Thus, by means of VISTA instrument and using deposition and desorption processes, it is possible to have a Fig.1. SH1-BB laboratory set-up used for contamina- characterization of organic compounds. In particular, tion, saturation and TGA tests. organic compounds have been well characterized during the deposition and TGA thermal processes obtain- SH1-BB has been used as cold sink, i.e. at -20°C (dep- ing some physical-chemical parameters, e.g. enthalpy osition process) for contaminant molecules, to obtain of sublimation/evaporation, ΔH and vapor pres- sub,evap ΔH and to perform thermal cycles with the integrat- sures, P (at a given temperature and pressure). sub vap ed heaters (TGA measurement).

VISTA Instrument and applications: VISTA instrument has been studied for space applications and proposed for planetary in-situ misions, achieving different scientific tasks [1]. VISTA is composed by two different subsystems: a) Sensor Head 1 (SH1), able to work at low temperaures (down to -200°C) and to perform thermal cycles (up to 100°C); b) Sensor Head 2 (SH2), able to performs TGA measurements with a low power budget [1].

Performance Tests: VISTA SH1 Breadborad (SH1- BB) has been used to demonstrate the capability to detect contaminant depositions at low temperatures in

vacuum chamber. The setup used fro experiment is Fig.2. Contamination test and TGA cycles. shown in Figure 1. Contamination tests, saturation tests, regeneration tests The experimental procedure consisted in performing a (TGA) and organic compound characterization using contaminant deposition, i.e. Adipic acid, sample on the Langmuir relation [5] have been performed: sensing crystal's surface, heating the effusion cell up to 100°C, for a total deposition of 29.4 μgcm-2 (Fig. 2). Therefore, the crystal has been heated by using the (1) integrated heater (TGA cycles) up to 80°C (for regeneration) and ΔHsub has been calculated with eq. 1. The obtained result is in agreement with Albyn (2004), Saleh (2008), Bruns (2012) and Dirri (2016) within 5% and within the error with Booth (2010) (Tab. 1) and ΔHsub. The ΔHsub results obtained with SH2-BB and eq. show the good capability of VISTA SH1 to character- 2 are in agreement within 6% with the results obtained ize a contaminant sample. with SH1-BB and eq 1 (Tab.2). The vapour pressure average value for Adipic acid obtained with Calusius-Clapeyron and Langmuir equations is in agreement within 2 times the error bars of Test/ Adi- ΔHsub ± σΔH Trange (°C) Method Equation literature data [6]. pic acid (kJ/mol) PT-SH2 141.6±0.8 40-70 EM Van't Hoff Reference ΔHsub ± σΔH Trange Method DP-SH2 138.2±1.1 30-70 EM Langmuir CT II- SH1 (kJ/mol) (°C) 133.8±1.8 25-50 TGA Langmuir This work 135.3±1.3 25-72 TGA (3° cycle) ST - SH1 Dirri (2016) 141.6±0.8 40-70 EM 134.3±3.4 47-72 TGA Langmuir Bruns (2012) 136±10 71-110 ASAP-MS (2° cycle) ST - SH1 Booth (2010) 119±18 30-60 KEMS 138.2±2.2 44-69 TGA Langmuir Saleh (2008) 135±13 27-40 IVM (3° cycle) Albyn (2004) 129.3±2.5 33-55 EM Table 1. Enthalpy of sublimation results obtained in Tab.3. Enthalpy of sublimation results obtained with this work and compared with previous works. Effusion Method (EM) or deposition tests and Ther- mogravimetric Analysis (TGA). The enthalpy of sub- A dedicated setup has been also developed for SH2- limation results are the average values obtained from BB to characterize five organic compounds present in deposition rates data (DP tests, SH2) and desorption Earth's atmospheric aerosols, i.e. dicarboxylic acids, by rates data (ST and CT tests, SH1) in each temperature range listed. means of ΔHsub and Van't Hoff relation [13]:

(2) Moreover, the presence of some these organic compounds inside meteorites [7] make VISTA useful to know the chemical composition of these samples asso- The sample and the PCM were placed inside a vacuum chamber: the sample was heated from 25 to 80°C by ciated to primitive asteroids [8] (e.g. for Sample Re- means of a resistor whereas the PCM was cooled at - tunr Missions). 72°C by a cold finger (LiN2) in order to allow the molecules deposition coming from the sublimation process. By measuring two different deposition rates on References: [1] Palomba E. et al., (2017), 14th PCM, k1 and k2, at two different sample temperatures IPPW Abstract; [2] Wood B.E. et al., (1997), AIAA 97-0841; [3] Serpaggi F., et al. (1999), J. Solid State T1 and T2, it has been possible to infer the ΔHsub of the samples by means of eq. 2. Therefore, different deposi- Chem., 145, 580-586; [4] Elder J.P., (1997), Jour. tion curves (Fig. 3) in a similar temperature range have Thermal Analysis, v.49, pp.897-905; [5] Langmuir I. been obtained for five compounds. (1913), Phys. Rev., APS Jorn., 2, 329-342; [6] Bilde M. et al., (2015), Chem. Rev., 115, 4115−4156, 2015; [7] Andersen C.A. et al., (2005), Int. Jour. of Astrob., 4, 1, 13-17; [8] Dirri F. et al. (2016), OLEB Journ., doi:10.1007/s11084-016-9517-y.

Fig.3. Deposition rates of dicarboxylic acids.

The results are in good agreement with literature [6], demonstrating the VISTA SH2 capability to measure COSMORBITRAP: R&T DEVELOPMENT OF A NEW HRMS ANALYZER FOR FUTURE SPACE MISSIONS. C. Briois1, N. Carrasco2, G. Chalumeau1, F. Colin1, H. Cottin3, C. Engrand5, N. Fray3, B. Gaubicher1, N. Grand3, J.-P. Lebreton1, F.-R. Orthous-Daunay4, C. Pennanech3, S. Ruocco2, L. Selliez1,2, C. Szopa2, L. Thirkell1, V. Vuitton4, P. Zapf3 and A. Makarov6 1 LPC2E, Orléans, France, 2 LATMOS, Guyancourt, France, 3 LISA, Créteil, France, 4 IPAG, Grenoble, France, 5 CSNSM, Orsay, France, 6 ThermoFisher Scientific, Bremen, Germany; Corresponding author: [email protected]

The in situ exploration of the environments of various Solar System bodies with High-Resolution Mass Spectrometry (HRMS) requires the development of a new generation of space instruments.

Our team, the Cosmorbitrap Consortium, is developing for several years a new type of HRMS analyzer based on the use of the OrbitrapTM technology [1]. In laboratory and in commercial instruments this analyzer has allowed to reach mass resolution more than 105, up to 106 at m/z=200 [2],[3].

We are pursuing several international collaborations for the development of complete HRMS instruments based on the Cosmorbitrap coupled to various front- ends, each adapted to the application envisaged.

In this paper we provide a description of the Cos- morbitrap development and highlight the performances achieved so far with a laser ablation/ionization front-end. Examples of the envisaged space applications will be presented and the capabilities that would be offered by such HRMS instruments will be discussed.

References:

[1] Briois et al. (2016) Planetary and Space Sci- ence, 131, 33-45. [2] Makarov (2000) Analytical chemistry, 72, 1156-1162. [3] Denisov et al. (2012) Interna- tional journal of Mass Spectrometry, 325-327, 80-85. A Thermopile Based Heat Flux Sensor for Measuring Heat Flow on the Surface of Venus. Michael Pauken, Kevin L. Smith, Sutine Sujittosakul, Brian Phan, Samad Firdosy Billy Chun-Yip Li, George Nakatsukasa, Jean-Pierre Fleurial, Suzanne Smrekar. Jet Propulsion Laboratory, California Institute of Technology, 4800 Oak Grove Drive, Pasadena CA, 91109, [email protected].

Introduction: Among the great questions regarding (3) an array of thermopiles that are low thermal conduc- origins is “how are rocky planetary bodies formed?” tivity to create a temperature gradient and a resulting Measuring heat flow at a planet’s surface is one of a voltage in response to the heat flow and (4) graphite number of remote and in situ methods available for un- plates for rejecting heat flow to the environment. Insu- derstanding the structure of terrestrial planets. Heat flow lation encompasses the thermopiles to limit heat flow is a fundamental signature of planetary formation, struc- through the sensor except through the thermopile legs. ture and dynamic activity [1]. Radioactive decay and gravitational accretion are responsible for producing heat within a planet. This heat source can be responsible for molten layers which produce interior convection between the core and the crust. Planetary heat flow is typically non-homogeneous. Variations on Earth are a result of plate tectonics, crust composition and volcanic activity. On the moon, heat flow variations are thought to be a result of radiogenic material concentrations [2]. At Mars, variations are likely a result of volcanism. Global or widespread heat flow measurements on all the terrestrial bodies in the solar system will help formulate and evaluate models of planetary structure and evolution [4].

We present the design, development, fabrication and Figure 1. A schematic of the Venus Heat Flux Sensor, testing of a new planetary heat flow sensor specifically heat flows from the bottom through the carbon fiber in- designed to operate on the Venus surface. While it terface pad into the thermopiles. would not be able to provide global or widespread heat The Venus heat flow sensor is similar to thermoe- flow measurements on Venus (unless a number probes lectric (TE) generators that are composed of hundreds equipped with a heat flow sensor are deployed around of discrete low aspect ratio couples that produce output the planet), the sensor described here could provide di- power of 10’s to 100’s of watts. The sensor differs from rect in situ measurement of local surface heat flow. TE generators by using couples with a high aspect ratio, Measuring the heat loss at the surface of Venus as shown in Figure 2, to provide a voltage output signal could improve our understanding on the current geo- proportional to heat flow through the sensor. logic activity, lithospheric thickness and distinguish between various hypotheses of Venus’ evolution which can be classified into three different models; low (<20 mW/m2), intermediate (20-40 mW/m2) and high (>40 mW/m2) values of heat flow. Each model provides some insight on volcanism and plate tectonics activity in the past. The objective of this research is to develop a heat flux sensor capable of measuring heat flow from the surface of Venus over the range of 10 to 100+ mW/m2 with an accuracy of ±5 mW/m2 to differentiate between various thermal evolution models. The design of the sensor, as shown in Figure 1, consists of (1) a carbon fiber interface pad that deforms under light pressure to provide good contact with an irregular shaped surface, (2) a high thermal conductivity graphite plate that transfers the Figure 2. A high aspect ratio thermopile with 10 pairs heat flow from the interface pad into the thermopiles, of p-type and n-type skutterudite elements connected in series for signal amplification. during the transient response as it approaches equilib- The design approach assembles the high aspect ratio rium with the environment and the heat flow approaches thermopiles as shown in Figure 3. The thermopiles are zero. wired in series to provide a 90x voltage signal amplification from the temperature gradient generated during heat flow through the sensor. The bottom of the sensor picks up heat, while heat is rejected out the top of the sensor through graphite plates (not shown).

Figure 4. Predicted output voltage signal for the Venus heat flux sensor over the expected heat flux range.

References: [1] Pollack, H. N., et al. (1993). Rev. of Geophys., 31(3), 267-280. [2] Siegler, M. A., & Smrekar, S. E. (2014), JGR, 119(1), 47-63. [3] An- Figure 3. The inner workings of the heat flux sensor is drews-Hanna, J. C., et al. Science, 339(6120), 675-678. an array of 9 thermopiles brazed to a metal layer bonded [4] Hayne, P. O., et al, (2017) Planetary Science Vision to the graphite plate. 2050 Workshop.

An analytical model of the heat flux sensor has been Acknowledgement: This work was performed at the Jet developed to understand the performance and capability Propulsion Laboratory, California Institute of Technol- of the heat flux sensor and to guide the development of ogy, under contract with the National Aeronautics and a prototype sensor. The thermoelectric device-based Space Administration. sensor design constitutes a geometry that analytically satisfies the requirements for measuring the surface heat flux in the range of 10 to 100+ mW/m2 with a resolution of 5mW/m2. The Seebeck coefficient of the skutterudite elements is 340 mVolts/K at 750K, Venus’ surface temperature. This produces a voltage of 0.12 mVolts per element pair and for an array of 90 elements, the total voltage generated is 10.8 mVolts for the 5 mW/m2 heat flux resolution. The predicted voltage output from the sensor at both 25° and 470°C are shown in Figure 4. A series of tests are planned to evaluate the performance of the sensor’s response to heat flow. Initial tests will be at high heat flux values (on the order of 1 to 2 W/m2) by placing the sensor on a heated surface. A calibrated commercial heat flux sensor will be placed in series with the test article to measure the heat flux through the sensor under steady state conditions. A transient test with the heat flow sensor can be achieved while the heated surface cools down to ambient conditions. The resolution of the sensor would be determined IMPROVEMENT OF THE DYNAMICAL RESPONSE OF A SPHERICAL 3D WIND SENSOR FOR MARS ATMOSPHERE. M. Dominguez-Pumar, L. Kowalski, M.T. Atienza, S. Gorreta, V. Jimenez, L. Castañer MNT-Group. UPC-Campus Nord, Ed. C4. Jordi Girona 1-3. 08034 Barcelona.SPAIN.

Introduction: 3D wind sensing in Mars atmos- obtained model is summarized in what is called the phere is very challenging due to Mars low pressure (6- diffusive symbol of the structure. The diffusive symbol 12 mbar) and large temperature dynamical range of a first version of the sensor can be seen in Figure 4. (150K-300K). The objective of this paper is to present As it can be observed there are two significant time the improvement of the dynamical response of a spher- constants (5s and 67s), meaning that the time evolution ical 3D wind sensor. This sensor is a heritage of the under closed-loop operation is determined by the drift REMS (Remote Environmental Monitoring Station) of both state variables on the constant temperature con- wind sensor on board the Curiosity Rover. The main trol surface. The time response of this first version to a advantages of the proposed spherical geometry are the wind change can be seen in Figure 5. simplicity of the associated inverse problem (3D wind In a second version of this sensor, the thermal con- speed recovery from measurements), and the simplicity nection between the heaters and the spherical sectors of the system itself (6 controlled heaters in total). has been improved by using direct welding. The diffu- Sensor concept: The structure of the sensor can be sive symbol showed a reduction to only one significant observed in Figures 1 and 2, [1]. It is a spherical sensor time constant (30s), see Figure 6. The time response to composed of 4 sectors. These sectors are united to a a change in wind direction can be seen in Figure 7. As PCB (printed circuit board) that is used to support the it can be observed the system time response has been structure and provide electrical signal routing. Each greatly reduced to approximately 1-2 seconds. sector has a Pt resistor that is used both for temperature sensing and heating. Finally two additional Pt resistors References: [1] L. Kowalski, et al. (2016), IEEEE have been placed on the PCB also for temperature con- Sensors Journal, 16, 1887-1897. [2] M. Dominguez- trol. Pumar, et al. (2016) IEEE Trans. Ind. Electr., 64, 664- Sensor operation: The sensor is operated in a 673. closed loop enforcing constant temperature. The system output is the power required to keep constant the temperature of each sector. Under this mode of operation, it is possible to accelerate the dynamical response of the system. Main result: The main result of this paper is to show how the dynamical response of the sensor can be radically improved by ensuring a low number of state variables of the thermal filters associated with the sensor. In particular it is needed an excellent thermal connection between the controlling resistors and the spherical silver sectors. The thermal models of the structure used are based on Diffusive Representation [1]. The analysis of the dynamics of the thermal system working under closed loop operation has been made using the theory of Slid- ing mode controllers [2]. The main conclusion of this analysis is that in order to accelerate the time response under closed loop operation, it is necessary to have a Fig. 1. Schematic of the the tetrahedral wind sensor low number of significant state variables of the thermal used in the experimental measurements. Ra, Rb, Rc system. Therefore, ensuring a good thermal connection and Rd are the Pt resistors placed on each sector, and between heaters and the sectors is a key design factor. Rcore1 and Rcore2 are the Pt resistors controlling the Thermal modelling: The thermal models of the temperature of the supporting PCB. sensor are obtained from applying pseudo-random binary sequences of current to the heaters, while observing the time evolution of the temperature (open-loop actuation). An example can be seen in Figure 3. The

Fig. 2. Photograph of the prototype of the spherical

sensor anemometer. Fig.5. Experimental average power injected into one of the heaters of the two hemisphere wind sensor to PDLQWDLQ ǻ7 . At t=800s the wind is changed from 0.8m/s to 0.3 m/s.

Fig. 6: 9-order diffusive symbols for the sensor with improved thermal connection between heaters and Fig. 3. Top: Zoom of 50s of open-loop experi- sectors (same wind velocities as in Figure 6). mental data Vs predicted data. Middle: Injected PRBS current. Bottom: Wind velocity as a function of time.

Fig. 4: 9-order diffusive symbols of the first version of the sensor for 6 different wind velocities. Velocities: Fig.7. Experimental average power injected into the w1 = 0.5, 0.7, 0.9, 1.1,w6= 1.3 m/s and w6 = 1.5 m/s. sectors of the tetrahedral wind sensor. At t=20s the yaw angle is changed from 0º to 66º.

SOLID (SIGNS OF LIFE DETECTOR): A TRL 5-6 INSTRUMENT FOR WET CHEMISTRY ANALYSIS IN PLANETARY EXPLORATION. V. Parro1, M. Moreno-Paz1, J. M. Manchado1, M. García-Villadangos1, Y. Blanco1, E. Sebastian1, J. Romeral1, J. Sobrado-Vallecillo1, P. L. Herrero2, C. Compostizo2, and J. Gómez-Elvira1 1 Centro de Astrobiología (CAB, INTA-CSIC), Torrejón de Ardoz, 28850 Madrid, Spain, 2SENER Ingeniería y Sistemas SA, Las Arenas, Spain. [email protected].

Introduction: Liquid handling and solid sample emperature, and pressure sensors, as well as preparation for downstream wet analysis in planetary motors. exploration is challenging and not well established. 2. At -40 ºC, only tested the instrument re- This is one the main reasons for the very few wet ex- sistance, in off during the test. periments performed on the surface of Mars so far. 3. Testing sonication and the fluidics below 1 mb Life detection planetary missions like the Icebreaker and at -5 ºC. proposal to Mars [1] for finding organic biosignatures 4. Finally the temperature was increase to 80ºC in ice-rich soils of the northern plains, or the coming as a survival maximum temperature. In this missions to the Oceans Worlds Europa case we only tested the reading and data rec- (https://solarsystem.nasa.gov/news/2017/02/08/nasa- orded by the electronics. receives-science-report-on-europa-lander-concept) or Enceladus will benefit of the possibility to detect biomolecules in liquid extracts. We have developed SOLID (Signs Of Life Detector), an immunoassay- based instrument for detecting molecular biomarkers in planetary exploration [2]. Current SOLID3.1 consists of two functional units, the Sample Preparation Unit (SPU), for the extraction of organics into a liquid solution, and the Sample Analysis Unit (SAU), for immu- nological assays in antibody microarray format. SPU Technical Description: The SPU (160x170x230 mm and 4.25 kg), consisting of a single extraction cell, can process up to 30 x 0.5 g powder, soil, or ground or melted ice samples to produce a

liquid extract that can feed downstream analytical instruments. After sample loading, the procedure in- Figure 1. SOLID_SPU showing the main elements volves: i) addition of liquid solvent into the extraction (left), and the functional diagram (right) cell (EC); ii) homogenization and extraction of organic material into the solvent by ultrasonication; iii) evacua- Conclusion: We have tested the subsystems and tion of whole sample (solids plus solvent) into a depos- the whole SOLID-SPU system under Mars relevant it with cylindrical micrometer pore size walls; and iv) conditions. All the systems and components survived injection of filtered liquid extract into downstream at least 3 cycles with temperatures from -40 ºC to + 80 analytical instruments (e.g. the immunoassay LDChip ºC. Additionally, we successfully performed the func- or Life Detector Chip [2], or any other technique that tional tests below 1 mbar and -5ºC including liquid requires a liquid extract). Figure 1 shows a functional pumping, ultrasonication, and filtering. diagram of SPU, with valves, motors, sonicator, a pump and temperature and pressure sensors in several References: [1] McKay et al., (2013) Astrobiology, locations of the unit. 13, 334-353. [2] Parro, V. et al. (2011) Astrobiology Ultrasonication and liquid handling under Mars 11: 15-28. [3] Sobrado-Vallecillo, J. M. et al. (2014) relevant conditions with SOLID-SPU. In order to American Institute of Physics: Review of Scientific increase the maturity level, we performed a series of Instruments 85: 035111. tests under relevant environment: low pressure and Additional Information: Funded by Spanish Min- temperature in a Mars simulation chamber [3]. The set istry MINECO No. AYA2011-24803. of tests consisted in three replicate cyles comprising the next steps: 1. At -20ºC we tested the electronics and motion module performance, including the position, BAROCAP® NG: THE NEXT GENERATION PRESSURE SENSOR FOR PLANETARY ATMOSPHERIC SCIENCE. H. Kahanpää1,2, M. Hieta1,2, J. Polkko1, M. Genzer1, T. Nikkanen1,3 and A.-M. Harri1, 1Finnish Meteorological Institute, P.O. BOX 503, FI-00101, Helsinki, Finland ([email protected]), 2Aalto University / School of Electrical Engineering, P.O. BOX 15500, 00076 AALTO, Finland, 3Reaktor Space Lab, Mannerheimintie 2, 00100 Helsinki, Finland.

Introduction: Barocap® Next Generation (NG) is a gas pressure sensor developed by Vaisala Corporation (Figure 1). The Finnish Meteorological Institute has qualified this sensor for space applications. It has been used in the DREAMS meteorological instrument [1] onboard Exo-Mars Schiaparelli and selected to be used in the meteorological instruments of the Mars 2020 and ExoMars 2020 landers. We present results of the qualification tests of Barocap® NG and show that this sensor has several benefits compared to similar pressure sensors used in previous Mars landers.

Figure 2: Structure of the Barocap® NG pressure sensor

Figure 1: Barocap® pressure sensor. Right-bottom: three sensor chips. Top: sensor chip mounted on support plate. Left-bottom: sensor inside package with pressure inlet

Heritage: Before ExoMars Schiaparelli, the Finnish Meteorological Institute had provided pressure

instruments based on earlier versions of the Vaisala Figure 3: Principle of the Barocap® NG pressure Barocap® sensor for six planetary probes: Mars 96 [2], sensor Mars Polar Lander [3], Beagle 2 [4], Huygens/Cassini [5], Mars Phoenix [6][7] and MSL Curiosity [8]. All of In Barocap® NG the distance between the plates these sensors have worked without failures till the end reduces with growing pressure. This is a difference of the missions. compared to the LL- and RSP-type Barocap® sensors Technology: Vaisala Barocap® is a silicon-based used in previous Mars landers [8]. The upper limit of micromechanical pressure sensor. It is basically a plate the dynamic range is limited by the pressure where the capacitor whose capacitance is changed by pressure. electrodes touch each other's. This pressure depends on The other electrode is attached to a membrane made the thickness of the membrane and the gap between the out of single-crystal silicon and the other electrode is electrodes in vacuum. Sensors for different pressure stationary (Figure 2). The surrounding pressure ranges up till 10 bar can be developed by tuning these deforms the membrane, thus changing the distance two parameters. between the electrodes and hence also the capacitance of the sensor (Figure 3). Another difference compared to the LL- and RSP- Results: The test data was used to determine type Barocaps® is the mounting of the sensor chip to performance characteristics of the Barocap® NG tuned the support plate (Figure 1). The chip is attached from for the Martian pressure range. These characteristics the side so that thermal expansion do not deform the are summarized in Table 1 and compared to the chip. characteristics of the older Barocap® types RSP2M The Barocap® NG sensor weighs only 2 g. With and LL [8] used in previous Mars landers. proximity electronics included the masses of the As can be seen from Table 1, Barocap® NG has previous space instruments based on Barocap® sensors better resolution and repeatability than what RSP2M have weighted circa 40 g. The power consumptions of and LL have. The most important improvement is these instruments have been circa 15 mW. however the stability of temperature dependence. The Tests: The Barocap® NG was qualified for the temperature of a pressure sensor inside a Mars lander ExoMars Schiaparelli mission by the test campaign of may wary strongly during a sol. For example the Mars the qualification models of the DREAMS pressure Pathfinder pressure sensor experienced diurnal sensor [1]. These tests comprised of environmental temperature variations of 35 °C [9] and the Phoenix tests and calibration checks performed between the pressure sensor variations of 60 °C [7]. Potential environmental tests. The following environmental tests changes in temperature dependence between were performed: Random vibration, Pyroshock, Bake- calibration and mission were notable uncertainty out (16 h in +110 °C, 800 Pa) and Thermal Vacuum (8 sources in both cases [7][9]. Barocap® NG is almost cycles, -45 … +55 °C). The calibration checks were immune to this error source due to its stable performed in the Martian pressure range (400 … 1200 temperature dependence. According to initial results Pa) in four temperatures ranging from -45 °C to +55 also the absolute stability of Barocap® NG is better °C. than that of RSP2M.

Table 1: Performance characteristics of Barocap® pressure sensors used in Mars landers

Barocap® type NG for Martian RSP2M LL pressure Used in missions ExoMars Schiaparelli ExoMars Schiaparelli, MSL, Phoenix, Huygens, MSL Polar Lander, Mars 96 Resolution (time scale of seconds) 0.1 Pa 0.2 Pa 0.2 Pa Repeatability (hourly time scale) < 1 Pa < 2 Pa < 2 Pa Absolute accuracy (sol 0) 3 Pa 3 Pa 3 Pa Drift rate 1 0 to 2 Pa / year 3 to 5 Pa / year 0.5 Pa / year Temperature dependence 4 Pa / °C < 2 Pa / °C 3 to 15 Pa / °C (without compensation) Stability of temperature < 0.025 Pa / °C < 0.3 Pa / °C < 0.1 Pa / °C dependence 2 Warm-up time 3 200 s 1 s 150 s Outgassing time in 60 °C, 0 Pa ~7 days 1 day 1 day Testable in room pressure No Yes Yes Available Yes Yes No

1 Drift rates were determined using the data measured during the interplanetary cruises of MSL and ExoMars Schiaparelli 2 The stability of temperature dependence is a measure of how much the temperature dependence of the sensor may change between calibration and mission due to environmental stresses 3 The warm-up time is defined as the time between power-on and the instance when the pressure reading has stabilized

Conclusions: Our test results show that the performance of Barocap® NG fulfills the requirements of a barometric pressure sensor of a Mars lander. The possibility to adjust the dynamic range of the sensor by tuning its mechanical properties makes it also a possible sensor candidate for e.g. Titan probes. Table 1 shows that the benefits and drawbacks of Barocap® NG and Barocap® RSP2M complement each other's. Both of these sensor types can be used in the same instrument as they share common proximity electronics. This concept is used in the pressure sensors of the Mars 2020 and ExoMars 2020 landers. The expectation is that these instruments will have the benefits of both the Barocap® NG and Barocap® RSP2M sensors. References: [1] Esposito F. et al. (2017) Space Sci. Rev., Submitted. [2] Harri A.-M. et al. (1998) Planet. Space Sci., 46, 779–793. [3] Paige D. A. et al. (1998) LPI Contributions, 953, 30. [4] Towner M. C. et al. (2004) Planet. Space Sci., 52, 1141–1156. [5] Harri A.-M. et al. (2006) Planet. Space Sci., 54(12), 1117- 1123. [6] Taylor P. A. et al. (2008) JGR, 113, E00A10. [7] Taylor P. A. et al. (2010) JGR, 115, E00E15. [8] Harri A.-M. et al. (2014) JGR Planets, 119, 82–92. [9] Haberle R. M. (1999) JGR, 104(E4), 8957–8974. MELDI2 DO NO HARM TEST SERIES. G. T. Swanson1, J. A. Santos2, T. R. White3, W. E. Bruce4, C. A. Kuhl4 and H. S. Wright4, 1AMA Incorporated, NASA Ames Research Center, Moffett Field, CA 94035 USA, g.swan- [email protected], 2Sierra Lobo, NASA Ames Research Center, 3NASA Ames Research Center, 4NASA Langley Re- search Center

Abstract: Mars 2020 will fly the Mars Entry, De- This presentation will discuss the planning and exe- scent, and Landing Instrumentation II (MEDLI2) sensor cution of the MEDLI2 DNH test series. Selected high- suite consisting of a total of seventeen instrumented lights and results of each environmental test will be pre- thermal sensor plugs, eight pressure transducers, two sented, and lessons learned will be addressed that will heat flux sensors, and one radiometer embedded in the feed forward into the planning for the MEDLI2 flight thermal protection system (TPS). Of the MEDLI2 in- system certification testing. strumentation, eleven instrumented thermal plugs and seven pressure transducers will be installed on the heatshield of the Mars 2020 vehicle while the rest will be installed on the backshell. The goal of the MEDLI2 instrumentation is to directly inform the large performance uncertainties that contribute to the design and validation of a Mars entry system. A better understanding of the entry environment and TPS performance could lead to reduced design margins enabling a greater payload mass-fraction and smaller landing ellipses. To prove that the MEDLI2 system will not degrade the performance of the Mars 2020 TPS, an Aerothermal Do No Harm (DNH) test series was designed and conducted. Like Mars 2020’s predecessor, Mars Science Laboratory (MSL), the heatshield material will be Phe- nolic Impregnated Carbon Ablator (PICA); the Mars 2020 entry conditions are enveloped by the MSL design environments, therefore the development and qualification testing performed during MEDLI is sufficient to show that the similar MEDLI2 heatshield instrumentation will not degrade PICA performance. However, given that MEDLI did not include any backshell instrumentation, the MEDLI2 team was required to design and execute a DNH test series utilizing the backshell TPS material (SLA-561V) with the intended flight sensor suite. To meet the requirements handed down from Mars 2020, the MEDLI2 DNH test series emphasized the interaction between the MEDLI2 sensors and sensing locations with the surrounding backshell TPS and sub- strucutre. These interactions were characterized by per- Figure 1. Front (top) and Backside (bottom) of the forming environmental testing of four 12” by 12” test MEDLI2 DNH Test Panel panels, which mimicked the construction of the backshell TPS and the integration of the MEDLI2 sensors as seen in Figure 1. The testing included thermal vacuum/cycling, random vibration, shock, and arc jet testing. The test panels were fabricated by Lockheed Mar- tin, establishing techniques that will be utilized during the Mars 2020 vehicle installation. Each test panel included one thermal sensor plug (two embedded thermocouples), one heat flux sensor, and multiple pressure port holes for evaluation. HIGH TEMPERATURE CHARATERIZATION OF GALLIUM NITRIDE ULTRAVIOLET PHOTODETECTORS FOR IN-SITU SHOCK-LAYER RADIATION MEASUREMENTS. R. A. Miller1 and D. G. Senesky2, 1PhD Candidate, Department of Aeronautics and Astronautics, Stanford University, 2Assistant Pro- fessor, Department of Aeronautics and Astronautics, Stanford University.

Abstract: Thermal protection systems designed to protect payloads from the high heat load experienced during atmospheric entry are commonly sized based on convective heating. However, depending on the atmosphere, size of the vehicle, and velocity of entry, radiative heating can dominate [1]. For example, the radiative heating for a Titan mission is predicted to be four times greater than the convective heating with much of the heating in the ultraviolet (UV) regime [2]. Current commercial sensor technology is based on silicon which is not well suited for high temperature UV photodetection [3]. Gallium nitride (GaN) is a wide bandgap sem- iconductor that has been shown to be thermally stable as Figure 1. Test set-up for characterizing GaN-on-sap- well as highly responsive to incident UV radiation mak- phire photodetectors at elevated temperatures. ing it a good candidate material for high-temperature UV photodetection [4]. In this study, GaN-on-sapphire ultraviolet photodetectors were characterized at elevated temperatures as shown in Fig. 1. The current-voltage curves plotted in Fig. 2 show an increase in dark current as temperature increases from 22°C to 100°C which can be attributed to the increase in thermal carriers. At elevated temperatures the themal carriers begin to dominate the photo- generated carriers (normalized photocurrent-to-dark current ratio is reduced from 167.6 mW-1/2 at 22°C to 35.8mW-1/2 at 100°C). Nevertheless, the room temperature characteristics remain unchanged after this themal cycling. In addition to understanding the high tempera- Figure 2. Dark and illuminated (365nm) current-volt- ture response of these devices, testing in NASA’s Elec- age curves for the GaN-on-sapphire photodetectors for tric Arc Shock Tube (EAST) facility demonstrated the room temperature (22°C) and 100°C. feasibility of using GaN-based photodetectors to detect shock-layer radiation in a Titan simulate environment. The measurements made by the GaN photodetectors are in good agreement with the average UV radiance measured by the EAST facility spectrometers. While the electrical response of the photodetectors was not meas- urably degraded by the shock, Fig. 3 shows the packaged photodetector chip pre- and post-shock testing indicating the packaging is the limiting factor. References: [1] Cruden B. A. Prabhu D. and Martinez R. (2012) Figure 3. GaN-on-sapphire photodetector chip installed Journal of Spacecraft and Rockets 49, 1069-1079. [2] Bose in a shock tube port holder for testing in NASA’s EAST D. Wright M. J. Bogdanoff D. W. Raiche G. A. and Allen facility: (a) before shock testing and (b) after shock test- Jr. G. A. (2006) Journal of Thermophysics and Heat Trans- ing with the direction of the shock indicated. fer, 20, 220-230. [3] Lien W.-C. Tsai D.-S. Lien D.-H. Senesky D. G. He J.-H. and Pisano A. P. (2012) Electron Device Letters, 33, 1586-1588. [4] Monroy E. Omnes F. and Calle F. (2003) Semicond. Sci. Technol., 18, R33-R51. REMOTE RECESSION MEASUREMENTS OF SEEDED PICA. Bradley D. Butler1, Michael Winter, Alexan- dre Martin, and Sean C.C. Bailey. University of Kentucky, Department of Mechanical Engineering, 151 Ralph G Anderson, Lexington, KY, 40506. ([email protected])

Introduction: Typically the characterization of trometer that spatially resolves the stagnation line ablative material recession relies on thickness meas- ahead of the sample in the 220 nm to 550 nm spectral urements before and after testing, while the char depth range. Additionally, some of the test specimens were is measured by sectioning core extractions. These also observed using miniature fiber-coupled spectrom- methods are only able to provide information integrat- eters which combined cover a spectral range from 200 ed over the test time or flight time of a recoverable nm to 1700 nm. Earth entry mission. A spectroscopic method for de- Results: Feasibility tests at NASA Ames were car- termining the rate of recession of Phenolic Impregnat- ried out with seed materials extruded within an epoxy ed Carbon Ablator (PICA) material in real time has matrix to form seeding rods, which were then inserted been tested in different arc-jet facilities. Through the from the side into the samples. The second round of use of an onboard spectrometer imaging the post-shock arc-jet testing[2], conducted at NASA Langley, proved region, this method has the potential to be used for the encapsulation of seed materials in an epoxy carrier remote time resolved measurements of ablative heat to be impracticable due to the early pyrolysis of the shield performance. epoxy leaving a step face in the material when reces- Method: The basic idea is to seed the thermal pro- sion reached the seeding depth. For the seed material tection with a tracer material, with a distinguishable infusion method elemental Al was seen in the post- spectral signature, at a known depth and position. shock region ahead of the recession reaching the seed- When the material recession or char reaches the seeding depth, which indicates a potential for char depth or ing depth the seed material is gasified and the spectral internal temperature indication. Seed materials with signal is detected by optical emission spectroscopy. A higher melting points such as VB2 and TiC showed rate of material recession or charring can be deter- great promise as recession indicators. The performance mined based upon the properties of the seed material of the metallic wire seeding will be compared to that of and the time resolved spectra. The working principle is the infused material. illustrated in Fig. 1 with spectra shown from a feasibil- [1] Winter, M., et al. (2014) Remote Recession ity test[1]. Sensing of Ablative Heat Shield Materials, AIAA Aer- Three methods of seed material delivery and ospace Science Meeting. placement have been evaluated over the course of three [2] Butler, B., et al. (2016) Characterization of arc-jet test campaigns at NASA Ames and NASA Candidate Materials for Remote Recession Measure- Langley. The methods include encapsulation in an ments of Ablative Heat Shield Materials, AIAA Aero- epoxy disc, seed impregnation during PICA pro- space Science Meeting. cessing, and also the use of metallic wires. Various seed materials that have been tested including Al, Al2O3, V, VB2, Ti, TiC, Ag, In, Si, SiO2, HfO2, HfC, NaCl, and MgCl2. All test specimens were observed using a spec-

Fig. 1: Illistration of remote recession sensing technique; shown spectra obtained during feasability testing[1]. RECONSTRUCTING TPS RECESSION USING THE “ReGS” SENSOR. G. Vekinis and A. Marinou, Insitute of Nanoscience and Nanotechnology, NCSR Demokritos, Agia Paraskevi Attikis, 15341, Greece, [email protected]

Introduction: Being able to monitor in-situ the In Figure 2 the raw data of three flame tests using recession of an ablator Thermal Protection System ReGS sensors in ZURAM are shown. (TPS) is useful for developing new materials and correctly sizing the ablator shield for space missions requiring high speed atmospheric entry. A number of TPS recession sensors [e.g. 1-3] have been developed and some have been used in space missions in the past. Most rely on the ablator char completing an electrical circuit whose response changes with recession as the temperature profile shifts. This approach has proved rather unreliable in many cases since the char’s characteristics are often unpredictable and circuit continuity is generally not guaranteed.

The ReGS sensor: The “ReGS” (“Resistive Grid Figure 2. Raw data from three ReGS-in-Zuram reces- Sensor”) TPS recession sensor is currently being de- sion tests using oxy-acetylene flame burning. veloped with ESA funding [4] on the basis of the previously reported “ReWiG” sensor [5] and has been The data show that there is significant variation in re- used to monitor recession in the ZURAM [6] ablator as cession rate depending on how the tests were carried well as other ablating materials. out. In the test “short-4cm”, the gradual decrease in the The ReGS sensor consists of a thin metallic grid gradient of the curve indicates that the heat flux on the whose resistance increases almost linearly for the first surface of the ZURAM is decreasing, indeed as ex- half of its length as the grid elements melt away when pected since no effort was made to keep the flame- embedded in the ablator, as shown in Figure 1. surface distance constant. In the other two curves, at- tempts were made to keep the recession rate approximately constant but the data show that this was only partially successful as the curves show substantial non- linearity at various points during the test. Recession reconstruction: By comparing the data in Figure 2 with the known sensor response, it is possible to reconstruct the recession behaviour in each test and this is shown in Figureg 3.

Figure 1. A recessed ReGS sensor in ZURAM after burning using an oxy-acetylene flame.

Two types of ReGS sensor are under development: a “short” sensor with a capability of measuring recession up to about 6mm with an optimum resolution of about ±0.3mm and a “long” sensor for recession depths up to 12mm and a resolution of about ±0.6mm. Figure 3. Reconstructed recession-time curves from the tests in Figure 2. Various events can be discerned. Recession tests: Tests have been carried out using both oxyacetylene flame as well as plasma jet (which A comparison between the recession measured us- will be presented elsewhere). ing the ReGS sensor and the actual recession determined post-mortem is shown in Figure 4 for a number of different tests.

In all cases the sensor signal indicated slightly lower recession than the maximum recession in the ZURAM (which generally occurs close to the sensor in these tests) but higher than the average recession in the ZURAM. Development of the sensor has been aided by numerical modelling of the recession of the Zuram and the sensor which is presented in another paper in this conference. The work is continuing.

References: [1] E. Martinez and T. Oishi, “Current Developments in Sensors for Thermal Protection Systems”, Proc. of the 5th Eur. Workshop on TPS and Hot Structures, 17-19 May, 2006 at ESA/ESTEC, The Netherlands. Ed. K. Fletcher, ESA SP- 631, 2006 [2] T. Oishi, E. R. Martinez and J. A. Santos, “Development and application of a TPS ablation sensor for flight”, 46th AIAA Aerospace Sciences Meeting and Exhibit, 7-10 Janu- ary, Reno, Nevada, USA [3] G. Papadopoulos, N. Tiliakos, G. Benel, “Non-Intrusive Sensor for in-situ Measurement of Recession Rate of Heat Shield Ablatives”, AIAA Ifoftech@Aerospace 2010, 20 - 22 April 2010, Atlanta, Georgia, USA [4] ESA/ITI contract B000016985, 2016 [5] G. Vekinis, “ReWiG: a resistive grid TPS recession sensor”, IPPW12, 15-19 June 2015, Cologne, Germany [6] A. S. Pagan, J. Rieser, B. MassutiBallester, G. Herdrich, “Characterisation of the Lightweight Ablative Heat Shield Material ZURAM® in HighEnthalpy Air Flows”, 8th Euro- pean Workshop on TPS and Hot Structures, ESA/ESTEC, 19-22 April 2015, The Netherlands. EXPERIMENTAL EVALUATION OF AN IN-FLIGHT SURFACE RECESSION SENSOR FOR ABLATIVE THERMAL PROTECTION SYSTEMS. B. Massuti-Ballester1, A.S. Pagan1, G. Herdrich1, G. Vekinis2, A. Marinou2, Ch. Zuber3, and H. Ritter4, 1University of Stuttgart, Institute of Space Systems, Pfaffen- waldring 29, 70569 Stuttgart, Germany ([email protected], [email protected], [email protected] stuttgart.de), 2National Center for Scientific Research “Demokritos”, Advanced Ceramics and Composites Laborato- ry, Patr. Gregoriou E' & 27, Neapoleos str., PO Box 60037, Postal Code 153 41, Agia Paraskevi, Attica, Greece ([email protected], [email protected]), 3German Aerospace Centre Stuttgart, Institute of Structures and Design, Pfaffenwaldring 38, 70569 Stuttgart, Germany ([email protected]), 4ESA-ESTEC, Keplerlaan 1, NL-2200 AG Noordwijk, The Netherlands ([email protected]).

Abstract: Ablative Thermal Protection Systems (TPS) remain a critical and indispensable subsystem for spacecraft subjected to hyperbolic entry trajectories such as planetary probes or sample return missions. Depending on the specific mission profile, the heat shield may take up a considerate fraction of the probe’s mass budget. A more precise quantitative understanding of material ablation in actual flight conditions, as expressed through the resulting surface recession rate, is thus invaluable towards optimizing payload mass fractions and accordingly improving a mission’s scientific potential by effectively reducing the as of yet gen- erous margins on the dimension of ablative heat shields. A means of monitoring the recession of an entry probe’s heat shield in-flight would provide unadulter- ated data directly reflecting the performance of a given material and design, thus providing a concrete baseline for less conservative future design iterations. An in-situ surface recession sensor for ablative TPS, developed at Demokritos, is embedded into a cylindrical sample of the ablative heat shield material ZURAM®, developed at DLR Stuttgart and character- ised at the Institute of Space Systems (IRS) in Stuttgart [1], which is then subjected to simulated hyperbolic Earth re-entry conditions in the PWK1 plasma wind tunnel facility at IRS. The sensor’s performance is assessed and verified both through integral physical measurements and the additional application of a laser- based photogrammetric measurement setup [2]. References: [1] A.S. Pagan, Ch. Zuber, J. Rieser, B. Massuti-Ballester, G. Herdrich, H. Hald, S. Fasou- las. (2016) Characterisation of the Lightweight Abla- tive Heat Shield Material ZURAM® in High-Enthalpy Air Flows. Proc. 8th European Workshop on TPS & HS. [2] A.S. Pagan, B. Massuti-Ballester, T. Mayer, G. Herdrich, S. Fasoulas, R. Ogawa, Y. Kubota, K. Ya- suo, H. Hatta (2015). Proc. 8th European Symposium on Aerothermodynamics for Space Vehicles.

Antarctic Testing of the European Ultrasonic Planetary Core Drill (UPCD): Lessons Learned from the Field.

Ryan Timoney1, Kevin Worrall1, Xuan Li1, David Firstbrook1, Patrick Harkness1

1 Space Systems Engineering, School of Engineering, University of Glasgow, Glasgow, UK, G12 8QQ

Introduction: The Ultrasonic Planetary Core Drill (UPCD) project, funded by a ~ €2.5M European Union Seventh Framework grant, has seen a consortium of European partners develop a sample acquisition and caching system for future exploration of terrestrial and icy planetary bodies. The ultrasonic/sonic drilling technique which forms the basis of the technology was pioneered by NASA JPL [1] at the turn of the 21st century. The Uni- versity of Glasgow, principal investigators of the UPCD project, have worked towards optimizing the technique [2] and integrating the technology into an architecture which allows multi-drill bit coring and caching through Figure 2: Coal Nunatak Frost Polygon Fields a novel application of the bayonet connection method [3, 4]. The field test campaign allowed the team to push the The UPCD project culminated in a field trial at Coal complete UPCD architecture (Drill System, Sample Nunatak, Antarctic Peninsula, in the Antarctic summer Caching Carousel and Z-Axis Vertical Actuator, Figure 2016 (Figure 1). The presence of geology found only in 2) to the limits of its capability. The relatively unknown the polar regions of Earth and Mars such as a perma- subsurface ensured that the team would be drilling frost, frost polygons and sloped lineae (Figure 2) quali- blind, analogous to the conditions which are experi- fied the site as a suitable location for testing the hard- enced by instrument teams on existing and future plan- ware. etary missions. Testing the system in a Mars analog environment paves the way to achieving an advanced Technology Readiness Level (TRL) 5-6, further enhancing the complement of instrumentation available to the European and worldwide planetary science communities.

Lessons Learned from the Field: The field site was chosen with guidance from the British Antarctic Survey (BAS) who operate within the British Antarctic Territories in which Coal Nunatak lies. The location is deemed to be analogous to the conditions which might be expected in other polar locations which are more commonly used as Mars analog sites (Antarc- tic Dry Valleys and the Haughton Impact Crater, Devon Figure 1: Coal Nunatak Field Site Island), though had never been utilised by instrumentation teams seeking to test their hardware. This provided the UPCD team with an exciting opportunity to work in a relatively unexplored environment coupled with the peace of mind afforded by relying on the operational and logistical expertise of BAS, a longstanding leader in polar exploration. Over the course of a ten-day expe- dition, the UPCD team was met with a number of challenges which it attempted to overcome in order to push the drill harder, exposing areas where the system was most robust and where the technology was lacking. The terrain proved to be extremely challenging, providing the team with a breadth of lessons learned regarding, predominantly, the need for higher than anticipated drilling torques, amongst other findings which would only have been unearthed in a field test campaign such as that which was undertaken. Whilst the need for greater rotational performance was identified, the system proved itself as having a notable robustness in the assembly of complete drill strings, and the subsequent disassembly and caching of these sample-containing core drill bits. This area of research is particularly novel, thus the results are highly encouraging and may confer an exciting and novel ability to future mission planners.

Figure 3: UPCD Complete Architecture

We at the University of Glasgow are eager to present our findings to the worldwide planetary science community at the 14th International Planetary Probes Work- shop.

References: [1] Y. Bar-Cohen et al, (2000), SPIE Smart Struc- tures, 3992, 101. [2] P. Harkness et al, (2011), IEEE Ul- trasonics Control, Vol 58, 11. [3] R. Timoney et al. (2015). AIAA Space. [4] R. Timoney et al. (2016) ASCE Earth and Space

THE PERFORMANCE OF ULTRASONICALLY-ASSISTED HAMMER-ACTION PENETRATORS IN PLANETARY REGOLITH. David Firstbrook1, Kevin Worrall1, Ryan Timoney1, and Patrick Harkness1. 1University of Glasgow, Glasgow, G12 8QQ, U.K., [email protected].

Introduction: The subsurface of planetary bodies serves as a fascinating area of research and science opportunities. Protected from the often harsh environmental conditions present, it poses scientifically interesting region in which to find life outside of our planet [1]. Due to the low gravity of these planetary bodies, this is not an easy task. Accessing depths on the order of a few meters on Earth is normally done with compli- cated drill-string structures, however the difficulties of conducting these operations autonomously on the surface of another planet mean these are not often the best methods. Self-propelling penetrating moles offer an attractive solution to this, offering tens of meters depth at a fraction of the overhead weight. These moles utilize an internal mass to hammer the probe through the granular surface, and have been used in several missions such as The advantages of this manifest themselves into Beagle-2, the Philae lander, and the 2018 InSight mis- one of two ways: sion [2, 3]. A balance between progression rate and Firstly, using the same hammer mass, a mole would structural damage to the internal components of the be able to travel further into the ground, shortening the probe is essential, as impulses can reach potentially time required for travel. For small landers that operate destructive levels, as demonstrated with the PLUTO from battery power alone, this would enable a longer probe aboard Beagle-2 exhibiting hammer strikes as duration in which to conduct measurements at the tar- high as 8,000 m/s2 [2]. get depth. Ultrasonic vibration has been shown to reduce Secondly, reducing the hammering force would get over-head force requirements under progression the mole to the target depth in the same length of time through granular material [4]. The work covered in this whilst dramatically reducing the impact load. This paper will discuss the first steps into combining this could potentially enable more delicate and sensitive technology with the hammering process. A free-falling devices to be placed within the mole opening up an weight was used to provide the hammering impulse, as extended suite of scientific instruments that could be opposed to the spring-loaded method used for HP3, utilised. however the height of the drop was designed to deliver It could even be possible to form a compromise of an impact energy of 0.8 J, mimicking the energy used these two methods, simultaneously accessing greater in the HP3 tests [3]. depths with lower impact forces. A titanium penetrator, tuned to resonate at 20 kHz, References: [1] Hassler D. M. et al. (2014) Science with a 60º conical tip at 17 mm diameter was designed 343 (6169), 1244797. [2] Glaser D. et al. (2008) Mete- oritics & Planet. Sci., 43 (6).1021-1032. [3] Hansen- and subjected to the hammering process. Ultrasonic th excitation amplitudes between 0.4 and 1.6 μm were Goos H. et al. (2014) 45 Lunar Planet. Sci. Conf. pulsed simultaneously with the hammering-action, and Proc. 1325. [4] Firstbrook D. et al. (2017) Proc. R. the depth measured after each stroke. Soc. A 473 (2198) 1-16. An example of the initial results using a medium- grained quartz sand are shown in the following figure. The tests using vibration show that the final penetration depth of 190 mm can be attained in significantly fewer hammer strikes.

LITHOSPACE: AN AUTOMATED SYSTEM FOR IN SITU PETROGRAPHIC THIN SECTION PREPARATION ON MARS. F. Foucher1, N. Bost1, S. Janiec2, F. Westall1, P. Perron3, A. Fonte4, N. Le Breton2, J. Li5, T. Platel5, M. Tagger6, M. Viso7, P. Chazalnoël8, F. Courtade8 and M. Villenave8, 1CNRS, CBM, Orléans, France, 2ISTO, Université d’Orléans, France, 3Lycée Benjamin Franklin, Orléans, France, 4PRISME, Unisversité d’Orléans, France, 5Polytech’ Orléans, France, 6LPC2E, CNRS-Université d’Orléans, France, 7CNES, Paris, France, 8CNES, Toulouse, France. ([email protected])

Introduction: Optical microscopy in transmitted The LithoSpace project: The aim of the Litho- light is probably one of the most standard techniques in Space project, supported by the CNES since 2014, is to geology. It has been used for more than 150 years. work on the development of an automated system Indeed, atlases of microscopic images of rocks and permitting preparation of petrographic thin sections on their characteristic mineralogy and textures are widely extraterrestrial bodies, in particular, on Mars. Several used in geological departments [1]. Observation in studies and tests were conducted and most of the prob- transmitted light permits identification of rocks having lems solved. The final protocol follows the process similar chemical and mineralogical compositions but from obtaining a standardized drill core to observation different mineral organizations (basalt vs. gabbro for of a thin section in a fully automated way. instance). Using polarized light, it permits identifica- Proof-of-concept: The objective of the project is tion of most of rock-forming minerals in thin section now to make a “proof-of-concept” model. Thus, a first [2]. Observation in transmitted light is also essential numerical model of the system has been designed by for micropaleontology since it is the only way to ob- students in fifth year at the engineering school Poly- serve individual microfossils, if they are large enough tech’, University of Orléans. With the help of students (Fig.1), and to document the mineralogical and textural in BTS (two-year technical degree) “industrialization context in which the biosignatures occur. Nevertheless, of mechanical products” and in professional license of optical observation in transmission has never been “technical coordinator of industrialization methods”, carried out on Mars. Indeed, although optical micro- from the Benjamin Franklin high school of Orléans, scopes can be readily designed for space exploration, the objective is now to make this model real. Finally, thin section preparation is not easy to do in situ. In this study is a good way to raise student awareness of particular, it requires several human manipulations that space exploration by involving them directly at the are difficult to automate. start of the project. On the use of thin sections for in situ space exploration: In situ thin section preparation would be an important improvement for the geological and astrobiological exploration of the solar system, for Mars but also of other planetary bodies. The proposed instrument could incorporate new high resolution techniques, such as Raman mapping or micro-LIBS as well as optical microscopic observation in transmitted light. Acknowledgements: We acknowledge CNES for funding. References: [1] MacKenzie W.S. and Guilford C. (1980) Atlas of rock-forming minerals in thin section, ed. Pearson Prentice Hall, pp. 98. [2] Michel-Lévy A. (1888), Interference color chart.

Fig. 1 Pieces of rocks from (a) the Gunflint formation (-1,9 Ga) and (c) Svalbard, and (b,d) associated optical microscopy images in transmitted light. (d) Image in polarized-analyzed light.

LOAC-S, A LIGHT AEROSOLS COUNTER/SIZER FOR PLANETARY ATMOSPHERES. J.-B. Renard1, N. Verdier2, F. Poulet3, M/ Vincendon3, O. Mousis4, 1 LPC2E-CNRS (3A avenue de la recherche scientifique, 45071 Orléans cedex 2, France, [email protected]), 2 CNES (18 avenue Edouard Belin, 31400 Toulouse, France, [email protected]), 3 IAS (Rue Jean-Dominique Cassini, 91440 Bures-sur-Yvette, France, fran- [email protected], [email protected]), 4 LAM (rue Frédéric Joliot Curie, 13013 Marseille, France, [email protected]).

Liquid and solid aerosols are present in the at- project submitted to the European Space Agency in the mosphere of many solar system objects: terrestrial frame of the M5 cosmic Vision, to study the chemical planets (Venus, Mars), satellites (Titan, Encelade), gas and isotopic compositions as well as the dynamics of giants, inner cometary coma. Aerosols can form in situ Saturn’s atmosphere. LOAC-S will be dedicated to in the atmosphere or be related to surface/subsurface determine the size distribution and the nature of the materials. Measuring aerosols properties can thus pro- Saturnian aerosols. Given the versatility of the instru- vide major constraints about both atmospheric compo- ment, LOAC-S could be an excellent candidate of the sition/dynamic and surface erosional/climatic history. payload of any planetary space mission dedicated to While some aerosols properties, such as average grain bodies with atmosphere. size or main composition, can be estimated using remote measurements, their size distributions, which are related to their formation process and their origin, are often poorly known. Large uncertainties indeed remain in remote derivation of the size distribution of aerosols, which require more sophisticated in situ measurements. LOAC (Light Optical Aerosol Counter) is an ultra-light and compact optical counter/sizer to perform measurements of liquid and solid particles at ground and under all kinds of balloons in the Earth atmosphere (www.LOAC.fr). The particles cross a laser beam and their scattered light is recorded by photodiodes. In its The LOAC optical chamber current version, measurements are performed at two scattering angles. The first one is around 12°, and is almost insensitive to the refractive index of the parti- References: [1] Renard et al. (2016), Atmos. Meas. cles; the second one is around 60° and is strongly sen- Tech., 9, 1721-1742. [2] Moussis et al. (2016) PSS sitive to the refractive index of the particles. By com- 130, 80-103. bining measurement at the two angles, it is possible to retrieve the concentrations for 19 sizes between 0.2 and 100 micrometers and to estimate the main typology of the particles (droplets, carbonaceous particles, minerals particles, salts, ices), based on calibration charts obtained in the laboratory with real particles [1]. A new version of LAOC, called LOAC-S (S for “Spatial”) in under development in the frame of a “Re- search and Development” project funded by the French Space Agency CNES. The aim is to first improve the existing version of LOAC in terms of performance for the detection of sub-micronic aerosols and secondly to be able to use the instrument in space conditions: low temperatures, low and high pressures, electro-magnetic radiations. Also, the possibility of adding several angles of measurements is foreseen, so as to obtain the whole scattering function of the aerosols and therefore their typology. LOAC-S is involved in the HERA project [2], which is an entry probe dedicated to the study of the Saturn atmosphere (http://hera.lam.fr/). HERA is a SUSTAINABILITY FOR SPACE EXPLORATION – A CASE STUDY FOR COMMERCIAL EXPLORATION OF THE MOON. B. Pigneur1 et al. 1Centre for Systems Engineering, Mullard Space Science Laboratory, University College London.

Introduction: In the area of ‘New Space’, ‘Space name a few. Strategic knowledge gaps have been iden- 4.0’ and the growth of commercial use of outer space, tified and technology roadmaps have been produced to the meaning of sustainability is not limited to techno- address the technical challenges of human exploration logical sustainability anymore. This study aims to de- of the Moon [12], [13], [14], [15]. Several mission rive a common definition of sustainability for space architectures have been proposed or are being devel- exploration, looking at appropriate dimensions as well oped including commercial application. Last summer, as specific challenges linked to the space sector. An the 1st Silicon Valley workshop on lunar commerciali- attempt to develop a framework for future space mis- sation took place to discuss the future of lunar exploi- sion will be presented in this paper. Finally, this con- tation and Moon Village [16]. cept will be applied in the context of Moon explora- This paper presents the application of the findings tion. The authors will present the current findings of of the study onto a potential permanent lunar explora- this ongoing study. tion architecture. It assesses the sustainability of the Sustainability: The concept and definition of sus- proposed mission concept and makes recommenda- tainability is not new and has been used in many sections for future studies. tors including the space industry. From the authors References: point of view, it seems that there is an ongoing evolu- [1] UN Office for Outer Space Affairs (2017) tion of the definition for the space sector and therefore Guidelines for the long-term sustainability of outer an opportunity for a better understanding of the impli- space activities. cations for space exploration missions. [2] Johnson C. D. (2017) Handbook for New Actors Definition. This paper reviews definitions of sus- in Space, Secure World Foundation. tainability across space agencies and industries to es- [3] NASA’s Sustainability Portal tablish a common understanding [1], [2], [3]. It also www.nasa.gov/agency/sustainability learns from non-space sectors with a strong emphasis [4] Lombardi P. L. (1999) Understanding Sustain- on sustainability such as construction, green energy ability In The Built Environment: A Framework For and development projects [4], [5], [6]. Evaluation In Urban Planning And Design, PhD the- Dimensions. This study builds on prior research to sis. select and define appropriate dimensions and aspects [5] European Commission (2013) Horizon 2020 – of sustainability for space exploration. The four do- The EU’s new research and innovation Programme. mains of sustainability (ecology, economic, politics [6] Costanza R. and Pattern B. C. (1995) Defining and culture) discussed in the Circles of Sustainability and Predicting Sustainability, Ecological Economics. method [7] are studied through the lenses of the three [7] Paul J. (2015), Urban Sustainability in Theory pillars of economy, society and environment [8]. The and Practice: Circles of Sustainability, Routledge. seven modalities framework [9] is also reviewed as [8] Adams W. M. (2006), The Future of Sustaina- well as the circular economy approach and its ‘resolve’ bility: Re-Thinking Environment and Development in framework [10]. the Twenty-first Century, IUCN Report Challenges. Specific challenges linked to space [9] Thomas S. A. (2016) The Nature of Sustainabil- exploration are highlighted especially with regards to ity, Chapbook Press. technology exploration roadmap [11]. [10] The Ellen MacArthur Foundation (2015) To- Framework. The authors discuss a sustainability wards a circular Economy: Business rationale for an framework for assessment of future space exploration accelerated transition. concepts. This framework is based on definition and [11] Simpson M. et al. (2016) Space for the 21st dimensions of sustainability as well as specific chal- Century: Discovery, Innovation, Sustainability, Aero- lenges associated to space exploration. space Technology Working Group Vol. 5, CreateSpace A Case Study – Commercial Exploration of the Independent Publishing Platform. Moon: A strong advocacy for a return to the Moon [12] Shearer C. K. and Neal C. (2012) Strategic has been made by many people, centres, agencies and Knowledge Gaps for the “Moon First” Human Explo- organisations such as the Lunar Exploration Analysis ration Scenario, LEAG, SAT. Group, the International Space Exploration Coordina- [13] Hufenbach B. et al. (2015) International Mis- tion Group or the Lunar and Planetary Institute to sions to Lunar Vicinity and Surface: Near-Term Mis- sion Scenario of the Global Space Exploration Roadmap, IAC 2015. [14] Pieters C. M. et al. (2007) Summary and High- lights of the NRC 2007 Report: The Scientific Context for the Exploration of the Moon (SCEM) [15] Neal C. et al. (2009) The Lunar Exploration Roadmap - Exploring the Moon in the 21st Century: Themes, Goals, Objectives, Investigations, and Priori- ties. [16] Kapoglou A. et al. (2016) Outcome of the 1st Silicon Valley workshop on lunar commercialization to support a permanent human settlement on the surface of the Moon, Moon village stakeholder engagement working group.

MARVIN LANDING SITE CHARACTERIZATION L. Sander, R. Buchwald Airbus Defence and Space, Airbus-Allee 1,28199 Bremen, Germany, (mail to: [email protected])

Reference Mission: The performed landing site char- LRO LOLA data). Thus, in a second step, the data on acterization analysis has been established in the frame the relevant scale has to be extrapolated. of the Moon Advanced Resource Utilization Viability When selecting the adequate data set for this analysis, 1 Investigation (MARVIN) phase 0 study . MARVIN is the real data density included in the DTMs has to be aiming at landing on the near site of the moon to per- evaluated. The real data density is limited by the reso- form an ISRU demonstration of extracting oxygen lution of the camera and the applied process for calcu- from lunar regolith. lating local elevation data using double image photo- Motivation: Before selecting an adequate landing site, grammetry. As described by Barker et.al.[1] the real landing hazards and scientific objectives of the mission data density of the SELENE maps is most likely lim- have to be balanced. Besides illumination and commu- ited to 60-80m/px - independent from the available nication constraints, landing hazards to be considered spatial resolution of the DTM, leading to the decision are in particular craters, rocks and boulders and the to limit the spatial resolution of the SLDEM2015 also related effective slopes. Each of these hazards can lead to 60m/px. In order to avoid filtering effects due to the to a failure of the spacecraft at landing and the loss of applied down-sampling, the original SELENE data set the mission. The landing site characterization provides has been used for this analysis. both - the requirements for spacecraft sizing and rec- Figure 2 shows the derived slope frequency distribu- ommendations for landing site selection. tions at different base length (black line mean slope, Generation of slope maps: The first step of the land- red lines 1σ, 2σ and 3σ probability). ing site characterization is the generation of a slope map of the landing site based on available digital terrain models (DTMs). The DTMs are derived from remote sensing data of former missions (e.g. SELENE, LRO). Figure 1 shows an exemplary slope map of a landing area close to the Apollo 17 landing site at Mare Serenitatis.

Fig. 2: Extrapolation of slope vs. base length Based on Barker a cut of base length of 60 m per pixel has been used for the extrapolation of the data at lower base length. The visible deviation between the data points at lower base length and the calculated trend line confirms Barkers observation that the real data density contained in the SELENE DTM allows deter-

Fig.1: Mare Serenitatis slope map from Kaguya data mining slopes down to min. 60 - 80m scale directly. Extrapolation of slopes at relevant scale: For the References: design of the landing system effective slopes at lander [1] M.K. Barker et.al., A new lunar digital elevation footprint scale (2-5m) are needed. Available near glob- model from the Lunar Orbiter Laser Altimeter and al DTMs feature resolutions of 100m/px (GLD100 - SELENE Terrain Camera, Icarus, Vol. 273, p. 346- derived from LRO wide angle camera), 7.4m per pixel 355, 2016, doi: 10.1016/j.icarus.2015.07.039 (derived from SELENE terrain camera) or 60m/px [2] Carrier, W., et al.: Physical Properties of the Lunar (SLDEM2015 - combining SELENE WAC maps with Surface. LunarSourcebook - A User's Guide to the Moon. G. H. Heiken, D. T. Vaniman and B. M. French (Eds.), Cambridge University Press, pp. 475 - 594 1 Funded by the Federal Ministry for Economic Affairs (1991) and Energy on the basis of a decision by the German Bundestag. (Support Code: 50 JR 1603) DEVELOPMENT OF SHOCK ABSORPTION STRUCTURE FOR OMOTENASHI SURFACE PROBE. Tetsuya Yamada1 , Hideyuki Tanno2, and Tatsuaki Hashimoto3 : 1JAXA(3-1-1 Yoshinodai, Chuo-ku, Sagamihara, Kanagawa, JAPAN: [email protected])

Introduction: Recently, target bodies of the solar crucial to control the mechanical characteristics exploration have been extending to distant heavenly parametrically according to the mission requirement bodies beyond the Mars, and even to more distant (allowable maximum deceleration, instrument weight, ones. Because the overall mission duration tends to be etc) material properties are now under parametrical longer and longer in such high-energy missions, investigation. In addition to the shock absorption reliability requirements for the subsystems become characteristics, the radio wave transmitting severer. The compact probe for semi-hard impact characteristics are required especially for the landing is one of the key technologies to accomplish OMOTENASHI surface probe (SP). Thus, a certain such challenging exploration missions under small kind of urethane material is now the prime candidate weight budget restraints. One of the important issues for the material. concerned with the semi-hard landing compact probes In order to verify validity of the analysis and is secure functional reliability of the subsystem and extracting other issues concerned with the crushable impact shock absorption. The moon semi-hard landing material, the landing impact tests are being carried out mission OMOTENASHI suggested by JAXA has been by using the ballistic range in JAXA Kakuda Space selected for launch by NASA SLS in 2018. The present Center by impacting a small test model with onboard study shows state-of-the-art technology development measurement systems onto the simulated lunar surface of the crushable shock absorption structure together target as shown in Fig. 2. Measured impact data are with a designing of the small lunar semi-hard impact compared to the analytical prediction and discussed in surface probe of OMOTENASHI (Fig. 1). the present paper.

Crushable Structure and Impact Landing Tests: References: The crushable structure realizes not only chuteless [1] Hashimoto, T., Yamada, T., Kukuchi, J., et al., landing on the bodies with atmosphere but also on Nano Moon Lander: OMOTENASHI, submitted to airless ones by absorbing landing shock energy and ISTS, (2017). protects the inner instrument modules against the [2] Yamada, T., Tanno, H., and Ogasawara, T., et landing shock within a prescribed deceleration level. al., High-speed Compact Entry Capsule enhanced by These versatile landing is advantageous for secure Lightweight Ablator and Crushable Structure, Electric decent system in that ignition electronics, Transaction of ISTS, No.2015-e-25, (2015). parachute-triggering sensors, and triggering-control [3] Tanno, H., Komuro, T., Ito, K.:Free-flight logic circuits etc. are no longer required. Mechanical aerodynamic test with projectile-onboard recorder in a characteristics, mainly stress vs. strain characteristics, ballistic range, AIAA 2013-0475, 51st Aerospace have been investigated for various materials such as Science Meeting, Jan. 2013,Texas,(2013) porous metals, carbon structures and resins. It is

Fig. 1. Artist’s impression of semi-hard landing of the small Fig. 2. Impact landing test onto simulated-lunar soil Surface Probe for OMOTENASHI mission1. target. (30 m/s and 20° canted wrt the horizon. SCALE TESTING OF DYNAMIC BEHAVIOUR OF MOON LANDING E. Urgoiti1, J. De Miguel2, 1 SENER- Structures & Mechanisms Section, Av. Zugazarte 56, 48930 Las Arenas (Vizcaya) Spain Tel: 34944817500 Fax: 34944817603 email:[email protected], 2SENER-Structures & Mechanisms Section, Av. Zugazarte 56, 48930 Las Arenas (Vizcaya) Spain email: [email protected].

Introduction: defining the characteristics of the scaled model, the The use of a scale model for the determination of dy- scale laws and the tests cases equivalence to real land- namic behavior under non-Earth gravity conditions are ing cases on the Moon. It also include description of described in this paper. The base or the article are the the test campaign and images and videos of the differ- activities carried out under ESA contract for Landing ent cases performed with the outcome of the test cam- System Development in which part of the scope in- paign. cluded a scale model test campaign to validate simula- Test preparation and aspects on the test set up are de- tion models and provide qualitative assessment of the scribed to consider all those aspects found important dynamic behavior of a landing probe on the surface of for the representativity of the test. the Moon. The dimensioning of a scale model following scale Scale model testing provides a valuable tool to assess laws is shown for the Moon case and the presentation the dynamic behavior in different scenarios with gravi- of the extrapolation to a general case with other gravity ty conditions different than the one on Earth. This tool accelerations is also described is a cost effective approach to validate in first stages of the activity, which are the main parameters and modelling aspects to be considered for a reliable estimation of the dynamic behavior including turn over effects, rebounds and stability

The paper describe the activities performed in the frame of the Landing System Development project Investigations of the landing impact characteristics and the lander/lunar soil coupling interactions for a lunar lander

Yongbin Wang1, Shiqing Wu2, Xuyan Hou3, Pingping Xue4, Huan Liu5, Lei Wang6 and Shutong Chen7, 1Beijing Institude of Space Mechanics and Electricity (P.O.Box 5142-269, Beijing, China,100094, [email protected]), 2Beijing Institude of Space Mechanics and Electricity (P.O.Box 5142-269, Beijing, China,100094, [email protected]).

Introduction: With the background of China’s of the landing impact energy, and in this sense a half of manned lunar exploration project, in order to solve the the landing impact energy can be absorbed by the lunar landing impact security problems of manned lunar unconsolidated granule solid. Therefore, the study on lander, a full-scale prototype for the lunar landing the coupling response between the lander and lunar buffer of the manned lunar module has been designed, soil is very significant. developed and experimentally verified. Compared with the common finite element simula- Fig. 1 shows the landing gear structure, which are tion technology, the discrete element method (EDEM four-legged cantilever structures. software) combining theoretical analysis is adopt to

X analyze the landing characteristics of the lander, and herein the discrete element method is more appropriate for the study on the granular media as lunar soil. Based on the simulation model of lunar soil and the theory of the discrete element method, the theoretical O and discrete element models of interaction between the lander and the simulated lunar soil are presented re- 3 Z spectively. Through the comparison and analysis under 2 4 1 different working conditions, the influencing relations Y Fig. 1 lander structure and the position of stowed and between the lander and the lunar soil are obtained, and deployed mechanism of the the energy transmission and dissipa- In order to certify the integrated performance of tion between the two can be preliminarily revealed. We landing gear, landing impact system test was devel- believe this research may make contributions to oped. The test condition include different landing providing the experimental data and theory theoretical model, landing speed, landing topographical character- basis for the design of the security manned lu- istic. All results were effective and satisfy requirement, nar lander. the system structure had no destroy. Aluminum honeycomb stroked as design, it could bear impact load. The test certified this type of landing gear can satisfy each technical requirement. Fig. 2 shows the landing impact system test.

Fig. 2 landing impact system test The landing impact experiment shows that the simulated lunar soil contributes to about 50% absorptivity Near surface environment specifications for Lunar South Pole exploration sites F. Cipriani1 and Ilya Kuznetsov2, 1ESA/ESTEC, Keplerlaan 1, 2200AG, The Netherlands ([email protected]), 2 IKI, Российская академия наук, Profsoyuznaya ul., 84/32, Moskva, Rusland, 117342

Lunar exploration scenarii such as ROSCOSMOS Figure 1: context map (LRO/WAC) showing the Luna 25 to 29 landers planned between 2019 and 2030 area container (blue rectangle North of Boguslawsky include landing in the Lunar South Pole regions in or- crater). der to characterize trapped volatiles and usable re- A coupled PIC-Monte Carlo approach is used to sources, study the lunar dusty exosphere, plasma and calculate surface potential (e.g. Figure 2) and lunar radiation environments, and prepare for the future de- dust charging and near surface transport. ployement of robotic and humans explorers and humans infrastrucutres. In the years to come some steps will be undertaken to test technologies for landing, communication, and resources charachterization in a collaborative frame between ROSCOMOS and ESA [1,2]. Those activities will benefit from a better understanding of both the lunar regolith properties as well as close lunar surface environment characteristics at scales relevant to habitat, and surface operations. Here we present some predicted effects due to the Solar Wind Figure 2: surface potential map associated with a plasma and Solar EUV illumination interactions with DTM of a 130mx90m lunar surface patch (with re- the lunar surface such as regolith surface charging, duced resolution) at 60 illumination angle in the solar topology driven fields and potential surface dust mobi- wind. lization. This provides an illustration of possible study cases We explore a range of environmental parameters that can be performed for specific landers, rovers and corresponding to varying Moon location as well as payloads scenarios at the lunar surface. Solar activity while specific to the South Pole Region of the Moon at a possible area of interest North of the References: Boguslawsky crater (72.9°S, 43.3°E, ~100 km in diam- [1] Zakharov A. and Kuznetsov I. (2017) , Robotic eter), a region of interest for future landers such as Lu- Missions of the Russion Lunar Program DAP-2017. na Glob[3]. In particular we have used a LRO based [2] Carpenter J. et al. (2015) Lunar Exploration and DTM on a restricted (~100mx100m) 69.545°S, Science in ESA, Vol. 17, EGU2015-15783 43.544°E centered area of interest where some topo- [3] Ivanov M. A. (2015) PSS, Vol.117, 45–63 logical variations can be identified (see Figure 1).

MOON ADVANCED RESOURCE UTILIZATION VIABILITY INVESTIGATION STUDY. R. Buchwald1, S. Lentz1 , J. Bolz1 , T. Diedrich1 and J. Weppler2 1Airbus Defence and Space GmbH, Airbus-Allee 1, 28199 Bremen, Germany 2Deutsches Zentrum für Luft- und Raumfahrt e.V. (DLR), Königswinterer Str. 522-524, 53227 Bonn, Germany

mission concept, a lunar landing spacecraft and robotic Introduction: elements for the lunar regolith excavation and pro- This presentation presents the activities performed cessing within the ISRU feed and disposal system. in the frame of the Moon Advanced Resource Utiliza- The study was performed by Airbus Defence and tion Viability Investigation (MARVIN) phase 0 study*. Space On-Orbit Services and Exploration in close co- MARVIN was initiated by the DLR Space Administra- operation with the German aerospace centre DLR insti- tion, with the purpose to develop a lander mission tutes of Space Operations and Astronaut Training, Ro- which shall perform extraction and utilization of re- botics and Mechatronics Centre and the Institute of sources on the moon as a technology demonstration for Planetary Research and the Technical University of the usage within future exploration missions. Munich (TUM). In-Situ Resource Utilization (ISRU) is seen as a This presentation will provide status of the work key element to expand the capabilities for lunar explo- performed. It will briefly describe the results of the ration by breaking up the reliance on earth supplied trade-offs and assessments, outline the selected mis- resources for structures, power, propellant, life support sion elements and the follow-up roadmap up to the etc and thereby to lower overall launch mass of lunar targeted launch in 2025. exploration missions and associated cost for transportation of infrastructure and consumables from earth. References: Within the field of possible ISRU applications, [1] Lewis, J. S. et al. (1993). Resource of Near MARVIN focuses on the utilization of consumables Earth Space. The University of Arizona Press and in particular on the extraction of oxygen from lu- [2] Clark, D. L., Keller, B. W., and Kirkland, J. A. nar regolith as a useful product for human exploration (2009). Field Test Results of the PILOT Hydrogen life support, as well as an element for propellant or fuel Reduction Reactor. AIAA Sp. Conf. Expo. cell production. [3] Lee, K. A. et al. (2013) The ROxygen Project: Several processes for oxygen production from lu- Outpost-Scale Lunar Oxygen Production System De- nar regolith have been proposed and discussed within velopment at Johnson Space Center.” J. Aerosp. Eng., the last 30 years [1] and some have been matured up to 26(1), 67–73. a Technology Readiness Level of analog site testing [2] [3]. The missing step still is the prove of oxygen production on the lunar surface under lunar environmental conditions. The main objective of MARVIN is to close this gap and to demonstrate repeatable oxygen production on the moon as a technology demonstration that can be scaled pending on the needs of future missions. In order to achieve this objective the study overall includes not only the design of an oxygen reactor, but also the

* Funded by the Federal Ministry for Economic Affairs and Energy on the basis of a decision by the German Bun- destag. (Support code 50 JR 1603) This report was prepared by order of the Space Admin- istration of the German Aerosapce Center (DLR). The scope of work was provided by the Space Administration of the German Aerospace Center (DLR). The Space Administration of the German Aerospace Center (DLR) did not influence the results of this report; the contractor alone bears the responis- bility.

PROSPECT: ESA’S PACKAGE FOR RESOURCE OBSERVATION AND IN SITU PROSPECTING FOR EXPLORATION, COMMERCIAL EXPLOITATION AND TRANSPORTATION. R. K. Fisackerly1, J. D. Carpenter1, the PROSPECT User Group and the PROSPECT Industrial Team, 1ESA ESTEC, Keplerlaan 1, 2201AZ, Noordwijk, The Netherlands ([email protected])

Introduction: The Package for Resource Observa- tion and in-Situ Prospecting for Exploration, Commer- Sample heating and chemical extraction: Samples cial exploitation and Transportation (PROSPECT) is in are sealed in ovens, derived with heritage from those development by ESA for application at the lunar sur- developed for EXOMARS [4], Rosetta and activities face as part of international lunar exploration missions performed through the German LUISE programme. in the coming decade, including the Russian Luna-27 Samples can then be heated to temperatures as high at mission planned for 2021. 1000°C. Heating in vacuum extracts ices and solar Establishing the utilization potential of resources wind implanted volatiles and pyrolyses some volatiles found in-situ on the Moon may be key to enabling sus- from minerals. Reacting gasses may also be introduced tainable exploration in the future. The purpose of to the ovens to extract additional chemistry of interest. PROSPECT is to support the identification of potential A number of techniques are under investigation, based resources, to assess the utilization potential of those on a combination of flight heritage and laboratory in- resources at a given location and to provide infor- vestigations. These include combustion with oxygen mation to help establish the broader distribution. [5], oxidation using fluorine [6] and reduction using PROSPECT will also perform investigations into re- hydrogen and methane [7]. source extraction methodologies that maybe applied at Gas compositional analysis: Evolved gasses can be larger scales in the future and provide data with im- analyzed using an ion trap mass spectrometer [5] for portant implications for fundamental scientific investi- masses up to around 200AMU. This gives a qualitative gations on the Moon. measure of the composition. Objectives: PROSPECT aims to assess the in-situ Gas chemical processing: Target gasses are pre- resource potential of lunar regolith at any given loca- pared for isotopic analysis through refinement or con- tion on the Moon. In order to achieve this PROSPECT version to other chemicals [5]. Such conversion can is required to: prepare chemicals which are better suited than the • Extract samples from depths of up to 2m. original compounds to analysis using a mass spectrom- • Extract water, oxygen and other chemicals of eter and can remove isobaric interferences. interest in the context of resources. Gas isotopic analysis: Isotopes of the elements of • Identify the chemical species extracted. interest are measured using a magnetic sector mass • Quantify the abundances of these species. spectrometer, along with measurements of reference • Characterize isotopes such that the origins standards [5]. Using this technique accurate analysis is and emplacement processes can be estab- achieved, allowing comparison with laboratory meas- lished. urements on Earth. In the lunar polar regions PROSPECT is able to Conclusions: PROSPECT is a package for the in- target water ice. At all locations on the Moon vestigation of lunar volatiles and other potential re- PROSPECT is able to extract solar wind implanted sources with possible applications for both exploration volatiles from the regolith through heating and aims to and fundamental science. The package builds on flight extract oxygen and other chemicals of interest as re- heritage and a unique set of capabilities, developed sources from minerals by a variety of techniques. over decades by a number of groups across Europe. System Functions: PROSPECT is targeting flights in the early 2020’s Drilling and sampling: PROSPECT includes a drill as part of the first phase of lunar resource characterisa- that is required to access the subsurface to depths of up tion [8] and will be available as part of international to 2m. Once at the required depth a sampling tool re- missions in this time frame. moves small (~3 cm3) samples, whilst preserving sam- References: [1] Magnani P. et al. (2010) Proceed- ple temperature. Samples must then be extracted and ing of i-SAIRAS. [2] Marchesi M. et al. (2001), Pro- handled whilst minimizing alteration of the samples. ceedings of the 9th European Space Mechanisms and The drill is derived from that being developed for Tribology Symposium, 91 - 96. [3] Coradini A. et al. EXOMARS [1] and the Rosetta drill [2]. Modifications (2011) EPSC-DPS EPSC-DPS Joint Meeting 2011 are considered to account for unique lunar mission abstract. [4] Schulte W. et al. (2010), Proceeding of i- requirements and material properties. SAIRAS. [5] Wright I.P. et al. (2012) Planetary and Space Science, 74, 1, 254-263. [6] Sebolt W. et al, (1993) in Resources of Near Earth Space, The Univer- sity of Arizona Press, 129. [7] Schwandt et al., (2012) Planetary and Space Science, 74, 1, 49-56. [8] Carpen- ter et al. (2016) Space Policy, 37, 2, 52-57. Acknowlegement: The Industrial team is led by Leonardo S.p.A, Italy with development of the chemical laboratory led by The Open University, UK.

ESTIMATING THE SUNLIGHT POWER AVAILABILITY FOR A PROBE OR ROVER ON THE SURFACE OF MARS G. Beaufils1a , C. Baur1b, F. Cipriani1c, F. Forget2 1ESA ESTEC-TEC (Technical Directorate) Keplerlaan 1 2200AG Noordwijk ZH, The Netherlands, 1a TEC-EPM (Power Management Section), 1bTEC-EPG (Solar Generator Section),1cTEC-EPG (Space Environment and Effects Section), 2Laboratoire de Météorologie Dynamique, University Paris 6, 75252 Paris, France

Introduction: To support to the definition and de- The approach is a Monte Carlo modelling of pho- velopment of ESA missions on the surface of Mars, ton travel into the Mars atmosphere. The key parame- the ESA Technical Engineering and Quality direc- ters in this approach are the scattering and extinction torate has developed, over the past dozen years, a se- coefficients associated with the dust and ice particles ries of models to predict the performance of a solar floating in the atmosphere. The presently active tool is array operating on the surface of Mars. The process derived from the work by Ockert-Bell and al. [1] for involves estimating, for a given set of atmospheric dust properties. The ice particle model is derived from (date and location) conditions: Warren [2] considering particles with 3μm effective 1- the sunlight intensity, spectrum and directivity, radius. Idealized reference conditions used for power 2- the local thermal environment and sizing scenarios of EXOMARS do not, however, con- 3- the solar cell performance in those conditions. sider ice clouds. For the environment modelling (sunlight and tem- Validation of the light spectrum on Mars surface perature), dedicated tools have been developed and and its directivity remains a concern. The performance validated, under ESA funding, by the Laboratoire de of triple-junction solar cells in the Mars environment is Méteorologie Dynamique (Paris) in parallel with the driven by the light in the 300 to 650nm range. Little Mars Climate Database [6]. suitable data have ever been published [3][4], nor ac- These environment tools are available to the indus- tually measured, that detail the light spectrum at the try and the Russian partners involved in the surface in this domain. Most direct data have been EXOMARS collaboration. derived from the analysis of the reflected light on the MER rovers coloured camera targets [5]. Need: The sunlight illumination on the Mars sur- The ESA tool is in fact indirectly validated against face, for a given location and time, is dependent of the orbital data of light reflected by the atmosphere [1]. actual atmospheric opacity condition of the day, referred through its “optical depth” conventionally Temperature Environment: The tool provides an measured at 670nm. The actual capability of a rover or estimation of the different thermal fluxes and the tem- probe solar power system on the Mars surface is fur- perature of the ground and the air (at 1m). ther affected by dust accumulation. The tool runs a simplified 1D model derived from The developed tools are the ESA reference to relate the General Circulation Model developed by the La- the meteorological hypothesis to practical inputs for boratoire de Méteorologie Dynamique for generating computing the power system performance capability. the early editions of the Mars Global Database. This covers the needs for: This model is extremely rapid, yet its outputs have a- defining reference scenarios against which the been successfully validated, for latitudes below ±50°, probe or rover power system shall be sized. to fit in a 5°C range against a wide sample of Path- b- supporting the landing site selection by relating finder or Mars Global Surveyor (TES) data. the available weather statistics to a (statistical) power system operation capability. Cell Performance: The present ESA internal ref- c- supporting, after actual landing, the early opera- erence is based on a detailed model of the AZUR tion planning using short term weather forecast, at Space 3G30 cell (EXOMARS baseline). This model least until accumulated telemetry data provide better synthesises the ESA accumulated technical knowledge inputs. and validation data on this cell in its various applica- Figure 1 presents a typical result for a reference tions, including low intensity low temperature (LILT). scenario analysis (1st day after landing). Figure 2 pre- This model definition is confidential but its results sents a typical result for landing site potential analysis. are used to check the models used by the different EXOMARS partners, which is important considering Sunlight Environment: The tool provides an es- that the mission operating conditions are far from the timation of the direct, scattered (and reflected) light well validated domain of commercial space applica- spectrum as a function of the incoming direction. tions. 6 Conclusions: Estimating the sunlight power availability for a probe or rover on the surface of Mars is mainly a matter of relating (complex) meteorological 5 (and dust deposit contamination) conditions to practical ambient light spectrum and direction, considering a

4 scarce availability of data for validation. The ESA reference status, used to support the EXOMARS project, has been presented.

3 Improving validation using new data and upgrad- ing tools if necessary represents a continuous maintenance effort. An update is under preparation that 2 should improve the tool fidelity in the most critical “blue” light domain, specifically for the share that scattered light actually represents. 1 References: [1] Ockert-Bell and al, Journal of Geophysical Re- 0 6 7 8 9 10 11 12 13 14 15 16 17 18 Time of Sol (1/24 Sol) search, Volume 102, Issue E4, p. 9039-905, 1997. Fig. 1 Predicted solar cell short circuit current (A) for a 1st [2] Warren, S. G., Optical constants of ice from the operation day at 5°N latitude, considering a landing at ultraviolet to the microwave, Applied Optics, Vol 23., Ls=324° (EXOMARS former 2018 landing hypothesis), and 1206–1225, 1984. EXOMARS reference scenario for optical depth and rover [3] G.A. Landis and al, The Solar Spectrum on the tilt (no dust deposit yet). Martian Surface and its effect on Photovoltaic Perfor- mances, IEEE 2006 39 Lat 25°N, OD NOM, Worst in ±10° [4] G.A. Landis and al, Photovoltaic Cell Opera- 38 Lat 25°N, OD NOM, Worst in ±10° Lat 15°N, OD NOM, Worst in ±10° th 37 Lat 15°N, OD NOM, Worst in ±10° tion on Mars, 19 European Photovoltaic Solar Energy Lat 5°N, OD NOM, Worst in ±10° 36 Lat 5°N, OD NOM, Worst in ±10° Conference, Paris France, June 7-11 2004, 35 Lat 5°S, OD NOM, Worst in ±10° Lat 5°S, OD NOM, Worst in ±10° [5] P.M. Stella and al, Design and Performance of 34

33 the MER (MER (Mars Exploration Rovers) Solar Ar- 32 rays, Conference Record of the Thirty-first IEEE Pho- 31

30 tovoltaic Specialists Conference, p. 626 - 630, 3-7 Jan.

29 2005. 28 [6] Millour E, Forget F, Spiga A, Navarro T, Mad- 27

26 eleine JB, Montabone L, Pottier A, Lefevre F, Mont- 25 messin F, Chaufray JY, Lopez-Valverde MA, Gonza- 24 lez-Galindo F, Lewis SR, Read PL, Huot JP, Desjean 23

22 MC, MCD/GCM development Team (2015) The Mars 21 Climate Database (MCD version 5.2). European Plane- 20 tary Science Congress 2015, held 27 September – 2 19

18 October, 2015 in Nantes, France 10:EPSC2015-438 17 16 0 10 20 30 40 50 60 70 80 90 100 110 120 130 140 150 160 170 180 190 200 210 220 Sol in mission(sol) Fig. 2 Predicted solar cell daily integrated short circuit current (Ah) over 220 sols, for 4 landing latitudes, considering a landing at Ls=324° (EXOMARS former 2018 landing hypothesis), and EXOMARS reference profile for optical depth, dust deposit and rover tilt. AN EUROPEAN CURATION OF ASTROMATERIALS RETURNED FROM EXPLORATION OF SPACE: THE EURO-CARES PROJECT. C. Smith1, S. Russell1, F. Foucher2 and the EURO-CARES team, 1Natural Histo- ry Museum, London, UK, 2CNRS-Centre de Biophysique Moléculaire, Orléans, France ([email protected]).

Introduction: The objective of the H2020-funded - Analogue Samples to determine which analogue EURO-CARES project is to create a roadmap for the proxies are necessary in a curatorial facility for testing implementation of a European Extra-terrestrial Sample sample handling, storage and preparation techniques; Curation Facility (ESCF) that would be suitable for the - Portable Receiving Technologies to propose methods curation of samples from all possible return missions for the recovery and transport of samples from the likely over the next few decades, to the Moon, aster- landing site to the permanent curatorial facility. oids and Mars. Public outreach: Along with the scientific and technical requirements, the EURO-CARES project is also focussed on a high impact public engagement plan that engages children, university students, the general public and policy makers, as well as our academic and industrial peers. A significant risk to the development of an ESCF is the public perception of extra-terrestrial samples, potentially containing biological entities, being deliberately returned to Earth without going through the “sterilising” process of exposure to cosmic-rays and space environment. This could be of great concern to many people and could lead to major delays in the establishment of an ESCF. Hence, open communication is of great importance. www.euro-cares.eu The planning of the facility design needs to start as early as possible (i.e., several years before the Study and long-term curation of extra-terrestrial planned return sample date), ideally to finish the con- samples imply keeping the samples as clean as possi- struction and interior design of the building at least one ble from any possible contaminants, while ensuring or two years before any sample return, to have enough they remain contained in case of biohazards, in particu- time to properly test the facility on analogue samples lar when considering a Mars sample return mission. and to train a dedicated team. Such a facility will have The requirements for a combined high containment to preserve (and protect) samples for decades of re- and ultraclean facility will naturally lead to the devel- search to be carried out on them, so its lifespan must opment of a highly specialised and unique facility that be sufficient enough. will require the development of novel scientific and Ackowledgement: We thank the European H2020 engineering techniques. program for funding (grant number 640190). EURO-CARES team work is organized around five distinct technical Work Packages (WP), led by institutions and scientists and engineers from all over Europe. These cover aspects including: - Planetary Protection to devise an effective, legally compliant and realistic, programme while minimising risk to current scientific study and optimising access to researchers for future studies; - Facilities and Infrastructure to define the state of the art facilities required to receive, contain and curate extra-terrestrial samples and guarantee terrestrial planetary protection; - Instruments and Methods to determine which analyses should be performed within the ESCF while ensuring minimal contamination and minimal damage to the sample; Sustained Mars Exploration Through Mars Atmospheric and Surface Resource Utilization. Keir Gonyea1, Robert P. Mueller2, Brandon Sforzo3, Laurent Sibille4, Hisham Ali5, and Robert D. Braun6, [email protected], [email protected], [email protected], [email protected], [email protected], [email protected].

Introduction: This study proposes a novel mission architecture to establish routine, Earth-independent transfer of large mass payloads between the Mars surface and orbit by leveraging both the Mars atmosphere and surface for use in in-situ resource utilization. The first stage of routine mission operations involves an atmospheric resource mining aerobreaking campaign following capture into a highly elliptical Mars orbit. During each pass through the atmosphere, the vehicle ingests the atmospheric oxidizer and stores it onboard, using solid oxide electrolysis to convert the primarily CO2 atmosphere into usable O2 for propellant. Power is made available through the use of magnetohydrodynamic energy generation, which converts the motion of the plasma in the shock later into usable electrical energy. Upon termination of the aerobreaking sequence, the descent vehicle detaches from the orbit stack, deor- bits, and executes the entry, descent, and landing sequence. Hypersonic deceleration is achieved via a deployable heatshield to lower the vehicle ballistic coefficient and supersonic and subsonic deceleration are achieved via retropropulsion. Ground operations involve resource mining of the Martian regolith to produce methane and oxygen propellant to be used for the subsequent ascent back to Mars orbit and the apoapsis raise maneuver to initialize the aerobreaking sequence in addition to providing fuel for propulsive descent. A preliminary, first-order feasibility and closure study was performed. Numerical models were developed to assess the performance of the aerobreaking atmospheric collection, CO2 to O2 conversion, magnetohydrodynamic energy generation, surface propellant production, and entry, descent, and landing capabilities. Closure was achieved for a human class mission with 81 orbital scooping passes, each atmospheric scoop varying in duration from 7.1 minutes to 5.3 minutes, to ingest approximately 34,939 kg total CO2. Due to the solid oxide electrolysis chemical conversion process and other efficiency losses, O2 product totaled 6,986 kg, which was sufficient for descent and landing retropropulsion propellant requirements. Surface propellant production rates were also found, which satisfied ascent and orbit maneuver propellant requirements, in addition to providing sufficient retropropul- sive descent fuel.

Mars 2020 Flight Computer Redundancy for EDL. Mallory Lefland1, 1Jet Propulsion Laboratory, California In- stitute of Technology, 4800 Oak Grove Drive, Pasadena, CA 91109

Abstract: The Mars 2020 Mission is primarily a build-to-print version of the Mars Science Laboratory (MSL) Mission, especially in the area of Entry, De- scent, and Landing (EDL) [1]. The time criticality of the EDL timeline makes it vulnerable to computer de- fects that could get in the way of a safe landing on Mars. This led the MSL team to create a safety net – a stripped down version of flight software known as Se- cond Chance (SECC) running on the backup computer that could take over in the event of a prime computer reset [2]. Additional capabilities existed in the flight software on the prime computer that could have pre- vented the computer from resetting, however, the MSL team decided to disengage this functionality in favor of using SECC. The Mars 2020 team has reevaluated the variety of methods available for utilizing redundant computers during EDL to select an appropriate strategy. This paper will review the avionics redundancy scheme inherited from MSL and the decisions that led to the configuration that flew on August 5, 2012, the methods in which redundant computer schemes can be utilized during EDL, the flight software enhancements that the Mars 2020 team has added to the system, and the configuration approach that the Mars 2020 Mission is planning for both flight computers during EDL. References: [1] Allen Chen. Et al (2014) Mars 2020 Entry, De- scent, and Landing System Overview, IPPW11 Presen- tation #8015. [2] Chris Roumeliotis et al. (2013) The Unparalled systems engineering of MSL’s backup entry, descent, and landing system: second chance, 8th Annual System of Systems Engineering Conference, 2013, 13-1045.

On the Development and Design for an Affordable Two-Wheeled Mars Rover, MARVIN: Eric Huynh1, Myles Oldroyd2, KoKo San3, Julio Serafin4, Richard Sok5 Dr. Periklis Papadopoulos6, Professor Marcus Murbach7 San Jose State University, 1 Washington Square, San Jose, CA 95112

The purpose of the MARVIN program is to design a cost effective rover for future interplanetary exploration missions. MARVIN is a two wheeled, self-stabilizing, all terrain, rover designed to operate on Mars, analyze surface samples and perform atmospheric and science measurements using nano sensor technology. MARVIN includes self-contained deployable wheels and a drilling system for sample analysis. The system is 3 feet long and it carries the payload rover package to a small Mars probe network mission. A sixteen spoke wheel architecture is designed to fold inside MARVIN’s main spacecraft structure. The wheels are autonomously deployed following Mars landing. A deployable robotic arm equipped with drilling and soil sampling technology enables the system to collect and analyze soil sample up to one meter below the surface. The poster presented at the IPPW-14 workshop will showcase MARVIN’s system and subsystem level architectures. Autonomous Mars Robotic Agro-Life Support System. R. Rosila1, S. Torres2, L. Murray3, J. Untalan4, L. Freitas5, J. Carpio6, J. Esquivel7, C. Espinoza8, E. Lopez9, R. Blanco10, T. Soares11, E. Marchen12, I. Quintero13, Z. Hughes14, B. Tesfamichael15, A. Gonzales16, E. Reyes17, 1San Jose State University, San Jose CA 95192 rob- [email protected], 2-17San Jose State University, San Jose CA 95192

Establishing a colony on Mars would mark a pivot- The mechatronic structure of the Farmbot mounted on al point in human exploration. To achieve such a feat, a bed of soil [1]. an understanding of food production on Mars is a necessity. An ideal approach would be to send a robot to References: Mars, such as the FarmBot [1], prior to the arrival of [1]"FARMBOT GENESIS." FarmBot. N.p, n.d. humans and begin food production ahead of time. Web. 07 Mar. 2017. FarmBot is an autonomous robotic system that handles the cultivation of fruits and vegetables. This autonomous system will be able to plant, water, and mend crops creating a sustainable food resource for future explorers. A desired launch window for FarmBot shall be in 2018, on board a Space X launch vehicle. Due to the martian atmosphere, implementation of the FarmBot will require innovation of the following: power management system, autonomy, and environmental protection. To allow plants to grow prior to arrival with minimal human interaction, the development of a greenhouse or another habitable environment is required. To assure the plants have sufficient oxygen to begin growth, oxygen must be pumped into the artificial environment prior to starting cultivation.Constructing and maintaining greenhouses with a low-pressure interior is optimal, but cultivating plants in these conditions poses some biological obstacles. Fortunately, headway has been made in genetic modification that can utilize the benefits of low-pressure environments for plants. Pow- er for the Mars FarmBot and the devices, such as heating lamps, required to operate the greenhouse/habitat will rely on solar panels. The work to be presented at the conference will show the system architecture including its key subsystems: main thermal-structure, controls, power, communication, etc. Functional operational requirements have been developed and will be presented as well. Detailed mission operations will also be presented and a preliminary design will be showcased.

BIOCONTAINMENT OF MARS SAMPLES RETURNING TO EARTH A.Fumagalli1, B. Spagnoli2, A. Terribile3, D. Indrigo3, J. Romstedt4, S. Vjendran4, G. Kminek4 1 LEONARDO ([email protected]), 2 POSITECH C/O LEONARDO ([email protected]; [email protected]), 3 ENIPROGETTI ([email protected]; [email protected]), 4 ESA/ESTEC ([email protected]; [email protected]; [email protected]).

In the framework of a potential international Mars Bio-container will be constantly monitored for leaks to Sample Return (MSR) mission, a Bio-Containment allow for a final decision on Earth re-entry. system was conceived, designed, breaboarded and During the workshop, the architectural design of deeply investigated by Leonardo S.p.A. (formerly the Biocontainment system will be presented. Exten- Selex ES) and Eniprogetti (formerly Tecnomare) under sive presentation of the most recent activities focused European Space Agency development contracts funded on investigating and testing of possible technological by the Mars Robotic Exploraiton Preparation (MREP) solutions to implement the design baseline will also be Programme. This abstract presents the results achieved provided. In particular, results on investigation of the in the development phase, started in 2011, and reports sealing issue based on gaskets of different types will be the results of several tests performed on the bread- discussed. In this respect, four different polymeric gas- boards. A road map for the full development of the kets have been tested through a test campaign aiming system is also presented. at subjecting the polymers to an ageing process/performance degradation due to environmental MSR foresees to collect and transport a set of soil conditions related to the mission scenario (i.e. radia- samples from the Martian surface back to the Earth. tion, vacuum, thermal ageing) and to finally test the Once collected, each sample will be stored into a dedi- two best candidates performing an ‘operative’ test with cated canister (Orbiting Sample, OS), which will be the gaskets individually installed in a breadboard of the launched into Mars orbit and captured by a Mars- container and of its lid. The test campaign led to the Orbiting spacecraft. This spacecraft will be equipped selection of two different polymeric gaskets, that with a robotic system capable of sealing the Orbiting proved to be promising both in terms of resistance to Sample into a Bio-Containment system. The Biocon- representative environmental conditions and in terms tainment system is then returned to Earth, transported of sealing performance.Concerning the third seal, of by a dedicated Earth Return Vehicle. The final stage of metal type, a thorough experimental investigation has the return is the release of the Earth Return Capsule, finally led to the selection of a configuration based on carrying the biocontainment system, into the Earth’s a metal seal with axial compression. This solution has atmosphere for an entry, descent and hard landing on proved to provide the required level of sealing compat- the surface. The objective of the Biocontainment sys- ibly with a reasonable compression force. tem is to meet the backward Planetary Protection requirements which are to avoid any particle larger than The technology for the sterilization to break the 10nm to escape from the cointainer itself, thus preserv- chain with Mars has also been investigated and tested. ing the Earth’s Biosphere from contamination of parti- The sterilization is achieved by bringing the surface of cles coming from Mars. The biocontainer concept con- the Biocontainer interface potentially contaminated by sists of a double vessel with three levels of sealing, the OS to a temperature of 500°C for at least 1 s. Re- sterilization capability and sealing failure detection. sults of these activities will also be presented at the The overall system, including handling elements, is workshop. capable to position the OS into the Bio-Container, to “break the chain with Mars” through a specific sterili- The original Bio-Container System concept has zation process, to provide a tight sealing of the OS for been validated and consolidated through a significant its safe return. The Biocontainer will be closed with experimental test campaign performed at breadboard two levels of gasket based sealings with each sealing level. The proposed next step is the development and level including three gaskets of different types. One of test of a more complete and representative Engineering the two containment compartments is pressurized in Model of the system, to turn a concept into a concrete order to enable a monitoring system to detect possible possibility. major failures by pressure measurements. During the whole return phase the pressurized compartment of the BLUNT BODY EDL SYSTEM PERFORMANCE IMPROVEMENTS THROUGH DIRECT FORCE CONTROL AND DEPLOYABLE TABS. A. M. Korzun1, R. W. Powell2, S. Dutta1, R. A. Lugo2, and A. M. Dwyer Cianciolo1, 1NASA Langley Research Center, M/S 489, Hampton, VA 23681, 2Analytical Mechanics & As- sociates, M/S 489, Hampton, VA 23681.

Introduction: Achieving precision landing, ac- Direct force control (DFC) guidance schemes take cessing lander sites at higher surface elevations, and advantage of the ability to independently adjust the lift increasing payload mass to the surface of Mars require and side forces separately to maintain a multi-degree- the ability to generate aerodynamic lift [1,2]. The two of-freedom reference trajectory over the large unex- Viking landers and Mars Science Laboratory (MSL) pected density variations typical of the Martian atmos- are the only vehicles flown to date at Mars that were phere. With DFC, each control parameter affects only designed to use lift to modify the entry dynamics. All one design metric. For example, controlling angle of three vehicles provided aerodynamic lift with an ax- attack controls the lift-to-drag ratio (L/D), and control- isymmetric blunt body through a radial offset of their ling side slip angle controls cross range. It may also be respective centers of gravity (CG). The Viking landers possible to modulate ballistic coefficient in real-time. achieved this radial CG offset through a combination Several approaches have been identified as potential of the stowed lander’s position within the aeroshell and mechanisms for DFC, including deployable flaps, a ballast mass [3,4]. MSL jettisoned two tungsten ballast moveable CG, shape morphing, etc. This analysis as- masses, totaling 153.5 kg, to achieve the required trim sumes deployable, moveable trim tabs in a configura- angle of attack during entry and then jettisoned another tion and scale similar to that shown in Figure 1 as the six masses, totaling 174 kg, prior to parachute deploy- mechanism for DFC. ment to re-trim the vehicle to a near-zero angle of attack. This concept of operation used a total ballast mass of 327.5 kg to create aerodynamic lift during the α hypersonic phase of flight [5]. For comparison, the mass of each Mars Exploration Rover was 174 kg. Constraints on the entry, descent, and landing (EDL) system may be loosened if similar or improved aerodynamic performance can be achieved with a less massive system. Recent mechanical and thermal de- β β sign to integrate an appropriately sized tab into the MSL entry vehicle indicated that the MSL trim performance could be matched using a deployable tab, with 1/10th of the mass required for the ballast-based system [6]. In addition to robotic scale missions, increased aerodynamic performance may extend the ap- α plicability of axisymmetric blunt body configurations to much larger payloads, such as those required for the Figure 1. Assumed trim tab configuration. Top human exploration of the surface of Mars. and bottom tabs control angle of attack (α); Side tabs NASA’s Entry, Descent, and Landing Architecture control sideslip angle (β). Study (EDLAS) provided an objective to consider alternative entry guidance approaches that minimize the risk of flow impingement on the payloads in the human Trim tabs are aerodynamic control surfaces that Mars exploration architectures [7]. This objective ne- permit an entry vehicle to achieve aerodynamic per- cessitates satisfying requirements on entry deceleration formance requirements without requiring an offset limits and maximum range error at landing, while min- center of gravity. These deployable surfaces provide imizing the required propellant, propulsion system the ability to modulate L/D or even ballistic coeffi- attitude control requirements, and angles of attack dur- cient, and a combination of tabs can provide increased ing entry. The design philosophy of this effort is to control authority in both crossrange and downrange. utilize aerodynamics during entry to minimize the per- The number of tabs, radial location, tab area, tab cant formance burden on as many other subsystems as pos- (deployment) angle, and tab aspect ratio are all param- sible, such as those required during the descent and eters that can be traded to achieve a desired perfor- landing phases of the mission. mance metric or range of performance metrics. [1] Lockwood M. K. et al. (2006) Journal of Spacecraft and Rockets, Vol. 43, No. 2, 258–269. [2] Korzun A. M. et al. (2013) AIAA 2013-2809. [3] Viking ’79 Rover Study Volume 1: Summary Report (1974) NASA CR-132417. [4] Edquist K. T. (2006) AIAA 2006-6137. [5] Schoenenberger, M. et al. (2009) AIAA 2009- 3914. [6] Ivanov, M. C. et al. (2011) NASA/TM-2011- 216988. Figure 2. Example of a mechanically deployed rig- [7] Dwyer Cianciolo A. M. and Powell R. W. id trim tab [6]. (2017) AAS 17-254.

Trim tabs are well-suited as a mechanism for DFC. Trim tabs are low-mass devices compared to ballast or CG movement devices and can alleviate packaging challenges or augment performance with a non-zero zCG offset. In the event of vehicle mass growth, maintaining required performance does not necessarily require a change to the size of the tab, since the performance can be regained by adjusting the tab cant angle. While deployable tabs have been of interest to the EDL community since the early 1960s, no such system has been fully characterized or flown on a blunt body. Development of a tab system will require testing and aerodynamic database development; trim tabs are not currently an “off-the-shelf” solution. Tabs are challenged by significant increases in localized heating compared to stagnation point heating, due to flow separation/reattachment and strength of the shock generated at the tab hinge point. Limited experience with analysis to predict aerothermal environments for the tabs, particularly for turbulent, potentially separated flow, remains a challenge in aerodynamic and aerothermal database development. This paper will discuss the development and performance of direct aerodynamic lift and side-force modulation through use of a predictor-corrector controller, with emphasis on sensitivity to ballistic coefficient. Both schemes are applied to cases of Mars entry from a 1 sol orbit (4.8 km/s entry velocity) with low and high ballistic coefficient vehicles. The reference vehicle is assumed to be a 70-deg sphere-cone with independently articulating flaps that serve as both lift and side force modulation devices. All cases transition to supersonic retropropulsion for the descent and landing phases. Performance metrics include landing footprint, fuel consumption, and engine performance (gim- baling/differential throttling, maximum throttle setting). Atmospheric parameters and aerodynamics coefficients are dispersed in this analysis.

References: THE MARS CLIMATE DATABASE (VERSION 5.3). E. Millour1, F. Forget1, A. Spiga1, M. Vals1, V. Zakha- rov1, T. Navarro1, L. Montabone1,2, F. Lefèvre3, F. Montmessin3, J.-Y. Chaufray3, M. A. López-Valverde4, F. Gon- zález-Galindo4, S. R. Lewis5, P. L. Read6, M.-C. Desjean7, F. Cipriani8 and the MCD development team, 1Laboratoire de Météorologie Dynamique (LMD), IPSL, UPMC, Paris, France, [email protected], 2Space Sci- ence Institute, Boulder, USA, 3Laboratoire Atmosphères, Milieux, Observations Spatiales (LATMOS), IPSL, Paris, France, 4Instituto de Astrofisica de Andalucia (IAA-CSIC), Granada, Spain, 5Department of Physical Sciences, The Open University, Milton Keynes, UK, 6Atmospheric, Oceanic and Planetary Physics (AOPP), Oxford, UK, 7Centre National D’Etudes Spatiales (CNES), Toulouse, France, 8European Space Research and Technology Center (ES- TEC), Noordwijk, The Netherlands.

Introduction: The Mars Climate Database (MCD) GCM from which the datasets are obtained includes is a database of meteorological fields derived from water cycle [4,5], chemistry [6], and ionosphere [7,8] General Circulation Model (GCM) numerical simula- models. The database extends up to and including the tions of the Martian atmosphere and validated using ther-mosphere [9,10] (~350km). Since the influence of available observational data. The MCD includes com- Ex-treme Ultra Violet (EUV) input from the sun is plementary post-processing schemes such as high spa- signifi-cant in the latter, 3 EUV scenarios (solar mini- tial resolution interpolation of environmental data and mum, average and maximum inputs) account for the means of reconstructing the variability thereof. impact of the various states of the solar cycle. An im- At the time of writing, we are building a new ver- provement in upcoming MCDv5.3, compared to previ- sion of the MCD, MCD version 5.3 which will be re- ous version, will be a better evaluation of these 3 EUV leased in May or June 2017. scenarios (based on the latest observations of that flux The GCM that is used to create the MCD data is over the last measured solar cycles). developed at Laboratoire de Météorologie Dynamique As the main driver of the Martian climate is the du CNRS (Paris, France) [1-3] in collaboration with dust loading of the atmosphere, the MCD provides LATMOS (Paris, France), the Open University (UK), climatologies over a series of dust scenarios : stand- the Oxford University (UK) and the Instituto de Astro- ard year (a.k.a. climatology) , cold (i.e: low dust), fisica de Andalucia (Spain) with support from the Eu- warm (i.e: dusty atmosphere) and dust storm, These ropean Space Agency (ESA) and the Centre National are derived from home-made, instrument-derived d'Etudes Spatiales (CNES). (TES, THEMIS, MCS, MERs), dust climatology of the The MCD is freely distributed and intended to be last 8 Martian years [11]. In addition, we also provide useful and used in the framework of engineering appli- additional “add-on” scenarios which focus on individ- cations as well as in the context of scientific studies ual Martian Years (MY 24 to 32) for users more inter- which require accurate knowledge of the state of the ested in specific climatologies than the MCD baseline Martian atmosphere. Over the years, various versions scenarios designed to bracket reality, as illustrated in of the MCD have been released and handed to more figure 1. than 350 teams around the world. Current applications include entry descent and landing (EDL) studies for future missions (Insight, ExoMars 2020), investigations of some specific Mar- tian issues (via coupling of the MCD with homemade codes), analysis of observations (Earth-based as well as with various instruments onboard Mars Express and Mars Reconnaissance Orbiter),... The MCD is freely available upon request (contact [email protected] or [email protected] ); a simplified web interface for quick browsing at MCD outputs is available on http://www-mars.lmd.jussieu.fr

Overview of MCD contents: The MCD provides mean values and statistics of the main meteorological Figure 1: Illustrative example of temperature pro- variables (atmospheric temperature, density, pressure files obtained with various MCD scenarios, compared and winds) as well as atmospheric composition (in- to the retrieved Opportunity entry profile (retrieved by cluding dust and water vapor and ice content), as the P. Withers). Provided outputs: The MCD provides users with: represent the exosphere where species evolve • Mean values and statistics of main with their own distinct scale heights. meteorological variables (atmospheric • An improved gravity waves schemes, which temperature, density, pressure and winds), as accounts for both horizontal and vertical well as surface pressure and temperature, propagation of these. CO2 ice cover, thermal and solar radiative fluxes, dust column opacity and mixing ratio, [H20] vapour and ice concentrations, along with concentrations of many species: [CO], [O2], [O], [N2], [Ar], [H2], [O3], [H] ..., as well as electrons mixing ratios. Column densities of these species are also given. • Dust mass mixing ratio, along with estimated dust effective radius and dust deposition rate on the surface are provided. • Physical processes in the Planetary Boundary Layer (PBL) [12], such as PBL height, minimum and maximum vertical convective winds in the PBL, surface wind stress and sensible heat flux. Figure 2: Temperature profiles at a given time and • A high resolution mode which combines high location extracted from MCDv5.2, illustrating of the resolution (32 pixel/degree) MOLA fact that the thermosphere, MCDv5.2 minimum, aver- topography records and Viking Lander 1 age and maximum EUV scenarios do not quite bracket pressure records with raw lower resolution the current solar cycle .The current redesigning of the GCM results to yield, within the restriction of minimum and average EUV scenarios in MCDv5.3 the procedure, high resolution values of will fix this issue. atmospheric variables. • The possibility to reconstruct realistic conditions by combining the provided Validation of MCDv5.3: As with previous ver- climatology with additional large scale sions of the MCD, this new version will be validated (derived from Empirical Orthogonal using available data from past and current missions: Functions extracted from the GCM runs) and e.g. Viking Landers, Curiosity, TES (onboard Mars small scale perturbations (gravity waves), a Global Surveyor), MCS (onboard Mars Reconnaisance scheme which has been improved for Orbiter), Mars Global Surveyor and Mars Express ra- MCDv5.3. dio occultations, etc. Results from this validation campaign will be presented at IPPW14. Improvements in MCDv5.3: This new version of the MCD is an upgrade from MCDv5.2 including the following additions and upgrades: References: [1] Forget F., et al. (1999) JGR, 104, E10. [2] Lewis S., et al. (1999) JGR, 104, E10. • Better documentation on guidelines for the in- th tegration of the MCD on Windows operating [3] Forget F., et al. (2014), 5 Int. Workshop on Mars systems. Atmosphere Modeling and Observations. [4] Made- • Improved interfaces with Matlab, Scilab and leine J.-B., et al. (2012) GRL, 39:23202. [5] Navarro T., et al. (2014) JGR (Planets). [6] Lefevre F, et al. IDL which will be fully interactive with the th MCD (Fortran) software (2011), 4 Int. Workshop on Mars Atmosphere Model- • Improved computation of some of the trace ing and Observations. [7] Gonzalez-Galindo F., et al. (2013) JGR (Planets), 118. [8] Chaufray J.-Y., et al. species column densities when in MCD “high th resolution” mode. (2014), 5 Int. Workshop on Mars Atmosphere Model- • Updated Extreme UV scenarios to better ing and Observations. [9] Gonzalez-Galindo F., et al. (2009) JGR, 114. [10] Gonzalez-Galindo F., et al. bracket the observed Sun cycle (see figure 2). th • Improved extrapolation around the MCD (2014) , 5 Int. Workshop on Mars Atmosphere Model- “model top” (~10-8 Pa, i.e. ~250km) to better ing and Observations. [11] Montabone L., et al. (2015) Icarus. [12] Colaitis A., et al. (2013) JGR (Planets). Status of the InSight Entry, Descent, and Landing System Eugene Bonfiglio 1 and Myron Grover 1, 1Jet Propul- sion Laboratory, California Institute of Technology (4800 Oak Grove Drive, Pasadena CA 91109).

Introduction: The Interior Exploration using Seismic Investigations, Geodesy and Heat Transport (InSight) is the 12th selection of NASA’s Discovery Class missions, set to launch in May 2018 and land on November 26, 2018. After landing, InSight will deploy two instruments on the surface of Mars, which will measure seismic activity and temperature gradients to investigate the fundamental processes of terrestrial-planet formation and evolution. The presentation will focus on the EDL portion of the InSight mission with highlights on other non-EDL aspects, such as mission design, instrument status, and Assembly Test and Launch Operations. The InSight spacecraft relies heavily on the Phoe- nix lander design and, thus, uses a nearly identical EDL design and architecture. Since the launch delay from the 2016 to the 2018 opportunity, a few key design elements have changed, including a novel change to the trigger used to determine when radar data can safely be ingested into the nav filter. This year’s presentation will include a discussion of radar ambi- guities the lander is susceptible to, which were a significant driver in the update to the trigger for ingesting radar data. It will also cover mission design and EDL interdependencies and subtleties that impact how the InSight navigation team was planning on targeting the Entry Interface Point (EIP). Other changes to the nominal EDL design will also be presented. MARS 2020 ENTRY, DESCENT, AND LANDING UPDATE. A. Chen. Jet Propulsion Laboratory, California Institute of Technology (Pasadena, CA, 91109, [email protected])

Abstract: Building upon the success of Curiosity’s landing and surface mission, the Mars 2020 project is a flagship-class science mission intended to address key questions about the potential for life on Mars and collect samples for possible return to Earth by a future mission [1]. Mars 2020 will also gather knowledge and demonstrate technologies that address key challenges for future human expeditions to Mars. Based on the highly successful entry, descent, and landing (EDL) architecture from the Mars Science Laboratory (MSL) mission [2], Mars 2020 will launch in July of 2020 and land on Mars in February of 2021.

The mission takes advantage of the favorable 2020 launch/arrival opportunity; this enables the delivery of a larger, heavier, and more capable rover to wider variety of potential landing sites. While Mars 2020 inherits most of its EDL architecture, software, and hardware from MSL, a number of changes have been made to correct deficiencies, improve performance, and increase the overall robustness of the system. The most significant of these changes is the recent addition to the baseline of a Terrain Relative Navigation (TRN) system, which will allow the vehicle to safely land at much more rugged and hazardous landing sites. Addi- tionally, Mars 2020 is engaging in significant parachute risk reduction activities and considering potential parachute design changes.

This paper presents an update on the development of the Mars 2020 EDL system. Additionally, the paper also summarizes the Mars 2020 landing site selection effort following the completion of the project’s 3rd landing site workshop.

References: [1] Mustard, J., et al. (2013) “Report of the Mars 2020 Science Definition Team,” Tech. rep., Mars Explora- tion Program Analysis Group (MEPAG). [2] Steltzner, A. (2013) “Mars Science Laboratory Entry, Descent, and Landing System Overview”, AAS 13-236.

MARS 2020 ON-BOARD TERRAIN RELATIVE NAVIGATION P. B. Brugarolas1, J. Casoliva1, A. Johnson1, S. Mohan1, A. Chen1, A. Stehura1, D. Way2 and S. Dutta2 1Jet Propulsion Laboratory, California Institute of Technology, 4800 Oak Grove Drive, Pasadena, CA, 91109 2NASA Langley Research Center, 8 Lindbergh Way, Hampton, VA, 23681

Abstract

The Mars 2020 project has baselined the addition of on-board Terrain Relative Navigation (TRN) to its Entry Descent and Landing (EDL) system. TRN enables access to a wider set of high scientific interest landing sites, which could have 100 m class landing hazards (e.g., large slopes, big boulders, scarfs, etc) within the landing ellipse. The top two candidate landing sites (Jezero Crater and North East Syrtis) require TRN to reduce the landing risks to a level that makes these landing sites feasible.

TRN adds two new elements to the EDL Guid- ance Navigation and Control (GNC) system, which is based on the Mars Science Laboratory EDL GNC system: a) a Lander Vision System (LVS) sensor, which provides map relative localization, and b) an autonomous on-board Safe-Target Selec- tion (STS) algorithm which identifies the safest landing target within the reachable region from an onboard safe targets map.

This paper provides an update to the STS preliminary design presented in [1]. It will discuss updates to the methodology and rationale to build the safe targets map, which takes into consideration terrain hazards, proximity between terrain hazards, and sensing and control errors. It will also discuss the updates to the safe target selection algorithm. Final- ly, this paper will discuss the key tuning and performance trades associated with adding TRN to the EDL GNC heritage system.

Acknowledgments: The research described in this article was carried out in part at the Jet Propul- sion Laboratory, California Institute of Technology, under a contract with the National Aeronautics and Space Administration.

Copyright (2017). California Institute of Technolo- gy

[1] Paul Brugarolas et al. (2015) On-Board Ter- rain Relative Safe-Target Selection for the Mars 2020 Mission, IPPW13. June 2016. Mars 2020 Parachute Modeling Updates D.W. Way1, S. Dutta2, C. Zumwalt3, A. Chen4, C. O’Farrell5, and I. Clark6. 1NASA Langley Research Center (Hampton, VA 23681, [email protected]), 2NASA Langley Re- search Center (Hampton, VA 23681, [email protected]), 3NASA Langley Research Center (Hampton, VA 23681, [email protected]), 4Jet Propulsion Laboratory (Pasadena, CA, 91009, [email protected]), 5Jet Propulsion Laboratory (Pasadena, CA, 91009, [email protected]), and 6Jet Propulsion Laboratory (Pasadena, CA, 91009, [email protected]).

Abstract: The Mars 2020 project is a flagship- Technology Conference, Datona Beach, FL, Mar. class science mission to land the next robotic explorer 2013. [5] J. R. Cruz, D. W. Way, J. D. Shidner, J. L. on Mars. With a state-of-the-art suite of scientific in- Davis, D. S. Adams, and D. M. Kipp, “Reconstruction struments, the new rover will conduct a search for the of the Mars Science Laboratory Parachute Perfor- evidence of past life, and for the first time, collect rock mance and Comparison to the Descent Simulation,” and soil samples for possible return to Earth. Based on Journal of Spacecraft and Rockets, Vol. 51, No. 4 the highly successful Mars Science Laboratory (MSL) (2014), pp. 1185–1196. [6] C. O’Farrell, E. J. Bran- Entry, Descent, and Landing (EDL) architecture and deau, C. Tanner, J. C. Gallon, S. Muppidi, and I. G. the Sky Crane landing system, the new mission will Clark, “Reconstructed Parachute System Performance launch in July of 2020 and will reach Mars in February During the Second LDSD Supersonic Flight Dynamics 2021. Test,” AIAA 2016-3242, AIAA Atmospheric Flight While the Mars 2020 EDL system inherits most of Dynamics Conference, Washington, DC, June 2016. its architecture, software, and hardware from MSL, a few minor adjustments and improvements have been made to either correct known issues or to improve the overall robustness of the system. Likewise, only minor adjustments have been made to the end-to-end EDL simulation and its subsystem models. One notable example is recent improvements in the simulation models of Disk-Gap-Band supersonic parachute. This paper will present several updates to the MSL parachute models, made to improve simulation predictions, based on new data from the MSL flight reconstruction and recent wind-tunnel experiments. While the MSL simulation performed very well with respect to predicting parachute performance for MSL, these changes are in-part motivated by the recent parachute failure observed on LDSD’s SFDT-2 flight test. These adjustments improve the accuracy of peak opening load and peak entry vehicle attitude rate predictions.

References: [1] Allen Chen et al. (2015) 2015 Update: Mars 2020 Entry, Descent, and Landing System Overview, IPPW12 Presentation #2104. [2] Way, D. W., Davis, J. L, and Shidner, J. D. (2013) “Assessment of the Mars Science Laboratory Entry, Descent, and Landing Simulation”, AAS 13-420. [3] Way, D. W. (2013) “Preliminary Assessment of the Mars Science Labora- tory Entry, Descent, and Landing Simulation”, IEEE- 2013-2755. [4] J. R. Cruz, D. W. Way, J. D. Shidner, J. L. Davis, R. W. Powell, D. M. Kipp, D. S. Adams, A. Sengupta, M. Kandis, and A. Witkowski, “Para- chute Models Used in the Mars Science Laboratory Entry, Descent, and Landing Simulation,” AIAA 2013- 1276, 26th AIAA Aerodynamic Declerator Systems TOP LANDING SITE CANDIDATES FOR THE MARS 2020 ROVER MISSION. E. K. Stilley1, A. Chen1. 1Jet Propulstion Laboratory, California Institute of Technology, 4800 Oak Grove Drive, Pasadena, CA 91109, [email protected].

Introduction: The Mars 2020 (M2020) mission is a high-heritage Mars mission based on the successful flight, Entry, Descent, and Landing (EDL), and surface operations of the Mars Science Laboratory (MSL). Many of the non-heritage changes come in the form of new or upgraded instruments chosen to help meet the ambitious mission objectives. The M2020 mission objectives include characterizing and sampling an as- trobiolocially-relevant ancient environment that is geo- logically diverse, and caching multiple samples of compelling rock and regolith for potential future return to Earth. The mission’s third Landing Site Workshop was held in February 2017. The scientists, engineers, and other community members participating were tasked with evaluating eight sites that had been scrutinized for landing safety, surface operability, and ability to meet the mission objectives, and recommending three or four sites to be carried forward for continued analysis. Based on engineering assessments performed by the M2020 team that were presented prior to the workshop, all eight landing sites proved to have relatively manageable safety from an EDL perspective. Only one site, Holden Crater, was graded to be less capable of meeting surface operability objectives compared to the others. Ultimately, the criteria used to evaluate which sites were chosen to move forward for future consideration by the project were judged primarily for their potential to meet the overarching scientific objectives. Three sites emerged from the three-day process (in alphabetical order): Colombia Hills, Jezero Crater, and Northeast Syrtis. Each of these sites presents unique exploration potential and challenges for the EDL system. This paper will present the approach and details of the EDL landing safety assessment as well as an overview of what makes these three sites relevant to the M2020 mission and to potential sample return.

References: [1] K. A. Farley, K. H. Williford (2017) 2020 Landing Site LSW#3 proceeding, Mars 2020 landing site down-select Feb 2017.pdf. [2] Allen Chen et al. (2016) 2016 Update: Mars 2020 Entry, Descent, and Landing Overview, IPPW13 Presentation.

COLORIMETRIC ANALYSIS TO HELP IDENTIFICATION OF DRILLED ROCK POWDERS ON MARS: THE CALIPHOTO METHOD. F. Foucher1, G. Guimbretière2, N. Bost1, A. Courtois3, E. Marceau3, P. Martin4, M. Bergougnioux5, M. Josset6, A. Souchon6, J.-L. Josset6, A. Verhaeghe6, N. Le Breton7 & F. Westall1. 1CNRS, CBM, Orléans, France, 2CEMHTI, CNRS, Orléans, France, 3Université d’Orléans, France, 4CNRS, LPC2E, Orléans, France, 5MAPMO, Université d’Orléans, France; 6Space-X Institute, Neuchâtel, Switzerland, 7ISTO, Uni- versité d’Orléans, France, ([email protected]).

Introduction: The objective of the ExoMars 2020 ments have shown that drill powders (fines) are char- mission (ESA-Roscosmos) will be to search for past or acterised by a more or less similar grain size (<60 µm). extant traces of life on the red planet. The originality of Secondly, the most important issue was the fact the mission is its drill which will permit production of that the colour of a powder is totally dependent on the centimetric drill-cores from up to 2 meters in depth. ambient luminosity. In order to solve the last issue, a During these drilling phases, the CLUPI camera will new method called CaliPhoto® was developed that observe the pile of rock powder forming at the surface. consists of adding a reference target to the field of The aim of the present study is to determine whether view of the camera, close to the powder. An image any geological information can be deduced from these processing algorithm is then used to calibrate the im- observations. On Earth there is no need to observe ages and permit comparison of the different powders. rocks in powder form and thus, to date, no investiga- Well characterized rocks were then crushed, sieved tion has been done to link the colour of a powder with and photographed using a commercial camera a particular type of rock. However, due to the particu- equipped with a detector similar to that used by the larity of the mission, colour information could become ExoMars Close Up Imager, CLUPI, i.e. a digital Fove- essential to improve the identification of the drilled on® sensor, in order to create a database of the colours rocks on Mars. of rock powders. Analogue sample selection: The majority of rocks The CaliPhoto® method: Several tests were car- on the surface of Mars are volcanic [1,2] thus, for this ried out. For rocks included in the database created so study, relevant samples were selected from the Massif far, the method permits identification with good accu- Central, in France, in order to cover a large range of racy. More interestingly, for those rocks not included volcanic rock types, as designated in the compositional in the database, it was still possible to make a good TAS diagram (Total Alkali Silica). The samples were match based on rocks of similar composition in the then crushed to less than 60 µm and displayed next to database. Moreover, rocks having similar elemental each other showing a large variety of colour (Fig. 1). composition but very different bulk colours, such as rhyolite (light) and obsidian (dark), have a similar powder colour, permitting thus a good identification. Finally, the CaliPhoto® method could be very useful on Mars to help identification of rocks during drilling without adding any new instrumentation. Acknowledgements: We acknowledge the Maison Fig. 1. Powders of various volcanic rocks from the du parc national des volcans d’Auvergne for permis- Massif Central, France. sion to sample. We thank CNES for funding. References: [1] McSween H.Y. et al. (2009) Main issues: Although the preliminary observation Science 324, 736. [2] Bost N. et al. (2013) Planetary shows differences in colour between the different sam- and Space Science 82-83, 113-127. ples, as shown in Figure 1, two main issues still remained to be tackled. First, we observed that the main difference between the different powders was the brightness; the different powders being more or less all included in tones of greys exhibiting darkish tones. Unfortunately, it also appears that decreasing grain size induces an increase in brightness. The powders must thus be compared after sieving, i.e. for similar grain size distribution. Nevertheless, this step remains relevant since experi- Mars Ascent Vehicle – Overview and Aeroheating/Thermal Protection System Design M. Lobbia1, A. Korzun2, J. Benito1, R. Shotwell1; 1Jet Propulsion Laboratory, California Institute of Technology, 4800 Oak Grove Dr., MS 301-490, Pasadena, CA 91109-8099; 2NASA Langley Research Center

Introduction: The most recent planetary science 2) a Mars orbiter containing on-orbit OS capture decadal survey indicated that one element of a poten- equipment and an Earth Entry Vehicle (EEV), and 3) a tial robotic Mars Sample Return (MSR) campaign MAV, OS, and a fetch rover (see Fig. 1). should be considered as the highest-priority large-class In the late 2016 timeframe, NASA drove towards mission [1]. While there is currently no specific plan or finalizing a Point of Departure (POD) design for a schedule for MSR, NASA is engaged in a variety of MAV implementing a single-stage hybrid propulsion related technology and risk reduction activities [2-3]. system using liquid injection thrust vector control Future Mars mission concepts are also being studied (LITVC) (see Fig. 2). The POD configuration, de- that could perform elements of the campaign, starting signed to enable launch from all candidate Mars 2020 with the sample caching function of the upcoming landing sites, would inject the OS into a 479 x 479 km Mars 2020 mission. Mars orbit. The notional Mars orbiter designed to ren- One key element of an MSR campaign would be a dezvous with the OS would employ optical tracking; Mars Ascent Vehicle (MAV) used to bring an Orbiting an active beacon integrated into the OS is also under Sample (OS) container from the Martian surface to consideration. Mars orbit. Mass reduction is a key focus of the The specific propellant combination being studied MAV/OS conceptual design efforts, as every kg of allows for storage temperatures as low as -72 deg. C, mass on the OS translates into many multiples of that aiding integration into the Mars environment. The required on the launch pad at Earth. While the Martian POD design is enclosed in a launch tube to provide atmosphere is much thinner than Earth’s, there is still thermal and environmental isolation for the MAV; sufficient density (coupled with relatively high veloci- current design trades are considering both a CO2 gas ties during ascent) that a thermal protection system gap or a gap filled with opacified aerogel insulation. (TPS) would be required to protect high-heating areas Aeroheating: Initial trades exploring the design on the MAV. Assuming the OS is attached directly to space for the MAV implemented lower-fidelity tech- the top of the MAV (as is the case in current design niques (e.g. Sutton-Graves approximation [4] for stag- iterations), the TPS design must consider both a low- nation convective heating) to determine the aeroheat- mass objective as well as sample loading techniques. ing environments and enable identification of relevant MAV System Overview: Various concepts for a TPS materials. Later iterations also included assess- MAV have undergone extensive design trades for sev- ments with the CBAERO tool [5], which can provide eral years. The current notional architecture is based on approximate estimates of heating on non-stagnation an MSR reference architecture consisting of three ele- regions using techniques such as Reynolds’ analogy to ments: 1) a sample-caching rover (such as Mars 2020), skin friction. To support TPS sizing for the POD de-

Fig. 2: MAV POD design employs a single-stage hy- Fig. 1: The notional MSR architecture is split into three primary brid propulsion system. sign, computational fluid dynamics (CFD) analysis was performed using the NASA LAURA code. This provided high-fidelity convective-heating estimates on the front of the MAV at discrete points along the MAV trajectory. TPS Considerations: Based on the relatively low convective heating environments for MAV ascent (as compared with planetary entry vehicle trajectories), a plethora of TPS materials can be considered to protect the OS during ascent. Both one-dimensional thermal/ablation analyses and three-dimensional thermal analyses have been conducted to verify transient thermal response to aeroheating environments. For the POD design, a Space Shuttle tile material- based TPS concept was baselined. This design was chosen based on a variety of considerations, including maintaining a robust outer surface to meet on-orbit capture/handling requirements, as well as using a composite substructure stood off from the OS to help with thermal isolation on-orbit. Subsequent to the POD design, it was determined that the relatively low heating can enable other concepts (such as metallic standoff structures based on Beryllium or other high- temperature alloys). Additional Discussion Planned for Paper: The full paper will provide additional background on the MAV system design, and the choices made to address the Martian environment. The paper will also discuss additional details of the aerothermal analysis, and TPS analysis results for a variety of material concepts. Fi- nally, some discussion of the design approach for a proposed upcoming MAV terrestrial demo will be presented.

References: [1] Squyres, S., et al., “Visions and Voyages for Planetary Science in the Decade 2013-2022,” National Academies Press (2011). [2] Perino, S., et al, “The Evolution of the Orbiting Sample Container for a Future Mars Sample Return,” IEEE Paper 2.0621 (2017). [3] Benito, J., et al, “Hybrid Propulsion Mars As- cent Vehicle Concept Flight Performance Analysis,” IEEE Paper 8.1304 (2017). [4] Sutton, K., and Graves, R. A., Jr., “A General Stagnation-Point Convective-Heating Equation for Arbitrary Gas Mixtures,” NASA TR R-376 (1971). [5] Kinney, D. J., “Aero-Thermodynamics for Conceptual Design,” AIAA Paper 2004-31 (2004).

PHOBOS SAMPLE RETURN: MISSION AND SPACECRAFT DESIGN. A. Wayman1, L. Peacocke1, M-C. Perkinson1, S. Vijendran2, T. Voirin2, J. Laranaga2 and D. Koschny2, 1Airbus, Gunnels Wood Road, Stevenage, SG1 2AS, United Kingdom, [email protected], 2European Space Agency, ESTEC, Keplerlaan 1, 2201 AZ Noordwijk, Netherlands, [email protected]

Abstract: Phobos Sample Return is a candidate The sampling is performed by a 4 degree of free- European Space Agency mission that would follow on dom robotic arm and rotary brush sampling mechanism from the ExoMars programme. The mission has been mounted on the LM. A verification system is included studied to Phase A level by European space industry, to ensure the appropriate mass and content of sample is and has been demonstrated to be technically and pro- obtained. The sample is then inserted into the ERC by grammatically feasible. Airbus has led one of the two the robotic arm, ready for the return to Earth. The cap- industrial contracts to investigate this pioneering mis- sule is covered in thermal protection material to sur- sion. vive the high-speed re-entry heat fluxes, and stiffened The Phobos Sample Return mission aims to travel by internal structures to survive the high g-loads at to the Martian system and characterize Phobos and impact. The sample itself will be protected by crusha- Deimos, before landing on the surface, obtaining a ble material to ensure it sees less than 2000g of accel- 100g sample of regolith and returning it safely to eration and is maintained in a state suitable for detailed Earth. Phobos is a prime target for Mars science due to scientific analysis. its spectral similarity to primitive D-type asteroids, its In the final phase of the study, a design-to-cost unknown origins, and its position as a witness plate to mission was developed. Several cost driving require- the formation and evolution of the Martian system. The ments were relaxed in agreement with the Science returned sample would enable analyses using full-scale Study Team, leading to the key science goals being laboratory instrumentation, significantly improving our satisfied at a significantly reduced. The design-to-cost understanding of these and other still unknown submission would no longer characterize Deimos and jects. would follow a simplified sampling site selection strat- The mission would launch in the 2024-2026 egy. This would significantly reduce the time spent on timeframe, on an Ariane 5 ECA or Ariane 6.4 launch the surface and allow the use of a simpler sample ac- vehicle from Kourou in French Guiana. A minimum of quisition system. A cold gas propulsion system is also one year is spent in the Martian system, with a return used to perform a powered descent down to the sur- to Earth in 2027-2032. The overall mission duration is face. 3-6 years depending on scenario and launch date. This presentation will present the results of the The composite spacecraft comprises four modules: Airbus Phobos Sample Return mission and spacecraft a chemical Propulsion Module (PM) that provides the design, with a focus on the key sample return techno- propulsion required to transfer the composite to the logioes as well as the cost and mass drivers of the mis- Martian system; a Lander Module (LM) to perform the sion. critical descent, using visual navigation, touchdown on landing legs and sample acquisition; an Earth Return Vehicle (ERV) to ascend from the surface and return the sample to Earth; and a passive Earth Re-entry Cap- sule (ERC) to protect the sample during atmospheric re-entry and hard landing at Woomera, Australia. A payload suite of cameras, spectrometers and a radio science experiment is carried on-board the LM to allow for detailed characterization of Phobos, and the selection of the preferred landing and sampling sites. The descent to the Phobos surface is performed using both a wide angle camera, with feature mapping and tracking software, and a radar altimeter. A final freefall from 50m will reduce the contamination of the surface by thruster plumes. The touchdown velocities are low, less than 1 m/s, and touchdown loads are attenuated using damping mechanisms integrated into the four landing legs. Altitude Control For High Altitude Balloons. V.D.Patel1 and P.E.Papadopoulos2, 1San Jose State University (One Washington Square San Jose, CA 95192, [email protected]), 2San Jose State University (One Washington Square San Jose, CA 95192, [email protected])

Introduction: [3] R.Z. Sagdeev, V.M. Linkin, J.E. Blamont, R.A. Venus has an extremely harsh environment, the Preston (1986). The VEGA Venus Balloon Experi- atmosphere’s chemical composition and temperature ment, Science, 231, 1407 makes it challenging to land a probe on the surface [4] A. Seiff, 'J.T. Scoffield, A.J. Kliore, F.W . Tay- without it melting or being crushed by intense pres- lor, S.S. Limaye, H.E. Revercomb, L.A. Sromovsky, sure. In the upper atmosphere of the planet, there is a V.V. Kerzhanovich, V.I. Morozand M.Ya. Marov layer that is very earth like in temperature and pres- (1985). Models of the Structure of the Atmosphere of sure. Past missions, VEGA 1 and VEGA 2, included Venus from the Surface to 100 km Altitude, balloons with Teflon like coating and were superpres- Adv.Space Res., 5, #I 1,3-58. sured but because of leaking they only about lasted 2 [5] B. Ragent, L.W. Esposito, M.G. Tomaslo, days each. JPL and NASA are working on balloons M.Ya. Marov, V.P. Shari and V.N. Lebedev (1985). using a combination of Mylar and Teflon and that will Particular matter in the Venus Atmosphere, Adv. be able to last almost 2 months. Superpressure balloons Space Res., 5 , #I 1, 8516 are designed to stay relatively stable in the upper at- [6] R.A. Preston, C.E. Hildebrand, G.H. Purcel1, J. mosphere but even then, there are 3-D winds that can Ellis, C.T. Stelzried, S.G. Finley, R.Z. Sagdeev, disturb the balloons and cause them to bob up and V.M. Linkin, V.V. Kerzhanovich, V.I. Altunin, down as the VEGA missions did. The problem with L.R. Kogan, V.I. Kostenko, L.I. Matveenko, S.V. these missions is that they’re all dependent on the wind Pogrebenko I, S. Strukov E.L. AlumY, u.N. Alexan- to move them around while having no real control over drovN, .A.Armand R.V.Bakitko, A.S.VysAh.lFo.vB, the direction. ogomYoulo. Nv,.Groshenkov, A.S.Selivanov, This paper will consider a design for an attitude N.M.Ivanov, V.F.Tikhonov, J.E.Blamont, L.Boloh, control system by implementing the use of venting, G.Laurans,A.Boischot,F.Biraud,A.Ortega-Malina, ballasting and a small He tank to increase stability and C.Rosolen, G.Petit (1986). Determinationof Venus mission lifespans of stratospheric balloons for Venus. Winds by Ground-Based Radio Tracking of the VEGA Since there is a significant time delay the altitude con- Balllons, Science, 231, 1414 trol system will have to be fully autonomous. For maintaining altitude two separate balloons will be used. One will be a superpressure balloon with a fixed pressure and the other a helium balloon with a vent to lower pressure when needed and increase pressure using the attached gas tank. The ballast will be attached to the payload and deployed when needed. With the combination of these three and on board pressure sensors, the balloon should be able to maintain a stable altitude.

References: [1] J. L. Hall, V. V. Kerzhanovich, A. H. Yavroui- an, G. A. Plett, M. Said, D. Fairbrother, C. Sandy, T. Frederickson, G. Sharpe, and S. Day, “Second generation prototype design and testing for a high altitude Venus balloon,” Advances in Space Research, Vol. 44, 2009, pp. 93-105 [2] V.V. Kerzhanovich, J.A. Cutts, A. Bachelder, J. Cameron, J. Hall, J. Patzold, M. Quadrelli, A. Ya- vrouian, J. Cantrel1, T. Lachenmeier, M. Smith (1999). Mars Balloon Validation Program. 13'h AIM Interna- tional Balloon Technology Conference, Norfolk, Vir- ginia, Collection of papers, 8. UNVEILING MERCURY’s MYSTERIES WITH BEPICOLOMBO – AN ESA/JAXA MISSION TO EXPLORE THE INNERMOST PLANET OF OUR SOLOAR SYSTEM. J. Benkhoff, ESA/ESTEC, SCI-S, Keplerlaan 1, 2200AG Noordwijk, Netherlands ([email protected])

Introduction: NASA’s MESSENGER [2] mission • to test general relativity with improved accu- has fundamentally changed our view of the innermost racy, taking advantage of the proximity of the Sun. planet. Mercury is in many ways a very different plan- Since and considering that the advance Mercury's peri- et from what we were expecting. Now BepiColombo helion was explained in terms of relativistic space-time [1] has to follow up on answering the fundamental curvature. questions that MESSENGER raised and go beyond. In addition, the BepiColombo mission will provide BepiColombo is a joint project between ESA and a rare opportunity to collect multi-point measurements the Japanese Aerospace Exploration Agency (JAXA). in a planetary environment. This will be particularly The Mission consists of two orbiters, the Mercury important at Mercury because of short temporal and Planetary Orbiter (MPO) and the Mercury Magneto- spatial scales in the Mercury’s environment. The fore- spheric Orbiter (MMO). The mission scenario foresees seen orbits of the MPO and MMO will allow close a launch of both spacecraft with an ARIANE V in Oc- encounters of the two spacecrafts throughout the mis- tober 2018 and an arrival at Mercury in 2025. From sion. Such intervals are very important for the inter- their dedicated orbits the two spacecraft will be study- calibration of similar instruments on the two space- ing the planet and its environment. craft. They also provide very scientifically valuable intervals to collect multi-point measurements in an Science goals: BepiColombo will study and under- environment where both spatial and temporal scales stand the composition, geophysics, atmosphere, mag- can be very short. netosphere and history of Mercury, the least explored planet in the inner Solar System. In particular, the mis- Instruments: The MPO scientific payload com- sion objectives are: prises eleven instruments/instrument packages; The • to understand why Mercury's density is mark- MMO comprises 5 instruments/instrument pachages to edly higher than that of all other terrestrial planets, the the study of the environment. Together, the scien- Moon included tific payload of both spacecraft will provide the de- • to understand and determine the status of the tailed information necessary to understand Mercury core of Mercury, and if the planet is still tectonically and its magnetospheric environment and to find clues active today to the origin and evolution of a planet close to its par- • to understand why such a small planet pos- ent star. The MPO will focus on a global characteriza- sesses an intrinsic magnetic field, while Venus, Mars tion of Mercury through the investigation of its interi- and the Moon do not have any, and investigate if Mer- or, surface, exosphere and magnetosphere. In addition, cury's magnetized environment is characterized by it will be testing Einstein’s theory of general relativity. features reminiscent of the aurorae, radiation belts and Major effort was put into optimizing the scientific re- magnetospheric sub-storms observed at Earth turn by defining the payload complement such that • to understand why spectroscopic observations individual measurements can be interrelated and com- not reveal the presence of any iron, while this element plement each other. The BepiColombo mission will is supposedly the major constituent of the planet complement and follow up the work of NASA’s • to investigate if the permanently shadowed MESSENGER mission by providing a highly accurate craters of the Polar Regions contain sulphur or water and comprehensive set of observations of Mercury. ice The mission has been named in honor of Giuseppe • to observe the yet unseen hemisphere of Mer- (Bepi) Colombo (1920–1984), who was a brilliant Ital- cury ian mathematician, who made many contributions to • to study the production mechanisms of the planetary research, celestial mechanics, including the exosphere and to understand the inter-action between development of new space flight concepts. planetary magnetic field and the solar wind in the absence of an ionosphere References: [1] Benkhoff, J., et al. (2010) Planet. • to obtain new clues about the composition of Space Sci. 58, 2-20. [2] McNutt R.L., S.C. Solomon, the primordial solar nebula and about the formation of R.E. Gold, J.C. Leary and the MESSENGER Team the solar system (2006) Adv. in Space Res. 38, 564-571.

BepiColombo – Engineering Challenges D. Stramaccioni & H. Ritter, European Space Agency, ESA/ESTEC, SCI-PBO, Keplerlaan 1, 2200AG Noordwijk, The Netherlands ([email protected])

Introduction: BepiColombo is Europe's first mis- been designed for temperatures up to about 500 °C sion to Mercury. It is planned to set off in 2018 on a whilst providing the demanding performance required journey to the smallest and least explored terrestrial by the Radio Science Experiment. A high temperature planet in our Solar System. BepiColombo is a joint Solar Array able to survive up to 220 °C operational mission between ESA and the Japan Aerospace Explo- limit. Most payloads have their dedicated thermal con- ration Agency (JAXA), executed under ESA leader- trol, based on low temperature sub-radiators, which ship. The mission comprises two spacecraft: the Mer- have to be highly decoupled from the internal space- cury Planetary Orbiter (MPO), provided by the Euro- craft environment. Pointing stability requirements of pean Space Agency, and the Mercury Magnetospheric the laser altimeter and the some optical instruments are Orbiter (MMO), provided by the Japanese Space satisfied mounting the instrument on dedicated Optical Agency. The transfer from Earth to Mercury will be Bench. performed by a dedicated module, the Mercury Transport Module (MTM), which makes use of a dedi- This talk will focus on the specific engineering cated solar electric propulsion system with additional challenges of the mission and will give an overview on chemical propulsion for the orbit insertion. A dedicat- the spacecraft design. A specific focus will be on the ed MMO Sunshield and Interface Structure (MOSIF), complex challenges faced by the thermal design of which provides thermal protection of the MMO space- MPO. craft, completed the BepiColombo spacecraft stack. Shortly before Mercury orbit insertion, the MTM is jettisoned from the spacecraft stack. The MPO then provides the MMO with the necessary resources and services until it is delivered into its mission orbit, when control is assumed by JAXA. The MPO orbit around Mercury will be 3-axis stabilized, planet oriented, with a planned lifetime of 1 year, and a possible 1- year extension. BepiColombo Spacecraft is facing a very harsh thermal environment in orbit around Mercury. The proximity to the Sun results in a solar irradiance ranging from 14,448 Wm2 (at Mercury perihelion) to 6,272 W (at Mercury aphelion). Additionally, Mercury rotates around its axis only three times in two Mercury years. In combination with the non-existent atmosphere these factors are responsible for surface temperatures ranging between 90 K and 725 K. Resulting from this, while passing over the planets subsolar point, the MPO will not only have to withstand the intense solar irradiation, but also needs to cope with a planetary IR load of up to 5000 W/m2. Design: The MPO thermal control design therefore comprises a radiator equipped with highly reflive fins to reflect Mercury IR environment while allowing heat rejection to the space. MPO is rotated by 180° twice per Mercury year in order to to keep the radiator face away from the Sun. A specially developed high temperature MLI able to withstand 450 °C covers the other five spacecraft surfaces. A complex network of embedded and surface heat pipes are used to remove the heat towards the radiator. The High Gain Antenna has The New Frontiers Venus In Situ Atmospheric and Geochemical Explorer (VISAGE) L.W. Esposito1, A. Allwood2, F. Altieri3, D. Atkinson2, S. Atreya4, K. Baines2,5, M. Bullock6, A. Colaprete7, M. Darrach2, J. Day2, M. Dyar8, B. Ehlmann9, K. Farley9, J. Filiberto10, D. Grinspoon8, J. Head11, J. Helbert12, S. Madzunkov2, G. Piccioni13, W. Possel1, M. Ravine14, A. Treiman15, Y. Yung9, K. Zahnle7 1LASP, Univ. Colorado, 2Jet Propulsion Laboratory, California Institute of Technology, 3Istituto Nazionale Di Astrofisica Inaf, 4Univ. Michigan, 5Univ. Wisconsin, 6Southwest Research Institute, 7NASA Ames Research Center, 8Planetary Science Institute, 9California Institute of Technology, 10Southern Illinois Univ., 11Brown Univ., 12Deutsches Zentrum Fuer Luft- Und Raumfahrt E.V., 13Istituto Nazionale Di Astrofisica Inaf, 14Malin Space Science Systems Inc., 15Universities Space Research Association The exploration of the inner solar system is driven by the overarching concept of comparative planetology - that understanding the structure, history, processes, and evolution of each inner solar system planet directly addresses the understanding of the other planets. The 2011 Planetary Science Decadal Survey [1] identified the Venus In Situ Explorer as the highest priority New Frontiers mission for future inner solar system studies. VISAGE, the Venus In Situ Atmospheric and Geochemical Explorer, addresses three fundamental goals: 1) to understand why Venus is so different from Earth by measuring noble gases to test models of the origin and evolution of Venus, and by measuring sulfur compounds and trace gas profiles to constrain atmospheric cycles, surface-atmosphere interactions, and climate models, 2) to understand whether Venus was ever like Earth by measuring surface and subsurface composition on the rolling plains of Venus, determining surface rock type, mineralogy, and texture to understand geochemical processes, weathering, and aeolian processes, and to provide ground truth for Magellan and VIRTIS, and 3) to understand what Venus can teach about extra-solar planets by modeling the history of Venus with an eye on predicting Venus’s future and to predict characteristics of Venus-like exoplanets and compare to signatures of exoplanets detected by Kepler and other observations.

The VISAGE Venus lander mission performs atmospheric and surface science investigations with a flyby carrier spacecraft that delivers a Lander and serves as the telecom relay. The VISAGE Lander science payload comprises five instruments: an Atmospheric Structure Investigation including Doppler Wind measurements, a Neutral Mass Spectrometer, an Imaging System, an X-ray Fluorescence experiment, and a Visible Near-Infrared Spectrometer. In the extreme surface environment of Venus, VISAGE is a relatively short (several hours) autonomous landed mission that requires no ground control. Once on the surface, VISAGE measures the mineralogy and elemental composition at two depths, with samples brought inside the Lander for analysis. Science investigations include measuring the inventory of noble gases and light stable isotopes, the abundance of trace and reactive gases from the surface to the clouds, and to provide descent imaging of the surface below 15 km. During descent, the thermal, compositional, and dynamical structure of the atmosphere along the probe trajectory is measured. Once on the surface, the elemental and mineralogical composition of surface rocks are measured, and panoramic images of the landing site are made. Measurements of noble gases in the atmosphere help discriminate between models of Venus’s origin, and the composition of the surface can elucidate the history of Venus.

VISAGE launches in December, 2024 with a targeted flyby of Venus in May, 2025 and Venus arrival in December, 2025. The VISAGE Lander descends under parachute and drag plate for one hour before landing at 8.8 m/s. Once on the surface, VISAGE conducts surface atmosphere measurements, surface and sub-surface composition measurements, and take panoramic images of the surface region for up to 3.6 hours. The total surface science data return is expected to be ~1.4 GBits.

Aerial Platforms for Venus Exploration: J. A. Cutts1, M.Pauken1, J. L. Hall1, K. H. Baines1, R. Grimm2 1Jet Propulsion Laboratory, California Institute of Technology, MS 321-550, 4800 Oak Grove Drive, Pasadena, CA 91109, [email protected], 2Southwest Research Institute, Boulder, CO

The dense atmosphere of Venus and the high tem- Several proposals have been made to apply this tech- peratures in the lower atmosphere and surface have pre- nology to a NASA or ESA mission. The VALOR Ve- sented impediments to the deployment of exploration nus Aerostatic-Lift Observatories for in-situ Research) techniques that work on airless bodies and planets with proposal was typical of these which focused on the at- thin atmospheres such as Mars. However, they also cre- mosphere [2]. The European Venus Explorer (EVE) ate opportunities for the use of aerial platforms to ex- conceived at about the same time, also focused primar- plore Venus in many different ways. NASA is currently ily on the atmosphere. Other more ambitious concepts conducting a study of the potential for of aerial plat- involved the deployment of sondes from the aerostat. forms as a contributed payload on the Russian Venera In this case, the aerostat serves as both a platform for precise deployment of the sondes and also as a com- D mission as well as for the primary science vehicle on munications relay. The proximity of the balloon to the a US Flagship mission. This paper reviews preliminary short lived sondes enables greater data return than results from the study. would have been possible for sondes communicating with an orbiting or flyby spacecraft., VEGA BALLOON MISSION The 2011 Planetary Science Decadal Survey recom- It is more than 30 years since the first and only aerial mended a Venus Climate Mission (VCM) as a small platforms were deployed at Venus, or indeed at any Flagship mission, comprising an aerostat, deep probe, planet, by the Soviet Union in 1985. Two VEGA aero- and two sondes. The objectives for the aerostat align stats implemented as 3.5m superpressure balloons were with previous atmospheric goals. The deep probe successfully deployed at Venus and were each tracked would be released during initial descent and provide from Earth as planned for about two earth days as they atmospheric and chemical data into the deep atmos- drifted halfway around the planet in the super-rotating phere, whereas the sondes could be released any time. atmospheric flow at an altitude of about 55 km. Alt- Recent work suggests valuable geoscience studies can hough the total payloads suspended beneath each aero- be performed from the aerostat itself. Infrasound sig- stat was only 6.9 kg, including sensors, batteries and natures of earthquakes can be detected in the atmos- communications equipment, VEGA remains an im- phere [3] and natural-source electromagnetic sounding portant proof of concept paving the way for more ambi- can probe the upper mantle [4]. Together, these tech- tious missions. niques can constrain the geodynamics of Venus without ever touching the surface. MISSION CONCEPTS Venus Geoscience Aerobot: More ambitious con- There has been no aerial platform mission to Venus cepts for the use of aerostats at Venus have also been since VEGA and currently none are under development. formulated. The Venus Geoscience Aerobot (VEGAS) However, in this period there have been several pro- concept [5] has a buoyant platform capable of making posals in both the US and Europe to fly more capable repeated short visits to the surface of Venus, and ex- aerostats at Venus. There have also been some im- tracting power from the thermal gradient in the atmos- portant technology developments and the option space phere in the process of conducting these maneuvers. for the use of aerial platforms at Venus has been exten- VEGAS would exploit the properties of water ammonia sively explored. mixtures for buoyancy and altitude control. VEGA-Type aerostats with larger payloads: One Aerostats and Sample Return: Venus Surface direction of research has been to develop an aerostat Sample Return (VSSR) has long been considered to be with a much larger payload capability than VEGA. JPL enabled by balloon loft of the ascent rocket [6]. An in- has been developing superpressure balloons tolerant of termediate mission concept – Venus In Situ Explorer both the sulfuric acid environment on Venus and capa- (VISE) advocated by the 2003 PSDS [7] was to perform ble of accommodating the diurnal stresses induced on sample analysis in a buoyant station at the clement bal- the balloon. A 5.5-m balloon with a payload capability loon-float altitude. A number of concepts for imple- of 45 kg is now at TRL 5 [1] and a 7.0-m balloon with menting VSSR and VISE have been considered includ- a payload of 110 kg is now under development. Demon- ing an innovative dual balloon concept. One spin off of strations have also been conducted of aerial inflation of this effort was a concept for a near surface balloon sys- superpressure balloons. tem called the Venus Mobile Explorer (VME) first identified in the NASA Solar System Roadmap of 2006 [8]. Altitude Control: In 2011, motivated by enduring This should also be a period for intensive technology questions about the nature of the mysterious time varia- investment in more capable systems that can make ex- ble ultraviolet haze in the Venus upper atmosphere, up- cursions in altitude both to 65 or 70 km near the top of per atmosphere, JPL began an investigation of ap- the cloud layer and downward to 40 km near the base of proaches to altitude cycling in the 55 to 70 km range.. the cloud layer. Other objectives would include systems Initially concepts using either ambient gas ballast capable of control in latitude including heavier than air (AGB) or Lift Gas Compression (LGC) were explored and hybrid technologies. There should be a focus on by the group at Smith College [9]. Subsequently, a con- systems capable of miniaturization enabling low cost cept for involving mechanical compressions by chang- missions with rapid turnaround ing the volume of the envelope was developed by Red Mid-Term 2025 to 2035: In this time frame, it Line Aerospace using their Ultra High Pressure Vessel should be possible to deploy aerial platforms with alti- (UHPV) technology [10] offering potential simplifica- tude control in the range of 40 km to 70 km. For the tions in fabrication and deployment of the aerostat. lower altitude range, these can use high temperature Aerial Platforms with horizontal control; Aero- electronics technologies that are maturing today. Given stats at a float altitude of 55 km will circumnavigate the the new science that will be enabled by the ability to planet in about five earth days as a result of the superro- repetitively profile in altitude, scientifically productive tating flow and are expected to gradually drift towards missions should be possible with modest payloads. The the nearest pole. The rate is believed to be small a few science would include investigations of a broad habita- meters per second but quite uncertain. Concepts for con- ble zone within the cloud layers trolling this motion have been studied in recent years. Technology work in this time frame should focus on A solar powered Venus aircraft can fly high in the systems for the lower 40 km of the atmosphere includ- clouds where there is sufficient energy. However, ac- ing the near surface environment. Success in this phase cording to Landis [11] in order to stay aloft it must “sta- will hinge on contemporaneous progress in high tem- tion keep” on the sun side of the planet by flying in the perature electronics. This phase could include tech de- opposite direction to the flow. mos of mobile systems with limited scientific measure- The Venus Atmospheric Maneuverable Platform ment capabilities in the near surface environment. (VAMP) concept developed by Northrop-Grumman Long Term 2035-2050: Aerial mobile exploration [12]is a semi-buoyant, maneuverable, solar powered air would be extended to the surface with sophisticated in vehicle conceived for flight in the Venus’ atmosphere situ measurement capabilities. The technology would on both the night and dayside. now also be ready to implement VISE the mission that the Decadal Survey originally conceived in 2003 – an FUTURE ROLE OF AERIAL EXPLORATION aerial platform that would raise surface samples to 55 Aerial platform technology must play a vital role in km for prolonged analysis under benign conditions. . the future exploration of Venus. We envisage a phased References.[1] Hall, J.L, Venus Balloon Technol- approach beginning with proven technologies that oper- ogy Summary, Report of Technology Focus GroupDec ate in the upper reaches of the Venus atmosphere where 7, 2015 [2] Baines, K.H. et al, Exploring Venus with temperatures are near Earth surface ambient. In subse- balloons, Science objectives and Mission Architectures, quent decades, aerial platforms would penetrate deeper IPPW-5 June 2007 [3]Cutts J.A.Probing the Interior in the atmosphere in step with advances in the technol- Structure of Venus, KISS Workshop, June 2014, [4] ogy for operating in those environments. Opportunities Grimm, R. et al. (2012) Icarus, 217, 462. [5] Nock, K should be taken to demonstrate these technologies in ad- E. Stofan et al, Venus Geoscience Aerobot, AIAA -99- vance of a major commitment of science payloads. 3856, 1999 [6] Friedlander A.L. and H. Feingold, Near Term (2016 to 2025): the focus should be on AIAA/AAAS Conf., 1978, #1438. [7] PSDS New Frontiers in formulating missions such as Venus Climate Mission, the Solar System 2003. [8] Lunine et el Editors, Solar Sys- endorsed by the 2011 Planetary Science Decadal Survey tem Exploration Roadmap, 2006 [9] Voss, P.B et al Al- and Venera D, a mission under study by a joint NASA- titude controlled balloons for Long Duration Flights on IKI SDT which includes an aerial platform option. Venus, Proceedings IPPW-11 Pasadena abs 8092 2014 These platforms would be based on mature technologies [10] De Jong, M,, Venus Lab and Technology Work- for light gas superpressure aerostats that operate near 55 shop,2015 [11] Landis G.A.et al, Atmospheric Flight on km altitude. In addition, to the atmospheric science Venus, AIAA-2002-0819 [12] Lee, G et al Venus at- these platforms can also address geophysical objectives mospheric Maneuverable Platform Science Vehicle through the use of infrasound generated by Venus concept, Venus Lab and Technology Workshop 2015 quakes, electromagnetic sounding using Schumann res- onances, and searching for remnant magnetism. CONSIDERATIONS FOR ATMOSPHERIC SAMPLE RETURN FROM THE HABITABLE ZONE OF VENUS. P. G. Athul1†, Y. Lu2†, E. Shibata3†, S. J. Saikia4†, and J. A. Cutts‡, [email protected], [email protected], [email protected], [email protected], †School of Aeronautics and Astronautics, Purdue University, 701 W. Stadium Ave., West Lafayette, IN, 47907, ‡NASA-Caltech Jet Propulsion Laboratory, 4800 Oak Grove Drive, Pasadena, CA, 91109, [email protected].

Earth’s “Twin” Planet: Venus is often called the Venusian cloud layers [2]. Three modes of Earth’s “twin” because it is also a rocky planet and aerosol particles corresponding to various sizes have has approximately the same size and mass as that of been identified. The smaller Mode 1 (~0.2μm) and Earth. Though the Venusian surface is a blistering Mode 2 (~1.2μm) particles have sulfuric acid as their 464° C and 92 bar pressure—conditions too hostile to primary constituent while the largest of the three support any form of life as known on Earth—the called the Mode 3 particles, ranging in radii from few atmosphere near 55 km altitude has a pressure of μm to tens of μm, are believed not to be primarily approximately one bar and a temperature of about 20° C, which could potentially harbor life forms or composed of sulfuric acid. An unknown ultraviolet may contain evidence of biological activity in the absorber in addition to the known compounds has past [1]. been identified to exist in the upper cloud layers. While numerous candidates have been suggested for both the Mode 3 particles as well as the unknown UV absorber, the composition of these particles remains unknown [2]. Grabbing Samples from Venus: An atmospheric sample return would accomplish the first and third of the Venus science goals. Although any samples returned would have relatively small masses compared to in-situ measurements, having a physical sample in a laboratory has its advantages [3]. Within a lab setting, more accurate and powerful equipment can be used without having to worry about the mass Figure 1: Venus photo from the Magellan mission. penalties on the spacecraft. With proper storage, [Source: NASA/JPL] these samples could be used time and again with new generations of analytical techniques and Venus Science Goals: Considering the similar technologies. As seen with lunar samples brought size and distance from the Sun, understanding how back by the Apollo spacecraft, any samples returned Venus came to be may lead to a better understanding can be retested with new hypotheses, without having of how planets are formed. In the Venus Exploration to send another spacecraft to perform more science Analysis Group (VEXAG) Goals, Objectives, and Investigations document [2], three primary goals are measurements. identified for future exploration: Search for Life: Of particular interest in the 48- 1. Understanding the atmosphere’s origins and its 60 km habitable zone in the atmosphere of Venus are evolution, as well as the climate history, the various cloud layers which are known to contain aerosol particles of various sizes and shapes which 2. Determining how the surface and interior are not yet fully understood. Numerous theories have evolved, and been proposed for how life might be able to evolve 3. Understanding the interior-surface-atmosphere and survive in these cloud layers. One hypothesis interactions over time, as well as if liquid water suggests that sulfur-based compounds can use was ever on Venus. ultraviolet light in a photosynthetic-like process [4]. Cloud and Haze Chemistry and Dynamics: Characterization of Mode 3 particles, which are One of the subgoals of the first objective is the larger than other cloud droplets as well as the morphology and chemical makeup of the aerosols in unknown UV absorber, will conclusively answer

1 speculations about life in the Venusian cloud layers. Atmosphere Sample Return: However, not all While the chances of life existing on the Venusian atmospheric sample return missions are equal in clouds are low, it cannot be ruled out completely terms of cost and complexity. There are mainly two until conclusive scientific data shows otherwise. types of atmospheric sample return missions: (1) a single spacecraft performs a high-altitude flyby over Venus and collects samples from the upper atmosphere (>100 km); (2) lower altitude atmospheric sample return that uses an additional element to dive deeper into the atmosphere to collect and lift gas samples from the lower atmosphere (~55 km) to rendezvous with an orbiting spacecraft which finally returns to Earth [7]. While the former concept requires fewer technology developments, the science return is incomparable to the latter concept which allows for the collection of particles and aerosols from the clouds in the lower altitudes. The astrobiological relevance and technical feasibility of Figure 2: Structure of the Venusian atmosphere returning samples using various sample collection [4]. methods from the lower Venusian atmosphere have been studied in the past [8]. The study investigated a

high speed atmospheric pass through the lower Challenges of Thick Venusian Atmosphere: atmosphere, a rotating probe tether system, and a Sample return missions from the Moon and comets balloon floatation system to collect atmospheric were successful, but to return surface samples from samples from the habitable zone in the Venusian Venus, it is imperative to have technology atmosphere [8]. developments and some precursor technology Baseline Mission Concept: Some initial results validation missions [5]. These include high of a preliminary mission concept were presented at temperature-resistant balloons that can be inflated at the NASA Planetary Science Vision 2050 Workshop a temperature of 460° C, guidance and control [9]. The baseline mission concept includes a Venus technology for a Venus Ascent Vehicle (VAV) that entry vehicle (shown in Figure 2), sampling can be launched from a balloon, and thermal control mechanism, and a VAV. To keep the cost and risk system for VAV and lander that is also compatible low, the baseline mission concept also uses mostly with the required sample retrieval activities. In heritage technologies to accomplish an atmospheric addition to the severe surface condition at Venus, sample collection, with the only enabling technology there exists highly corrosive species in the being the Heatshield for Extreme Entry Environment atmosphere (i.e. sulfuric acid) at altitudes below 50 Technology (HEEET) thermal protection system km, which further add to the complexity for surface material. On arrival at Venus, the return vehicle is sample return. At the current level of technology, inserted into Venus orbit, and the entry vehicle enters retrieving surface samples from Venus is difficult, the atmosphere and follows a conventional entry, whereas atmospheric samples can be captured and descent sequence. Multiple samples are then retrieved with relatively low cost [6]. A common collected during the descent phase at 60, 55, and 50 theme observed in planetary exploration is to first km altitudes, which is followed by VAV lifting off to perform a flyby, then have an orbiter, a lander, a a 300 km circular orbit where the Venus return rover, and finally a sample return to Earth for vehicle will rendezvous with the sample return analysis. However in the case of Venus, an canister holder and then return to Earth. atmospheric sample return is a logical step after landing probes due to the extremely hostile conditions on the surface which make rovers and surface sample returns outside reach both in terms of technical complexity and cost.

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capture, and sample handoff (see Figure 3). The cost, mass, technology levels, and science returns for an atmospheric and surface sample return will be compared to quantify the benefits and feasibility of an atmospheric sample return mission. Example: Trade Study for Venus Ascent Vehicle. A rocket is required to lift the samples from the Venus atmosphere; however, there are many ways that can be achieved. If samples are being collected at mid-altitude and at subsonic speed, the ascent ∆V is >8000 m/s which mandates a multi-stage rocket, preferably more than three stages to ensure a Figure 2. Entry vehicle for baseline concept [9] reasonable payload mass fractionEach stage can be either solid rocket or liquid rocket. Sample Canisters. To ensure the success of Future Work: The following list are key points sample collection and to reduce risk, samples will be to be looked at in the future: collected at various altitudes and several sample ● If only considering atmospheric sample collection mechanisms will be employed. The sample return, Cubesat-related technologies may be collection mechanism will also be designed to trap leveraged to reduce the mass and cost. and store aerosol particles in addition to atmospheric ● Reevaluate previous sample return gas. A prototype for collection and storage of architectures with state-of-the-art atmospheric samples from the surface of Mars has technologies. been done recently at Jet Propulsion Laboratory ● Technology development plan to enable (JPL) and has been used as a baseline for the canister atmospheric and surface sample return from design for the Venus mission [10]. The canisters, Venus made of aluminum and coated with teflon, will have ● Investigate techniques for collection and a minimum volume capacity of 50 cc of gas, which, storage of aerosol particles. at Earth’s surface condition (equivalent to conditions at 55 km at Venus), can hold ~60 mg of atmospheric References: [1] Grinspoon D. H. and Bullock M. samples in addition to aerosol particles. Previous A (2007) American Geophysical Union. [2] Herrick sample return missions, such as Stardust, only R. et al. [3] Drake M. J. et al. (1987) Eos, 68. [4] returned 1 mg of sample from comet Wild 2 [11]. Schluze-Makuch D. and Irwin L.N (2010) Cosmic Trade Space Studies: In this study, we focus on Biology. [5] Gershman R. et al. (2000) Aerospace the trade studies for a mid-altitude atmospheric Conference Proceedings. [6] Sweetser T. et al. sample return mission. Trades involving all phases of (1998) AIAA/AAS Astrodynamics Conference. [7] the mission, from launch to sample handoff, will be Sweetser T. et al. (2003) Acta Astronautica, 52. [8] considered, and the mass, risk, and science return will Schluze-Makuch D. et al. (2002) First European be compared to the aforementioned baseline concept. Workshop on Exo-Astrobiology. [9] Shibata E. et al. The phases of the mission that are considered are (2017) PSV2050. [10] Eric Kulczycki et al. (2013), launch, cruise to Venus, Venus capture, terminal IEEE. [11] NASA, (2006) Stardust Sample Return. descent, sample collection, Venus ascent, orbiter/sample rendezvous, cruise to Earth, Earth

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Figure 3. Trade space for all the phases. The blue boxes indicate which options were selected for the baseline.

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LITHIUM COMBUSTION POWER SYSTEM FOR SPACECRAFT LANDERS – POSTER & PRESENTATION. Christopher J. Greer1, Michael V. Paul2, and Alexander S. Rattner1. 1Department of Mechanical and Nuclear Engineering, The Pennsylvania State University, University Park, PA 16802. Email: [email protected], [email protected], 2Applied Physics Laboratory, John Hopkins University, Laurel, MD 20723. Email: [email protected]

Introduction: There is a mission class to solar-de- technology has been demonstrated to deliver high- prived situations where metal-fueled, combustion-based power (0.5-1.15 MW) and low-power (3-25 kW) oper- power plants have been proposed [1]. The surfaces of ation at Earth-based conditions [5]. Due to the extreme Venus and Europa are two primary examples of such environments of both Venus and Europa, detailed mod- situations. The longest lasting Venus lander, Venera 13, eling and experimental characterization of the proposed survived for just over two hours due to power and cool- power plant concepts is needed to assess its potential. ing system limitations [2]. The low energy density and Present Investigation: A detailed thermodynamic operating temperature constraints of current battery and heat transfer model is being developed to predict technologies prevent their use for longer duration mis- performance at target operating conditions. The model sions. The opacity of the Venus atmosphere prevents the receives inputs of oxidizer mass flow rate and ambient use of solar photovoltaic generators for the desired conditions to predict the thermal power output of the re- power range in the hundreds of Watts [3]. Plutonium actor system, accounting for heat losses. thermoelectric generator powered systems would sig- Preliminary model results show that a LCP system nificantly increase hardware and launch costs. There- with an in-situ, carbon-dioxide oxidizer could power a fore, a new power source will be required to enable fu- Venus lander for up 5 days with 300kg of fuel at 2.3kW ture missions in the Venus environment. of power. Greater mission durations are possible up to The COMPASS lab ALIVE (Advanced Lithium Ion 30 days with varying power outputs. Similarly, model Venus Explorer) report suggests the use of a chemi- results indicate a LCP system with a sulfur-hexafluoride cally-fueled power and cooling source for a Venus oxidizer could power a Europa lander at 200 W with lander mission [2]. The chemical power source is a lith- Thermoelectric Generators (TEG) for up to 14 days or ium combustion reactor attached to an Advanced Stir- with a Stirling engine up to 30 days. ling Radioisotope Convertor (ASRG) power generator In ongoing work, two different experiments are be- and refrigeration system. This technology presents an ing initiated. An Experimental Lithium combustion Fa- in-situ opportunity to utilize the mainly carbon-dioxide cility (ELF) is under construction to validate the heat atmosphere of Venus as the oxidizer for the combus- transfer model. This facility will be used to perform tion-based power plant. The products of lithium com- combustion experiments with both carbon-dioxide and bustion with carbon dioxide are denser than the reac- sulfur-hexafluoride oxidizers. A second experimental tants (condensed phase), which means that the Venus study under the NASA Hot Operating Temperature atmosphere would be passively drawn in by the com- Technology (HOTTech) program will test a concept to bustion reaction, but there would be no exhaust from the convert the combustion heat to electrical power with a power system to contaminate the environment. Accord- high-temperature turbine power cycle. The system ing to the ALIVE mission concept, estimated power re- model will be refined with experimental data to aid de- quirements for their future Venus lander is up to 0.3 kW sign of a future packaged prototype system. This com- electrical power and 2.0 kW cooling power [2]. This bination of experimentation and modeling will enable mission design could enable a Discovery-class mission the characterization of LCP technology for spacecraft with a duration of hundreds of hours on the surface, applications. without the need for high-temperature electronics. Conclusions: The goal of this study is to character- A proposed mission for a Europa Lander suggests ize the proposed power plant concepts for their target batteries as the power source with an expected mission environments, and inform engineering design. The ver- duration of 10 days [4]. A lithium-combustion power satility a LCP system provides a lander opportunities, (LCP) system with a sulfur-hexafluoride oxidizer could including: in-situ resource utilization, higher power in- enable a mission of the same duration with the added strumentation, variable power output, and longer mis- benefits of the sensible heat keeping the spacecraft sion durations. Experimental and modeling efforts will warm and the ability to deliver high thermal or mechan- indicate the potential for LCP systems. These findings ical power output on demand. TRL 9 underwater power will also enable assessment of the potential to meet systems called Stored Chemical Energy Power Source other space exploration needs. (SCEPS) system reacts on board sulfur-hexafluoride Acknowledgements: This work is supported and liquid lithium to power undersea vehicles. SCEPS through NASA Innovative Advanced Concepts (NIAC) Phase II funding, and NASA’s Planetary Science Divi- sion’s HOTTech funding. References: [1] Miller, T. F., Paul, M. V., & Oleson, S. R. (2016). Combustion-based power source for Venus surface missions. Acta Astronautica, 127, 197–208. [2] Michael Paul, Steven R. Oleson, Melissa L. McGuire, Carl E. Sandifer II (2012) COMPASS Final Report: Advanced Lithium Ion Venus Explorer (ALIVE) -CD–2012–72. [3] Geoffrey A. Landis, and Emily Haag (2015) Analysis of Solar Cell Efficiency for Venus At- mosphere and Surface Missions. [4] Berger, Eric. "At- tempt No Landing There? Yeah Right—we’re Going to Europa." Ars Technica. N.p., 17 Nov. 2015. Web. 25 Jan. 2017. [5] T.G. Hughes, R.B. Smith, and D.H. Kiely (1983) Stored Chemical Energy Propulsion System for Underwater Applications.

NUMERICAL AND EXPERIMENTAL INVESTIGATION OF BOUNDARY LAYER MANIPULATION WITH MAGNETIC FIELDS FOR HEAT FLUX MITIGATION DURING ATMOSPHERIC ENTRY. P.P. Upadhyay1, R.Tietz2, M. Dropmann3, S. Fasoulas4, R. Laufer5, T.W. Hyde6 and G. Herdrich7, 1,2,3,4,7(Institute of Space Systems (IRS),University of Stuttgart, Pfaffenwaldring 29,70569 Stuttgart, Germany) 5,6(Center for Astrophysics, Space Physics and Engineering Research-CASPER, Baylor University, Waco, Texas 76798-7310) [email protected], [email protected], [email protected], [email protected], [email protected], [email protected], [email protected]

Introduction: The high heat loads during plasma wind tunnel PWK1 [1] at the Institute of Space atmospheric entry make a thermal protection system Systems (IRS) at the University of Stuttgart, have been (TPS) necessary for all missions that enter the conducted to analyze the direct influence of a dipole atmosphere of a planetary body. As the TPS can be of magnet on the compression shock in front of a probe considerable mass it is desirable to reduce the heat flux body in a supersonic high enthalpy plasma flow [2]. to as much as possible to keep the respective mass low. Second experiments in a GEC reference cell at the Magneto-hydrodynamic(MHD) assisted hypersonic Center for Astrophysics Space Physics and flow control has experienced a renewed interest over Engineering Research (CASPER) at Baylor University the past decade : as it is foreseen a way to mitigate the have been conducted to study the separation of charges incoming surface heat flux on re-entry vehicles, by within the boundary layer when a magnetic field is modifying the dynamics of flow field and control of present [3]. the shock position through strong applied magnetic Experiments in PWK1. Using the magneto-plasma fields. Magneto-hydrodynamic flow interactions for dynamic plasma generator RD5 [4] in the plasma wind high speed aerospace applications have been studied tunnel facility PWK1 at IRS a set of experiments have dating back to the late 1950ies and early 1960ies. been conducted where a probe body based on the These studies focused primarily on the flow control ‘European Standard’, with integrated dipole magnet applications possible with MHD interaction for with dipole axis parallel to the plasma flow, has been purposes such as drag and peak heating modulation. exposed to a high enthalpy plasma jet. For simplicity The basic theory focuses on generation of a strong magnetic field around the body, which produces the the noble gas Argon was used for the experiment and a Lorentz force acting on the flow through an interaction plasma jet with a velocity of 3100 m/s and an enthalpy with the ionized flow behind the strong bow shock. In of 50 MJ/kg was formed and directed at the probe this paper the possibility of heat flux mitigation or body. For reference both experiments with and without redistribution by using MHD effects is investigated magnetic field where performed. The boundary layer both experimentally and numerically. was then observed with an emissions spectroscopy setup and two-dimensionally spatially resolved spectra have been taken. Experiments in GEC reference cell. In the GEC reference cell a static cold Argon plasma has been produced and a magnet covered by a glass plate has been introduced into the plasma. Micrometer sized Fig 1. MHD effect in front of Magnetic Probe mono disperse particles have been introduced into the plasma and tracked with a particle velocimetry setup. Due to the electric charging of the particles in the plasma the particle trajectories could be used to determine vector fields of the electric forces within the boundary layer that allow to draw conclusions about the charge distribution caused by the plasma magnetic field interaction. Respective experiment have been done with magnetic dipole axis parallel and normal to the surface and furthermore without magnet for reference. Fig 2. Magnetoplasmadynamic generator RD5 Experimental Results: The experimental results

from PWK1 yield information regarding both the Experimental Setup: Two sets of experiments charge separation and the boundary layer detachment have been conducted to investigate two aspects of the while the results from the GEC reference cell only boundary manipulation. First experiments in the provide information about the charge separation as no RD5 (Fig.2) [6, 7] analysed the interaction shock is present in the static plasma which could be investigated. Shock Detachment. The results from PWK1 show that the radiation maximum, which corresponds to the shock position, shifts with changing magnetic field strength. An increase of magnetic field strength leads to an increase in shock distance from the probe. An increase in shock distance is favorable as an increased shock distance leads to a decrease in heat flux. . These Charge Separation. The results from the GEC investigations in combination with the results in reference cell indicate a strong charge separation due references [6, 7] affirmed the effect of charge to the magnetic field plasma interactions. Electrons are separations due to the morphology of the applied transported along the magnetic field lines, while ions magnetic field. are too heavy to be strongly influenced by the magnetic field. This leads to strong electric fields Numerical Results: which then can have a strong effect as well on ion motion. The charged particles appear to be shielded from regions where the magnetic field lines are horizontal to the surface and are transported to regions where the field lines intersect with the surface. The resulting qualitative electric fields are presented in Figure 3. The emission spectroscopic results confirm this observation, showing increased ion emission in regions to which the charged particles are being transported, while the shielded regions feature decreased ion emission [4].

Fig 4. Numerical simulation of RD5 generator with magnetic probe Within the current investigation, the above mentioned experiment has been numerically simulated (Fig.4) as an attempt for validation using In-house MHD-code SAMSA. Further the limits of the experimental approach were extended to higher magnetic field strengths with prime focus on behaviour of the shock, in particular shock standoff distance or the boundary layer geometry respectively.

Fig 3. Qualitative electric fields References: [1] Knapp A. et al. (2012) Open Plasma Phys. J., 5, Numerical Simulation: 11-22. The numerical code ‘Self and applied field MPD [2] Dropmann M. et al. (2015) Phys. Rev. E, 92, thruster Algorithm’ (SAMSA) [5] is used for 023107. investigations of basic plasma processes in [3] Auweter-Kurtz M. et al. (1998) in Proc. Third Eur. magnetoplasmadynamic (MPD) thrusters and MHD Symp. Aerothermodyn. Space Veh., 529 relevant plasmas. The code is an in-house code under [4] Dropmann et al. (2016) Trans. JSASS Aerospace development at IRS. Based on an axis-symmetric finite Tech. Japan 14(ists30), Pe_21-Pe_26. volume method on unstructured, adaptive meshes and [5] Haag D. 40th AIAA Plasma dynamics and lasers under the assumptions of continuum flow, condition of conference,(2009) quasi-neutral plasma the argon flow is modeled as two- [6] Knapp A. et al. 40th AIAA Plasma dynamics and fluid plasma consisting of heavy particles (neutral and lasers conference, (2009) ionized argon) and electrons. [7] Ono N. Journal of IAPS, Vol.16 No.1, pp. 1-6. , Problem definition and modeling for Simulation. June,(2008) Early investigation as discussed in the experimental part conducted at IRS using plasma facility PWK1 together with magnetoplasmadynamic (MPD) source SUPERSONIC PARACHUTE TESTING USING A MAXUS SOUNDING ROCKET PIGGY-BACK PAYLOAD. J. S. Lingard1, A. Saunders1, J. C. Underwood1, S. B. Rogers1, J. Merrifield2, J. Caldwell2, J. Longo3 and L. Ferracina4, 1Vorticity Ltd, Chalgrove, Oxfordshire, OX44 7RW, UK, 2Fluid Gravity Engineering Ltd, Ems- worth, Hampshire, PO10 7DX, UK, 3European Space Agency, Noordwijk, Netherlands, 4ATG Europe B.V. on behalf of the European Space Agency, Noordwijk, Netherlands.

Introduction: A re-entry vehicle, SuperMAX, has sentative. Therefore, an aerodynamic database for the been designed under European Space Agency funding specific geometry of the piggy-back vehicle was creat- that can be used as a test bed for supersonic parachutes. ed in order to investigate the vehicle’s flight dynamics. The vehicle will be launched aboard MAXUS-9 sound- The vehicle was designed to align itself with the ing rocket from the Swedish Space Corporation’s facil- flow during the early part of entry so that it can with- ity, Esrange, in Kiruna, Sweden in April 2017. stand the significant peak aerokinetic heat flux, the Supersonic tests in large wind tunnel facilities are total heat load and the aerodynamic forces whilst ac- costly and, since the mothballing of the AEDC 16S commodating the test parachute and instrumentation. facility, parachute testing has been limited, by tunnel The recorded data includes high- and low-speed video, blockage, to articles with a diameter <0.83 m or to GPS position and velocity, 3-axis accelerations, 3-axis modest Mach numbers. rates and vehicle body temperatures thermocouples Full-scale supersonic parachute testing is also ex- positioned at selected points within the vehicle’s nose pensive, on the order of tens of millions of dollars for a and adjacent surfaces. The recorded data will enable series of tests. Therefore, a test technique that allows deployment and flight analysis of the parachute from representative supersonic testing of subscale para- the supersonic regime through transonic to subsonic. chutes at modest cost is required. The measurement of capsule heating will provide in- To obtain useful data from sub-scale parachute situ re-entry aerothermal data. Finally, the accelerome- tests, in addition to accurately scaling the model, it is ter and rate gyros will provide flight stability data that necessary to scale the test parameters appropriately: can be used to correlate CFD modelling. Mach number, parachute model porosity, parachute The parachute that will be tested on this mission stiffness, forebody Reynolds number. Therefore, the will be a 1.25 m nominal diameter Huygens heritage selected trajectory must provide suitable deployment Disk-Gap-Band. It was selected since this type of para- conditions such that the matching criteria can be met. chute is well characterized, allowing a good assessment The SuperMAX parachute test vehicle will be re- of the piggy-back test methodology. leased from the upper stage of MAXUS-9 at an altitude The vehicle comprises a stainless steel forward aer- of ~192 km and upward velocity of ~3 km/s. The vehi- oshell incorporating a tungsten ballast mass to ensure a cle will then continue on a ballistic trajectory up to an forward center of gravity position. The cylindrical aft apogee of ~710 km before descending and starting its cylinder is manufactured from CFRP with a refractory re-entry. Prior to atmospheric interface the vehicle will coating for thermal protection. Many of the internal be travelling at approximately Mach 12. components and mechanisms were manufactured from The physical characteristics of the piggy-back vehi- engineering plastics using rapid prototyping in order to cle were chosen to prevent interference with the optimize their geometry and reduce costs. MAXUS upper stage. The mass limit was 15 kg with a The vehicle has been design to be optimizable so maximum diameter of 290 mm and a maximum overall that not only parachutes but also other forms of decel- length of 248 mm. The interface with the MAXUS erators, such as inflatables, and thermal protection ma- vehicle also required that the piggy-back vehicle inter- terials can be tested. faced with three pre-existing mounting locations on the This paper will present an overview of the mission MAXUS payload. Assuming the vehicle could fit with- design, the capsule design and the results of the flight in this mass and volume envelope the design was com- in April 2017. Comparisons will be drawn between the pletely optimizable. Following a configuration trade results from this test and those obtained by convention- study a 60° sphere cone nose and truncated conical aft al wind tunnel techniques. surface was selected. Since the interface requirements of the piggy-back vehicle with the MAXUS rocket was most easily achieved with a cylindrical aft body with a flat aft face, neither the aerodynamic database of the Mars Micro- probe nor the Stardust vehicle were suitably repre- Upcoming Exo-Brake and Nano-Sat Advanced Flight Experiments –TechedSat 6,7,8 M.Murbacha*, A. Guarneros-Lunaa, R. Alenaa, A. Dono Perezb, A. Tannerb, J. Whelessb, S. Smith, A.Salas, C. Priscalc J. Pleaterc, a NASA, Ames Research Center, Moffett Field, CA, U.S. 94035, b Millennium Engineering & Integration Company, NASA Ames Research Center, Moffett Field, CA, U.S. 94035 cStinger Ghaffarian Technologies, Inc, NASA Ames Research Center, Moffett Field, CA, U.S. 94035

Introduction: With the success of TechEdSat- 5/Exo-Brake mission conducted in March, 2017, the following incremental developments are discussed for the currently planned successive experiments. The TechEdSat-6 will evolve from the predecessor experiments with a more reliable and improved manner of controlling the Exo-Brake. In addition, with the improved sensor suite, an on-board guidance (with over- ride capability) will be instituted for the first time. The TechEdSat-7 will incorporate a high-density Exo- Brake packing system in order to improve the volumetric efficiency and lower the ballistic coefficient. This may also have some utility as a nano-sat system for rapid disposal. The TechEdSat-8 is designed as a long 6U which will incorporate a high temperature Exo- Brake and improved front end ablator. Lastly, the manner in which an Exo-Brake system may contribute to nano-satellite scale planetary missions will be discussed

Surviving the Impact: Core Quality Testing in a New Orbiting Sample Container for Potential Mars Sample Return Scott V. Perino1†, Chad Truitt1, Darren Cooper1, and Tom Komarek1 1Jet Propulsion Laboratory, California Institute of Technology, 4800 Oak Grove Drive, Pasadena, CA 91109, †[email protected]

Disclaimer: Although NASA has no official plans within the EEV. As a result, each of the hardware com- at this time for a mission to return samples from Mars, ponents in the load path between the UTTR soil and the the Program Formulation Office of the Mars Explora- samples could affect loading on the samples. tion Program sponsors ongoing mission concept studies, This paper describes the first efforts by ADT to systems analyses, and technology investments which quantify the effect of Earth impact on Martian rock core explore different strategies for the potential return of quality in a representative environment. An instru- samples from Mars, consistent with the charter of the mented ‘EEV-like’ reusable penetrator called ‘OSV-A’ program and stated priorities of the science commu- was designed, analyzed and fabricated at JPL to support nity. the test effort. Furthermore, JPL’s most recent OS con- Background: The notional Mars Sample Return cept called ‘OS-3E’ [4] was also fabricated for this test (MSR) campaign [1] architecture calls for three effort and fitted with two instrumented sample tubes. A launches from Earth, spaced over several years. The CAD image of the assembled test article is shown in the first mission, called Mars 2020, is a rover that collects figure below and details of the hardware and fully con- and preserves Martian geological samples inside spe- tained instrumentation system are discussed. cially designed tubes. The next mission called Sample Return Lander (SRL), is a rover with an attached launch vehicle used first, for collection and then, for Mars orbit insertion of the samples within in a spherical container called an Orbiting Sample (OS). The last mission called Sample Return Orbiter (SRO) is a spacecraft that provides communications relay, orbital rendezvous and capture of the OS, and Earth-return delivery of the sample container. After the SRO captures the OS, it is then inserted and sealed within an onboard reentry vehicle called an Earth Entry Vehicle (EEV) [2]. Once the SRO approaches Earth, the EEV is released from the SRO, reenters the Earth’s atmosphere, and intentionally impact lands without parachute on a designated landing Reusable ‘EEV-like’ penetrator with OS, Tubes, site at the Utah Training and Testing Range (UTTR) [3]. and Core Samples For a potential MSR campaign to be successful, some knowledge of the as-of-yet unfunded SRO and Impact testing was conducted at JPL using a 26 m SRL mission hardware is needed. To address this issue, tall, purpose-built impact tower. The tower has large JPL has assembled the Mars Advanced Development container filled with specially selected soil at its base Team (ADT). Acting on behalf of the future missions, and is fitted with a pulley system connected to a pneu- ADT performs testing, analysis, and design explorations matic piston and pressure tank. The system can acceler- to ensure that current Mars 2020 hardware is compatible ate EEV-like penetrators and other hardware into the with the future mission needs. soil at speeds in excess of the expected terminal velocity Abstract: One of the most significant loading of an EEV. To date, 27 penetrator impact tests have been events of a potential MSR campaign would likely be the conducted using the system. Of those 27, the last seven EEV impact landing at UTTR. There is concern is that tests were conducted for core quality testing using the the EEV impact landing event could significantly re- OSV-A + OS-3E penetrator. Each test included two Old duce the scientific value of the returned Martian sam- Dutch Pumice rock cores that were made using the Mars ples. As a result, both analysis and testing has been con- 2020 developmental percussion drill [5]. Tube orienta- ducted at JPL to determine the effect of the impact land- tion at impact is a key parameter of interest. Tubes were ing on core quality. During the impact landing the sam- tested in three different orientations relative to the im- ples would be sealed in the tubes, the tubes would be pact surface plane: three tests at +90̊ ‘standing up’, three secured in the OS, and the OS would be contained tests at 0̊ ‘laying down’, and the remaining one test at +45̊ ‘diagonal’. The soil and impact conditions were kept constant as much as possible for all seven tests. Soil and impact conditions were chosen with the goal of hitting the 1300 G peak acceleration requirement levied on the tubes and OS. Post-impact core quality assessments indicate that although some damage did occur to the cores, the damage was not significant enough to reduce the scientific value of the cores. Additionally, tube/sample orientation was also not observed to be a driver for increasing core damage. These findings, although preliminary, indicate that the OS can be oriented with tubes ‘standing up’, a configuration deemed beneficial to many post-M2020 hardware sub-systems.

References: [1] Mattingly, R., and May, L. (2011) “Mars Sample Return as a Campaign,” IEEE Aerospace Conference, Big Sky, MT: IEEE, 2011, pp. 1–13 [2] Dillman, R., and Corliss, J. (2008) “Overview of the Mars Sample Return Earth Entry Vehicle,” Sixth International Planetary Probe Workshop, Hampton, VA: pp. 1–8. [3] Fasanella, E., Jones Y., Knight, N., and Kellas, S. (2001), “Low Velocity Earth Penetration Test and Analysis” AIAA, A1338 [4] Perino, S., Cooper, D., Rosing, D., Giersch, L., Ousnamer, Z., Jamnejad, V., Spurgers, C., Redmond, M., Lob- bia, M., and Komarek T. (2017) “The Evolution of an Orbiting Sample Container for Potential Mars Sample Return,” IEEE Aerospace Conference, Big Sky, MT: IEEE, 2017, pp. 1–17 [5] Chu, L., Brown, K., Kriechbaum, K., (2017) “Mars 2020 Sampling and Caching Subsystem Environmental Development Testing and Preliminary Results,” IEEE Aerospace Conference, Big Sky, MT: IEEE, 2017, pp. 1–10

SPLASHDOWN ON TITAN: PEAK LOADS, PLUNGE DEPTH AND RESURGE TIME FOR CAPSULE IMPACT IN A METHANE SEA. E. A. Leylek1, R. D. Lorenz2, G. Vassilakos3, M. R. Grover1, and J. O. Elliott1, 1Jet Propulsion Laboratory, California Institute of Technology (4800 Oak Grove Dr., Pasadena, CA 91109, USA, [email protected]), 2Johns Hopkins Applied Physics Laboratory, 11100 Johns Hopkins Road, Laurel, MD 20723, USA, [email protected], 3Science and Technology Corporation, Hampton, VA

Introduction: The hydrocarbon seas of Saturn’s Example Results: The results of six tests cases moon Titan are targets of astrobiological interest and will be presented. The first evaluates the maximum to study oceanographic processes, such as air-sea ex- plunge depth into the minimum density sea composi- change. Delivering a capsule to sample and explore the tion expected at Kraken Mare, 500 kg/m3, from a verti- seas is of keen interest for future missions. Under- cal entry. The other five cases investigate the splash- standing the impact loads at splashdown is important down impact loads for varying combinations of entry for the probe design and ensuring instrument survivea- vertical velocity, horizontal velocity, and attitude. The bility. Also, knowledge of the plunge depth and re- highest expected density of 630 kg/m3 is used for the surge time may open the way to making brief scientific seas in these cases to maximize the impact load. A measurements below the surface. Here we explore the time history of the acceleration loads for the impact splashdown behavior of two novel capsule configura- cases is shown in Fig. 2, and a visualization of the im- tions into Kraken Mare, Titan’s largest sea. pact is shown in Fig. 3. Results will presented in more Previous work applied the Von Karman momentum detail for all cases and probes. approach to evaluate splashdown loads for the Mercury and Apollo program [1,2], as well as for basic splashdown analyses in the event that the Huygens probe landed in liquid [3]. Hirano and Miura [4], motivated by Apollo, developed closed-form analytic solutions for spherical and conical nosed bodies. Seiff et al [5] compared the results of these analytical solutions to experimental tests of scale models dropped into water. Lorenz [6] conducted small-scale Huygens tests into kerosene, and more recently performed an extensive series of tests into water of a 1/8 scale model of the proposed Titan Mare Explorer [7].

The present work employs the finite element code Figure 2. Filtered results of LS-DYNA simulations of LS-DYNA to predict the splashdown loads, plunge resultant G-loads at CG of probe on sea impact depth, and resurge time for two candidate probe designs impacting the methane seas of Titan. For the simpler vertical liquid entry cases, the LS-DYNA results are compared to analytical calculations. The applicability of LS-DYNA in the design process of liquid entry planetary probes is assessed. Probe Configurations: Two candidate designs for Figure 3. Response sequence of one case in LS-DYNA probes are investigated (Figure 1). The first is a 45 degree spherecone shape, which leverages stability and References: a sharp nose to reduce impact loads. The second has a [1] Stubbs S. M. (1967) NASA TM D-3980. cylindrical body with hemispherical nose. This allows [2] McGehee J. R. et al. (1959) NASA Memo 5-23-59L. for more flexibility in packaging. A drag plate provides [3] Lorenz R. D. (1994) ESA Journal, 18: 93–117. stability in descent. [4] Hirano Y. and Miura K. (1970) JSR, vol7, 762-764. [5] Seiff, A. et al (2005), PSS, 53: 594-600 [6] Lorenz, R. (2003) IPPW-1 Proceedings, ESA SP- 544, 117-124 [7] Lorenz, R. et al (2015), Aero. J, 119, 1214, 409-431

Figure 1. Probe models INTRODUCING THE LASER-ENHANCED ARC JET FACILITY. P. T. Zell1, G. J. Hartman2, and G. A. Cushman3 1NASA Ames Research Center (Mail Stop 244-19, PO Box 1, Moffett Field, California 94035 Pe- [email protected]), 2Sierra Lobo, Inc. (Mail Stop N229-3, PO Box 1, Moffett Field, California 94035 [email protected]), 3Jacobs Engineering Group, Inc. (Mail Stop N230-3, PO Box 1, Moffett Field, California 94035 [email protected]).

A new combined, convective and radiative heating test capability for entry-vehicle ablative thermal protection materials is now available at NASA’s Ames Research Center. This upgrade to the existing Interac- tion Heating Facility (IHF) arc jet involves the addition of high-power fiber lasers that project 1070nm radiative energy into the existing vacuum chamber. The optically expanded laser beam enters a side window of the test chamber to form a uniform 150 x 150 mm square spot on a wedge shaped model holder that is exposed to the convective arc jet plasma stream. Combined heating levels of up to 180 W/cm2 convective and 300 W/cm2 radiative were demonstrated during a test conducted earlier this year. Future work is planned to increase the radiative spot size up to 425 x 425 mm at radiative heating levels of 100 W/cm2; and to allow the arc jet and lasers to simulate dynamic entry-heating profiles encountered in the entry trajectory. In-Situ X-ray Tomography of Ablation in a Portable Arcjet Facility Isil Sakraker1 and Hannah Böhrk1, 1Researcher in Deutsches Zentrum für Luft- und Raumfahrt, Stuttgart, Germany. [email protected] and [email protected]

Introduction: One of the major issues of the an industrial computed tomograph [1]. The sample hypersonic entry of spacecraft is the interaction of was heated by three halogen lamps focusing on the the vehicle surface with the planet’s atmosphere. center of the circular surface. It was equipped with The thermal protection systems (TPS) must be sized four type K in-depth thermocouples while being sufficiently however without excessive safety mar- exposed to X-rays. The char front propagation was gins that could lead to inefficient designs. Low den- visible in the 2D images (Fig. 2) and could be sity ablative TPS materials provide significantly correlated to the temperatures measured by the inefficient insulation during high speed flights such as depth thermocouples. Initial results showed that Moon return or Mars entry. DLR has the compe- when the char front reaches the thermocouples, the tence and experience in manufacturing a variety of measured temperatures (Fig. 3) agree very well with TPS materials. This study includes two ablative the reaction temperatures previously determined materials: DLR Cork and ZURAM®, which is a from the thermogravimetric analysis (TGA) [2]. carbon fiber preform impregnated in phenolic resin. It is our aim to extend the limits of the current abilities of ground test facilities e.g. arcjets, inductively couples plasma tunnels, etc. with in-situ X- ray tomography of ablation testing. Tomographic techniques can provide time resolved information about char front/pyrolysis layer propagation, pore evolution, fiber shape change, tortuosity, etc. How- ever, it is expensive and difficult, if not impossible, to combine a large scale high enthalpy facility with X-ray tomography. Therefore, two portable high enthalpy facilities have been designed in DLR Stuttgart; a radiation-only and an Arcjet configuration. Such in-situ experimental data will significantly improve the thermo-physical model accuracies where commonly, the pre- and post-test virgin and charred states are interpolated. Fig. 2: t=0 s (top); Charring cork sample t= 24 s (bottom) [1].

Fig. 1: Radiation facility [1]. Fig. 3: In-depth temperatures and chamber pressure during Previous Work: A preliminary study has been radiation exposure [1] compared to the HEATS model. conducted with the radiation furnace (Fig. 1) by The HEATS code of DLR, which is originally exposing a DLR Cork sample to radiative heat flux 2 developed for modelling the transpiration cooling of of 230 kW/m at 265 hPa, while being monitored by TPS, is improved to model ablative terms in 2D [3].

1 A comparison example is shown in Fig. 4 where the 2D image at time step t=127 s is calibrated against the char and virgin densities of DLR Cork. HEATS will be further improved with more experimental data, also to include the swelling. In-situ Arcjet Experiment: A second facility is equipped with an Arcjet and has a similar size to the radiation facility (Fig. 5). The aim is to monitor the microstructural changes caused by ablation processes in the presence of a plasma flow, inducing diffusion and convective/radiative heating to the sample surface. A hypersonic nozzle of area ratio 15 is used and the preliminary testing gas is argon. The flow characterization is ongoing by performing reference heat flux and total/dynamic and reservoir/chamber pressure measurements, as well as plasma radiation by means of emission spectroscopy. The testing conditions and facility operating envelope are currently being identified.

Fig. 5: Mini arcjet facility.

References:

[1] Hannah Böhrk and Rauf Jemmali, "Time resolved quantitative imaging of charring in materials at temperatures above 1000 K," Review of Scientific Instruments, vol. 87, no. 7, p. 073701, 2016. [2] H. Böhrk, "Kinetic Parameters and Thermal Properties of a Cork-Based Material," in 20th AIAA International Space Planes and Hypersonic Systems and Technologies Fig. 4: Comparison of HEATs and X-ray data for Conference, 2015. t=127 s [3]. [3] H. Böhrk, "Thermal Response of an Ablator: The first in-situ experiments of DLR-Cork Model and Time Resolved Imaging," in 21st ablation will be performed in an industrial AIAA International Space Planes and computed tomography system and the results Hypersonic Systems and Technologies will be reported at IPPW-14. At a later stage, Conference, China, 2017. further experiments will be conducted in a synchrotron facility, which will allow for submicron resolution in-situ tomography of lower density ablators thanks to its high photon energies. The results from this in-depth in-situ analysis of the transient phenomena will provide essential knowledge necessary for improving TPS design to higher standards of safety, quality and cost-effectiveness.

2

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TOWARDS IN SITU EMISSIVITY MEASUREMENT IN THE HYMETS ARC-JET FACILITY. Bradley. D. Butler1, Michael Winter1, Scott C. Splinter2, Paul M. Danehy2, Jeffrey G. Gragg2, and David E. Glass2. 1University of Kentucky, (151 Ralph G Anderson, Lexington, KY, 40506, [email protected]), 2NASA Lang- ley Research Center, (8 West Taylor Street, MS 190, Hampton, VA, 23681, [email protected]).

A method of determining a thermal protection system material’s spectral emissivity at high temperatures has been implemented for the Hypersonic Material Environmental Test System (HyMETS) at NASA Langley Research Center and will be presented. A set of fiber-coupled spectrometers, covering the spectral range from 350 nm to 1700 nm, measure the absolute spectral radiance emitting from a material sample 42° from normal. The normalized spectral shape is used to determine the material’s temperature and then the measured spectral radiance is compared to a blackbody at the same temperature to compute the spectral emissivity. The sample surface temperature is also monitored using a two color pyrometer, which is compared to the value obtained by the spectrometers. The HyMETS facility[1] is a 400kW segmented Mach 5 arc-jet with discretized gas injection capable of running variable gas composition including an Earth or Martian simulant atmospheric composition. Depending upon the mass flow rate and delivered power, the conditions vary from a stagnation pressure of 10 Torr to 100 Torr, heat flux of 100 W/cm2 to 500 W/cm2, and a bulk enthalpy of 5 MJ/kg to 50 MJ/kg. [1] Splinter, S., et al. (2011) Comparative Meas- urements of Earth and Martian Entry Environments in the NASA Langley HYMETS Facility, AIAA Aero- space Science Meetings. DESIGN OF A JET INTERACTION GROUND EXPERIMENT IN MARS ENTRY CONDITIONS. E. Tri- foni1, G. Ceglia1, A. Martucci1, A. Schettino1, R. Votta1, V. Mareschi2, A. Ferri2, N. Simioni3, N. Bellomo3, L. Fer- racina4, S. Vijendran5, 1Centro Italiano Ricerche Aerospaziali, Capua, Italy, 2Thales Alenia Space Italia (TAS-I), Torino, Italy, 3T4i Technology for Propulsion and Innovation, Padova, Italy 4ATG Europe B.V. on behalf of ESA/ESTEC, Noordwijk, The Netherlands, 5ESA/ESTEC, Noordwijk, The Netherlands.

Introduction: In order to meet the demanding landing accuracy requirements (<10 km) of a future Mars Precision Lander mission, a key aspect is the need for a guided entry thanks to dedicated attitude manoeuvres performed by thrusters at hypersonic/supersonic conditions along the trajectory. It is well known that jet interaction of the thrusters with the hypersonic/supersonic flow can cause unde- sired forces/torques, which could jeopardize the con- trollability of the entry probe. In this work is presented the design of a ground experiment aimed to assess the jet interaction in Mars entry conditions. CFD simulations of a reference probe were performed at hypersonic/supersonic Mars entry conditions with thruster off/on and the jet interaction resulted in all cases very small in terms of delta aerodynamic moments, but higher at hypersonic speed. To reproduce on ground the jet interaction predicted in hypersonic flight, a test campaign was envisaged in SCIROCCO arc-jet facility. Based on literature survey and on CFD results at flight conditions, the most appropriate flight to ground correlation parameters for jet interaction were defined. The reference Mars entry probe with hot gas thrusters was properly scaled to a ground test article with cold gas thrusters suited for SCIROCCO and the facility upgrade required for the thrusters feeding system was designed. CFD simulations were performed at ground conditions with thruster off/on to confirm the representa- tiveness of the selected test conditions with respect to the flight ones and to estimate the heat fluxes for test article design. Several measurement techniques have been taken into account to evaluate quantitatively and qualitatively the jet interaction effect.

Proposed Abstract – IPPW 14 June 13-17, 2017 Modeling, Simulation and Testing OR Airless Session

MATURATION OF THE ASTEROID THREAT ASSESSMENT PROJECT

J. O. Arnold(1), C. D. Burkhard(2), and E. Venkatapathy(3) \

(1)NASA Ames Research Center, Moffett Field, CA 94035 (650) 604-5265 [email protected] (2)NASA Ames Research Center, Moffett Field, CA 94035 (650) 604-1170 [email protected] (3)NASA Ames Research Center, Moffett Field, CA 94035 [email protected]

Abstract As described at IPPW 12 [1], NASA initiated a new research activity focused on Planetary Defense (PD) on October 1, 2014. The over-arcing function of the Asteroid Threat Assessment Project (ATAP) is to provide capabilities to assess impact damage that any Near-Earth Object (NEO) could inflict on the Earth. The activity includes four interrelated efforts: Initial Characterization (at the atmospheric entry interface); Entry Modeling (energy deposition in the atmosphere); Hazards (on the surface including winds, overpressures, thermal exposures, craters, tsunami and earthquakes) and Risk (physics-based). This paper outlines progress by the ATAP and highlights achievements that are complimentary to activities of interest to the International Planetary Probe community. The ATAP’s work is sponsored by NASA’s Planetary Defense Coordination Office (PDCO), a part of the agency’s Science Mission Directorate.

#&$ ! ! !! &' &("&)!'%&( VISUAL ATMOSPHERIC TURBULENCE DETECTION FOR MOVING OBJECTS IN THE ATMOSPHERE

YUE ZHANG1, YUCHEN LIU1, SHIKUI DONG2, XUN WANG3, YUN SU1

1Beijing Institute of Space Mechanics & Electricity, Beijing, China 2Institute of aerospace Thermophysics, Harbin Institute of Technology, Harbin, China 3Mianyang Institute of fluid physics, China Academy of Engineering Physics, Mianyang, China

Introduction: Atmospheric turbulence will have a atmosphere, and parachute deploying in the great effect on aerospace craft descending and landing atmosphere. in the atmosphere, supersonic aircraft or missile flying Imagi ng posi t i on Imagi ng posi t i on in the atmosphere, and parachute deploying in the of gr ound obj ect s of gr ound obj ect s atmosphere. Control action and flying attitude of those t hr ough di st ur bace moving objects are variable, which are extremely Δ y important stage in the process of flight. So it is a very Imagi ng important technology to detect the interaction between lens Image sur f ace atmosphere and moving objects. Flow form of the di spl acement ε moving air is invisible, how to make it visible to of di st ur bace Deflection realize visual monitoring for atmospheric turbulence is angl e the problem that will be solved in our researching ε work. di st ur bace Background oriented schlieren is a quantitative Uc measurement technique for flow density field, which NO n( x, y, z, t ) n'(x,y,z,t) di st ur bace Refractive comprehensively applied schlieren imaging principle, i ndex f unct i on x x digital camera technology and cross-correlation Groud light algorithm, and has a great application potential in flow field measurement. Background oriented schlieren can implement effectively both in flow filed measurement and temperature filed measurement, and can provide a simple and effective method for large field and Figure 1 Detection principle of Backgroud oriented quantitative flow measurement. Background oriented schlieren schlieren was used to visual optical atmospheric turbulence detection, caused by moving objects in the References: [1] Michael J. Hargather and Gary S. atmosphere. Based on the numerical simulation Settles. (2010) Exp Fluids, 48, 59–68. [2] Toshiharu analysis results of turbulent flow characteristics for M., Stephen B. and Paul M. D. et al. (2015) AIAA moving objects in 10km and 1.2-2.2Ma, flow SciTech, 2015-1692. [3] Masanori O., Kenta H., disturbance detection system was designed and Horoko K. et al. (2011) Meas. Sci. Technol., 22, simulated. Working band, resolution, and frame were 104011. designed, and high precision I-PIV algorithm was improved to extract small atmospheric turbulence image to identify atmospheric turbulence of moving objects. Visual optical atmospheric turbulence detection technology with background oriented schlieren can display ambient airflow field in the whole process for objects moving in the atmosphere. The visual atmospheric turbulence has a great effect on aerodynamic shape optimization and structural improvement. Our research results can be widely applied in atmospheric turbulence detection for aerospace craft descending and landing in the atmosphere, supersonic aircraft or missile flying in the ENHANCED MODELLING OF MARS ENTRY TRAJECTORIES FOR HIGH-MASS PAYLOAD MISSIONS. M. Braun1, L. Peacocke1, and P.J.K. Bruce1, Department of Aeronautics, Imperial College London, London, SW7 2AZ, United Kingdom.

Abstract: Future Mars missions will require high- mass payloads on the order of tens of tonnes to be delivered to the Martian surface. Existing Mars entry vehicles are limited by current launch vehicle fairings to diameters of 4.5 m, which in turn limits the useful payload to approximately 1 tonne, as demonstrated in NASA’s Mars Science Laboratory Mission. Deploya- ble aero-decelerator heatshields that extend to larger diameters following launch are a key enabling technology for such high-mass payload missions. Imperial College London has developed a Mars entry trajectory simulator within Matlab that allows rapid analysis of entry vehicle concepts, including the different potential aeroshapes of deployable heatshields. This model utilizes atmospheric data from the Europe- an Mars Climate Database, and determines the aerodynamic forces experienced by the vehicle using the Modified Newtonian method. The simulator was integrated in an iterative mass minimisation approach to find optimised entry vehicle shapes and corresponding landing strategies. The focus of this initial optimisation exercise was on combining deployable heatshield technology with supersonic retro-propulsion so that the use of parachutes is no longer required. This greatly extends capabilities of future Mars entry, descent and landing systems. Opportuni- ties to minimise entry masses have been identified by tuning the ballistic coefficient, enabled solely by the use of deployable heatshield technology. In addition, a new metric for assessing heatshield performance has been proposed, the heatshield area density, ϕshield. Work is ongoing to develop the simulator to a 6 degree-of-freedom model, and couple it with an aero- structural tool to model flexure and stability during entry. Correlation of the simulator with industry results is planned, in a collaboration with Airbus Defence and Space. This will then allow new and innovative configuration concepts for the deployable entry vehicle to be assessed and optimized, prior to a high-speed wind tunnel test campaign. This presentation will outline the benefits and concepts for deployable aero-decelerator heatshields, the structure of the rapid entry trajectory simulator, the latest results of the mass minimisation and conceptual design study, and the future work planned at Imperial. HIAD Tensile Testing Optimization: J. S. Cheatwood and J. M. Wells

Optimization of Thermal Environment Control for Tensile Tests of HIAD materials. J. S. Cheatwood1, J. M. Wells2, and R. K. Johnson3 1Virginia Polytechnic Institute and State University 2 AMA Incorporated, NASA Langley Research Center, 3NASA Langley Research Center

Introduction: Based on calibration runs, a series of step functions can be defined to achieve a desired oven temperature. In order to quantify the behavior of high- Several of these sequences were developed to achieve a temperature materials for the Hypersonic Aerodynamic variety of target temperatures. By creating an algorithm Inflatable Decelerator (HIAD) technology, a custom that utilizes direct amperage inputs, temperature designed thermal oven for tensile tests is required to oscillations were mitigated, and the role of represent thermal conditions for atmospheric entry. The thermocouple readings was changed from control to testing methods should be suitable for the various monitoring. In the end, results for tensile strength as a structural components of the HIAD that include straps, function of temperature became significantly more cords, and fabrics. Achieving the desired extreme accurate and repeatable. temperatures requires material test specimens to be heated at a rate of 150 degrees C per minute, with a test temperature of up to 500 degrees C. In order to get the most accurate trendline for the strength of material specimens, uniform oven temperature, with minimal fluctuation of that temperature, is desired. Monitoring differences in thermocouples located at various locations on the specimen, such as front to back and top to bottom, provides a measure of the thermal conditions and uniformity throughout the oven. Insulating the oven to minimize the heat loss allows it to reach high temperature and achieve a more uniform temperature profile needed to accurately assess the strength of the material specimens at a given temperature. There were significant variations between thermocouple readings which resulted in uncertainties regarding the uniformness of the thermal environment Figure 1. Temperature Profiles with Thermocouple of the oven. In order to make the thermocouple Feedback Control variations negligible, many modifications to the oven were attempted. Initially, the thermocouples were covered with aluminum foil to reflect back radiation from the lamp, decreasing front-back variations slightly. The size of the glass view window was decreased significantly by adding insulation covered with foil to reduce the view opening to 2” square, to reflect more heat back into the oven, thus minimizing the amount of heat lost through the window. While the temperature variations were rendered negligible as a result of these modifications, there were significant temperature oscillations that still needed to be addressed. Originally, the heating process modulated energy input based on thermocouple feedback to Figure 2. Improved Temperature Profiles using control. This approach introduces an inherent lag in the Amperage-driven Control control process causing temperature to overshoot and undershoot the desired profile. To mitigate the oscillations an amperage-driven control is implemented. MISSION ASSURANCE AND RESIDUAL RISK: THE PERFORMANCE VERIFICATION CHALLENGE FOR TECHNOLOGY INFUSION.

P.J. Gage1and E. Venkatapathy2

1 [email protected], Neerim Corporation, NASA Research Park, Mail Stop 19-46A, PO Box 1, Moffett Field CA, 94035-0001 2 [email protected], NASA Ames Research Center, Moffett Field, CA.

Introduction: Space exploration is inherently Elsewhere, NASA provides risk management pro- risky. Venturing into an uncertain environment is in- cedural requirements for all projects [6], and guidance trinsic to exploration, and the space environment is on implementation [7]. Risk-Informed Decision Mak- certainly hazardous. The combination of hazard and ing (RIDM) is emphasized. Hence, alternate criteria uncertainty constitutes risk to mission success. are available for characterization of maturation residu- NASA’s Policy on Mission Assurance [1] acknowl- al risk. Application of such criteria may alter decision- edges this situation by concentrating on the responsi- making with respect to adoption of new technology. bility to accept residual risk, which is defined as “the In a recent paper presented by Margret Frerking, et remaining risk that exists after all mitigation actions al. [8] proposed a new approach at JPL for evaluating have been implemented or exhausted in accordance technology readiness in order to find the right balance with the risk management process”. between beenfits and risk levels in order to encourage NASA’s Program Management Policy [2] requires new advances in science due to emerging new technol- application of systems engineering processes across ogies. The technology assessment approach (TRA) the life cycle of all projects. One of the metrics defined is a systematic, metrics-based process that assesses the in NASA’s Systems Engineering Policy [3] is Tech- maturity level of new technologies and facilitates the nology Readiness Level (TRL), which uses a 9 level handoff from technology development to engineering development. In the proposed presentation, this ap- integer scale to assess capabilities with respect to de- proach will be assessed for technnolgies associated sign, analysis, manufacture and test. Initial maturity with thermal protection system (TPS). assessment is conducted during Formulation, and up- The complete submission will provide details on dated assessments are required at Key Decision Points some of the challenges associated with TRL assess- (KDPs) of NASA projects. The NASA Systems Engi- ment and its interpretation in a risk framework. The neering Handbook [4] provides guidance on Technolo- specific wording of criteria for achieving TRL 5, 6 and gy Readiness Assessment. 7 will be interpreted and challenged. For a particular The TRL scale is commonly used as a gate for subsystem of interest to the planetary probe communi- technology infusion into flight projects: the recent New ty, Thermal Protection System for Atmospheric Entry, Frontiers Announcement of Opportunity [5] states a range of risk scenarios that are consistent with TRL 6 “Proposals with a limited number of less mature tech- designation will be included. An alternate framework nologies and/or advanced engineeringdevelopments for assessment of development risk will also be pre- are permitted as long as they contain a plan for matur- sented. Finally, the potential for different decision out- ing systems to TRL 6 … by no later than Preliminary comes based on the alternative criteria will be dis- Design Review (PDR) and adequate backup plans that cussed. will provide mitigation in the event that the systems cannot be matured as planned.” Here the TRL scale is effectively being used as a qualitative risk metric: it is assumed that maturity below TRL 6 at PDR is too References: risky to be flight-ready in the mission implementation [1] NASA Policy for Safety and Mission Success timeframe. The nature of the criteria for the higher NPD 8700.1E, Revalidated (2013) levels on the TRL scale introduces a strong bias for [2] NASA Engineering and Program/Project Man- adopting technologies that have been flown previously. agement Policy NPD 7120.4D, (2010) Furthermore, there is an implicit assumption that all [3] NASA Systems Engineering Processes and Re- technologies have a similar maturation rate: no candi- quirements NPR 7123.1B, (2013). date technology below TRL 6 can mature sufficiently [4] NASA Systems Engineering Handbook NASA quickly to reach flight readiness in the time between SP-2007-6105 Rev 1 (2007) PDR and Launch. [5] NASA Announcement of Opportunity New Frontiers 4, NNH16ZDA011O, (2016) [6] Agency Risk Management Procedural Re- quirements, NPR 8000.4A, Revalidated Jan 2014 [7] NASA Risk Management Handbook, NASA SP-2011-3422, Nov 2011 [8] JPL Technology Assessment Guideline, IEEE 2016.

GENERAL REFLECTIONS ON THE DEFINITION OF ANALOGUES. F. Foucher1, F. Westall1, J. Zipfel2, N. Bost1 and the EURO-CARES Team, 1Centre de Biophysique Moléculaire, CNRS, rue Charles Sadron, 45071 Orle- ans, France ([email protected]), 2Forschungsinstitut und Naturmuseum Senckenberg, Senckenberganlage 25, 60325 Frankfurt am Main, Germany.

Introduction: Most astrobiological investigations environments, i.e. study convergent evolution to find have been, are, and will be, focussed on solid materials convergent biosignatures. including rocks, soil, and ices. However, natural mate- Moreover, depending on the precision needed and rials can be very complex in composition, and the po- on their specific use, the degree of analogy of sites or tential traces of life and/or molecules of astrobiological samples may be more or less important. For instance, if interest that they could contain may be very subtle and the aim is to test rover mobility, the composition of the challenging to detect; hence, the importance of prior soil is not relevant and only its mechanical proper-ties preparation for the missions using analogues. Ana- are pertinent. It is thus possible to classify the analogues are terrestrial sites or samples having properties logues into different categories. more or less similar than those expected on a given Analogues and other samples for the EURO- extraterrestrial body. There is a huge variety of ana- CARES project: Analogue samples are complemen- logues on Earth that can be used for many purposes: to tary to other samples used during instrument develop- test spacecraft landing and rover mobility, to test and ment, which are not necessarily relevant to the extra- calibrate instruments and sample preparation systems terrestrial body being studied (such as a colour target for in situ missions before launch, to help interpreta- used to calibrate a camera or a piece of silicon used to tion of data acquired during missions, and to carry out calibrate a Raman spectrometer, for example). laboratory experiments. Analogue samples include The objective of the H2020-funded EURO-CARES minerals and rocks, as well as chemical, biological and project (grant number 640190) is to create a roadmap material samples. for the implementation of a European Extraterrestrial On the use of analogues: It is crucial to cross- Sample Curation Facility (ESCF) that would be suita- calibrate the payload of a mission before launch using ble for the curation of samples from all possible return analogue samples. Thus, we have developed a collec- missions likely over the next few decades, i.e. from the tion of analogue rocks, the International Space Ana- Moon, asteroids and Mars. The EURO-CARES Work logues Rockstore (ISAR, www.isar.cnrs-orleans.fr) in Package n°5, led by Frances Westall and Jutta Zipfel, Orléans (France) that can be used to test and calibrate is dedicated to the different kinds of samples needed in space instruments[1]. This collection was used to make such a facility, including analogues sensu stricto as a “blind test” consisting of analyses of two unknown well as calibration, reference, witness and voucher samples using a part of the ExoMars 2020 payload[2]. samples and standards. More information about the Data from each instrument were presented to geolo- project can be found on the website: www.euro- gists having no prior knowledge of the rocks for inter- cares.eu. pretation. The geologists were able to make relatively Acknowledgements: We acknowledge CNES and detailed interpretations, thus demonstrating that the use the European H2020 program for funding. of complementary payload data can compensate for the References: [1] Bost N. et al. (2013) Planetary technical limitations of the instruments (when com- and Space Science 82-83, 113-127 [2] Bost N. et al. pared with laboratory instruments). (2015) Planetary and Space Science 108, 87-97. On the limit of analogues: The term analogue may be confusing when applied to astrobiology. For instance, while basalts can be considered as an analogue of Martian rocks (since basalts have been found on Mars), analogue life forms, e.g. terrestrial extremophiles are less obvious simply because life has never been found on Mars. The latter may thus be deemed putative analogues to be used for scientific purposes more than for instrument testing. For instance, study the metabolisms of extremophiles living in analogue environments is interesting to search for the fundamental requirements necessary to live in these particular LAGRANGIAN PARTICLE TRACKING IN A DISCONTINUOUS GALERKIN FRAMEWORK TO SUPPORT HIGH-FIDELITY DUSTY FLOW SIMULATIONS OF MARS REENTRY ENVIRONMENTS. E. J. Ching1, Y. Lv1, and M. Ihme1, 1Flow Physics and Computational Engineering, Department of Mechanical En- gineering, Stanford University, Stanford, CA, USA.

Abstract: Recent interest in human-scale missions larly, in contrast with conventional finite-volume tech- to Mars has sparked active research into high-fidelity niques, heating results computed with discontinuous simulations of hypersonic reentry flows. These com- Galerkin methods exhibit substantially lower sensitivi- plex aerothermodynamic environments are character- ties to grid topology and viscous-convective flux for- ized by such physical phenomena as strong shocks, mulations. turbulence, dissociation, radiative heat transfer, and Given the need to accurately assess the impact of surface ablation. A key feature of the Mars atmosphere dust particles on Mars re-entry vehicles, a high-fidelity is the high levels of suspended dust particles, which particle method for simulating particle-laden reacting can significantly influence the flow environment. Dur- flow environments is developed. To exploit the ad- ing dust storms, suspended particles, on the order of vantages of high-order numerical schemes, this method microns in size, can reach altitudes up to 60 km [1], at is formulated in a discontinuous Galerkin framework which approximate flight velocities and Mach numbers using a mesocopic modeling approach. Currently, one- are 7.4 km/s and 40, respectively, based on previous way Lagrangian point-particle tracking is implemented trajectory analyses [2]. in a high-order parallel discontinuous Galerkin solver capable of simulating viscous perfect-gas flows in the Dust particles can interact with shocks, shock lay- hypersonic regime. Ongoing work entails improving ers, and vehicle surfaces [3]. For instance, particles the physics beyond simple Stokes drag, enhancing the crossing a strong shock can induce shock perturbations thermodynamic capabilities, and the consideration of that may affect downstream heat transfer. In addition, two-way coupling between the carrier and dispersed particles may experience break-up and phase changes, phases to allow for energy and momentum exchange. which can alter mass fluxes at the vehicle surface. In The focus of the particle method will be on accurately the post-shock region, particles and the surrounding predicting particle trajectories, shock and flowfield flow exchange energy and momentum. In addition to perturbations, and heat flux augmentation. Develop- experiencing slowdown and deflection, particles can ment will be guided by previous experimental investi- undergo considerable heating, which can subsequently gations of particle-induced heating augmentation in induce radiation to the vehicle surface. Furthermore, hypersonic flows over various blunt bodies [9]. particles colliding with the vehicle surface can enhance The first part of the presentation will provide a erosion of thermal protection systems, and the result- brief overview of the dusty environment on Mars, its ing debris can interact with both particles and the flow. effects on reentry vehicles, and motivation for a high- Finally, particle deposition can increase surface rough- fidelity physics-based particle method. The second part ness, facilitating laminar-turbulent transition and in- will discuss current progress in the development of the creasing surface heat fluxes. particle method and future work. The complex physics prevalent in aerothermody- References: namic environments makes accurate flow simulations [1] P. E. Papadopoulos (1994) Stanford PhD thesis. extremely difficult. Current simulation capabilities are [2] P. Subrahmanyam (2009) Int. J. Aerospace restricted not only by inaccurate models but also by the Eng., 2009, Article ID 326102. limitations of conventional numerical methods [4, 5]. [3] I. Montois et al. “Overview of Dust Effects As such, an attractive alternative to widely used se- During Mars Atmospheric Entries: Models, Facilities, cond-order finite volume schemes is finite-element- and Design Tools.” th based discretizations, namely high-order discontinuous [4] P. Gnoffo and J. A. White (2004) 37 AIAA Galerkin (DG) methods [6]. This family of numerical Thermophysics Conference, AIAA Paper 2004–2371. methods offers a number of advantages over classical [5] K. Kitamura et al. (2010) AIAA J., 48, 763–776. schemes, such as arbitrarily high spatial order of ac- [6] Z. J. Wang et al. (2013) Int. J. Numer. Meth. curacy, geometric flexibility, a high degree of scalabil- Fluids, 72, 811–845. ity, and the utilization of advanced hp-adaptation strat- [7] G. E. Barter and D. L. Darmofal (2010) J. egies. Having demonstrated success in areas such as Comput. Phys., 229, 1810–1827. th aerodynamics, combustion, and turbulence, discontin- [8] E. J. Ching et al. (2017) 55 AIAA Aerospace uous Galerkin schemes have begun to show potential Sciences Meeting, AIAA Paper 2017–0311. for the prediction of hypersonic flows [7, 8]. Particu- [9] L. E. Dunbar et al. (1975) AIIA J., 13, 908–912. FLUID STRUCTURE INTERACTION REBUILDING OF PARACHUTE TESTS J. S. Lingard1, A. Saun- ders1, J. C. Underwood1, S. B. Rogers1, L. Marraffa2 and L. Ferracina3, 1Vorticity Ltd, Chalgrove, Oxfordshire, OX44 7RW, UK, European Space Agency, Noordwijk, Netherlands, 3ATG Europe B.V. on behalf of the European Space Agency, Noordwijk, Netherlands.

Abstract: The European Space Agency recently Within this program further rebuilding of both the funded testing of disk-gap-band parachutes in super- supersonic and subsonic tests has been carried out. sonic flow and sub-scale subsonic wind tunnel and free flight testing of a range of parachute types. The aims of Three supersonic cases: x/D = 9.0 at Mach 1.8 and these Technology Research Programs were to expand Mach 2.25 and x/D = 7.5 at Mach 2.0 were simulated. the knowledge of parachute inflation and flying charac- Additionally eight subsonic tests configurations are teristics in subsonic and supersonic flows. In the sub- being rebuilt including matching particle velocimetry sonic program novel parachute designs were manufac- data at zero degrees incidence and at 7 degrees inci- tures in an attempt to improve both drag and inflation dence, matching a free flight test at high and low alti- characteristics. tude conditions and matching disk-gap-band static aerodynamics testing. Supersonic tests were conducted in the CNRC 1.5 m by 1.5 m Trisonic wind tunnel in Ottawa, Cana- A final simulation will be conducted to derive dy- da. The tests gathered data on inflation, inflation sta- namic aerodynamic coefficient data (pitch damping) bility, drag coefficient and general flight stability at using the free oscillation technique. Mach numbers between 1.6 and 2.2 in the wake of a probe of similar configuration to Stardust. Parachute The paper will discuss the tests conducted and the forces were measured and visual data obtained using results obtained. The FSI simulations of the tests will high speed, standard video and Schlieren. The model be described and results presented Disk-Gap-Band (DGB) parachutes had diameters of 1.5 and 3.0 times the probe diameter and were flown at trailing distances ranging from 7.5 to 10.5 forebody diameters.

The subsonic wind tunnel testing was conducted in the CNRC 9 m x 9 m tunnel. Tests evaluated inflation and drag performance and measured static aerodynamic coefficients at incidence. Further testing of a disk- gap-band parachute was conducted in the CNRC 3 m x 2 m tunnel and particle image velocimetry data obtained.

Subscale free flight tests were conducted using (6 kg) instrumented test vehicles released from helium balloons following ascent to above 28 km. Accelera- tions, rates, orientation, dynamic pressure and GPS position / velocity were measured and the descent recorded with an onboard camera.

Three free flight tests at full scale were also conducted using an instrumented test vehicle released from a helicopter.

In support of the post-test analysis and evaluation of the supersonic tests a single test was rebuilt using fluid structure interaction (FSI) simulation. The match to the test drag data was very good.

Time Resolved X-Ray Imaging and Modelling of an Ablator Hannah Böhrk1 and Isil Sakraker1, 1Researcher in Deutsches Zentrum für Luft- und Raumfahrt, Stuttgart, Germany. [email protected] and [email protected]

Introduction: To shield atmospheric re-entry ve- (TGA) measurements are obtained at different heating hicles from extreme heat load, thermal protection sys- rates and used in a parallel distributed activation ener- tems are required. In order to design thermal protection gy model to describe the overall rate of decomposition. systems or to select suitable shielding materials, the A model for the thermal response of decomposing ma- thermal response of the material to the load must be terial is introduced and the results were compared to known. In this study, ablation is observed in-situ by X- the X-ray images sequence. ray imaging and the data are compared to numerical Comparison: Correlating the images to thermo- computations. couple data agrees well with pyrolysis temperature and In-situ X-Ray Imaging of Ablation: A sample of thermal response. Thermal response and char layer a cork-based material, instrumented with in-depth propagation coincides well among model and experi- thermocouples, is encapsulated in an evacuated cell ment. The preliminary experiment-model comparisons (Fig. 1) heating a sample surface with a radiative heat of in-depth temperatures are plotted in Fig. 3. More flux of 230kW/m2 [1]. Simultaneously, the sample was details on material characterization, modelling and imaged in an X-ray computed tomograph. The 2D im- experimental results will be provided in a poster. ages show the sample surface and the in-depth progression of the char front dividing the char layer from the virgin material as shown in Fig. 2.

Fig. 3: In-depth temperatures and chamber pressure during radiation exposure [1] compared to the HEATS model. Fig. 1: Radiation facility: X-ray setup.

Modelling: The in-house HEATS code was im- References: proved to model ablative phenomena. For the computations, material values like thermal conductivity, spe- [1] Hannah Böhrk and Raouf Jemmali, "Time resolved quantitative imaging of charring in materials at cific heat and reaction rates previously had to be de- temperatures above 1000 K," Review of Scientific termined. The cork-based ablative material is investi- Instruments, vol. 87, no. 7, p. 073701, 2016. gated for the rate of thermal decomposition from room temperature up to 1823 K. Thermogravimetric analysis

Fig. 1: Sequence of 2D X-ray scans. Development of High-Fidelity Material Response Modeling for Woven Thermal Protection Systems. D. Z. Dang1, E. C. Stern2, and I. D. Boyd1, 1University of Michigan, 2NASA Ames Research Center.

Introduction: The current work is focused on characterizing and modeling key physicochemical phenomena for woven thermal protection systems (TPS) through the development of a coupled thermostructural ablation solver. The architecture and material composition of woven TPS (figure 1, [1]) may be precisely tailored to meet the requirements of a specific mission, making this technology a promising potential for future exploration missions. The Adaptive Deploy- able Entry and Placement Technology (ADEPT, [2]) in Figure 1. Schematic of 3-D Woven TPS [1] figure 2 and Heatsheld for Extreme Entry Environ- ments (HEEET, [3]) in figure 3 are the main projects intended to leverage the promising potential of woven TPS. For TPS in general, mechanical ablation is a phenomenon of interest; however, for woven TPS where tows or layers may be lifted into the boundary layer and quickly ablated, this becomes a unique problem of interest. Under certain arc jet conditions, increased recession has been observed for both the ADEPT’s heatshield and HEEET. It is expected that this is at- tributable to thermo-structural phenomena both in- depth and at the surface. The project intends to implement high-fidelity stress-dependent material property Figure 2. Arc jet Air Testing of ADEPT [2] models derived from experimental data (or through microscale modeling, if necessary) in the thermostructural ablation solver. Key challenges include coupling of mesh motion induced by both surface recession and structural deformation and coupling of thermo-structural parameters (e.g., porosity-stress, temperature-strain relationships). The modeling will use a macroscopic approach, using typical volume-averaged material response (solid material decomposition, pyrolysis gas continuity, and solid and gas energy) and structural (linear elastic) equations. Results from arc jet tests performed by NASA woven TPS development Figure 3. Arc jet Testing of HEEET [3] projects will be used to assess the validity of the modeling by the end of the project. The methodology, implementation, and preliminary results for simplified test cases will be presented in this workshop. The overall project is expected to yield results that will serve as innovative tools, with key physicochemical phenomena properly characterized, modeled and simulated, to better guide woven TPS design for future space exploration missions. References: [1] Venkatapathy, E. (2014) OPAG Meeting [2] http://www.nasa.gov/image-feature/nasa-completes-successful- heat-shield-testing- forfuture-mars-exploration-vehicles. (2015) [3]https://science.nasa.gov/technology/technologystories/speciali zed-weaving-techniques-enable-new-heatshield-planetary- exploration. (2017). Figure 4. Preliminary Coupling Scheme Sensitivity of Oxidation Processes on Heat Fluxes Determination M. Mione, B. Massuti-Ballester, A.S. Pagan and G. Herdrich Institute of Space Systems, University of Stuttgart, Pfaffenwaldring 29, 70569 Stuttgart, Germany Email: [email protected]

Abstract: Atmospheric re-entry provides one of 765). [3] M. Mione. “Determination of Transient Heat the harshest environments a spacecraft can face, since Fluxes of High Temperature Materials in Plasma Wind high heat loads are transferred structure. With special Tunnel”. Master Thesis. Institute of Space Systems, regard to LEO re-entry environment for re-usable University of Stuttgart. January 2017. [4] F. Sanson, N. Thermal Protection Systems (TPS), the complexity of Villedieu, F. Panerai, O. Chazot, P. M. Congedo, and the heating problem lays behind the interdependence T. E. Magin. “Quantification of Uncertainty on the between gas-surface interaction mechanisms, especial- Catalytic Property of Reusable Thermal Protection ly catalytic recombination at surface, oxidation behav- Materials from High Enthalpy Experiments". In: Ex- ior and surface emissivity change [1]. A test campaign perimental Thermal and Fluid Science 82 (2017), pp. performed on several Silicon Carbide (SiC) enabled to 414-423. investigate the oxidation/catalysis coupling and to assess the sensitivity of oxidation processes on the determination of heat fluxes. Pyrometric measurements of sample’s front and back surfaces allowed to derive transient heat fluxes, provided with consistent thermal properties [2]. The sensitivity analysis showed that the derivation of heat fluxes from plasma wind tunnel measurements is prone to large uncertainties deriving from the assessment of oxidation behavior of the sample’s front surface. Specifically, large uncertainties have to be expected from the assumption on the spectral emissivity of the sample’s front surface, where the assessment of SiO2 layer formation plays a key role, which in turn drive the overall uncertainties on the determination of heat flux distribution on the tested specimen and sample holder [3]. Sample’s backside constellation in the so called Standard 50 mm material probe has been many times treated as adiabatic [2], [4], and with an effective emissivity of 1 [2]. The validity of these two assumptions have been examined within this work. The outcome are opposite for each assumption. While the heat losses from the sample’s back surface to the insulation structure reach 10 % of the radiative heat flux and adiabatic assumption can not be made, it was proven that effective emissivity behind the material sample stays above 0.99 for steady-state conditions. A parametric study of both effective emissivity and heat losses have been performed under variation of incoming heat fluxes and material properties.

References: [1] G. Herdrich, M. Fertig, D. Petkow, A. Steinbeck, and S. Fasoulas. “Experimental and Numerical Techniques to Assess Catalysis”. In: Pro- gress in Aerospace Sciences. 48-49. January 2012, pp. 27-41. [2] B. Massuti-Ballester, S. Pidan, G. Herdrich and M. Fertig, Recent catalysis measurements at IRS (2015), in: Advances in Space Research, 56:4(742- 6-DOF ENTRY TRAJECTORY SIMULATOR FOR DEPLOYABLE MARS AERO-DECELERATORS. L. Peacocke1, P.J.K. Bruce1, and M. Santer1, Department of Aeronautics, Imperial College London, London, SW7 2AZ, United Kingdom.

Abstract: Imperial College London and Airbus Defence and Space are collaborating on a research project to investigate deployable aero-decelerator heatshields for future high-mass payload missions to Mars. Such deployable entry vehicles would extend to diameters much larger than the launch vehicle fairings that take them to space, enabling payloads much larger than the current 1 tonne capability to be delivered to the Martian surface. A 6 degree-of-freedom simulator has been developed within Matlab to assess different entry vehicle diameters, aeroshapes, entry conditions and ballistic coefficients, building on prior work performed at Im- perial. This simulator outputs the attitude variation of the entry vehicle, allowing the study of dynamic stability, of critical importance at lower Mach numbers. In addition, work is ongoing to couple the simulator to the commercial finite element solver Abaqus to allow modelling of the aero-structural behaviour during the entry and descent. The 6-DOF tool will continue to be developed for the duration of the research project, and will be used to analyse and optimize a number of configuration concepts for the deployable entry vehicle. The design and assessment of the configuration options is the next step in the research project, leading on to a high speed wind tunnel test campaign in the later phases to validate the preferred configuration. The simulator consists of the following modules: • Atmosphere model based on the European Mars Climate Database to calculate the freestream properties. • Geometry model with variable dimensions and shapes to generate a mesh for aerodynamic analysis. • Aerodynamic model based on the Modified Newtonian method to determine aerodynamic forces and coefficients. • Preliminary heating model based on the Sut- ton-Graves and Tauber-Sutton relations. • Trajectory model based on a set of equations of motion and a Runge-Kutta numerical solver. This poster will outline the structure of the entry trajectory simulator, the methodologies and candidate geometries used, the validation cases that have been correlated against Airbus Defence and Space in-house results, and early results from the aero-elastic coupling. SPECIAL CHECK-OUT EQUIPMENT FOR ENTRY, DESCENT AND LANDING GNC VERIFICATION AND VALIDATION. E. Rodriguez1, A. Ayuso2, I. de Miguel3, R. Haya4 1SENER Ingeniería y Sistemas, S.A. (Severo Ochoa, 4, 28760 Tres Cantos, SPAIN [email protected]), 2SENER Ingeniería y Sistemas, S.A. (Severo Ochoa, 4, 28760 Tres Cantos, SPAIN [email protected]), 3SENER Ingeniería y Sistemas, S.A. (Severo Ochoa, 4, 28760 Tres Cantos, SPAIN [email protected]), 4SENER Ingeniería y Sistemas, S.A. (Severo Ochoa, 4, 28760 Tres Cantos, SPAIN [email protected]).

Introduction: Guidance Navigation and Control Controller. It coordinates the real time operation (GNC) for Entry, Descent and Landing vehicles re- of all of the components of the GNC SCOE. quires a complex verification and validation campaign. Units EGSE. They provide the Front End Electron- SENER has designed and developed the Special ics (FEE) interfaces to the equipment under test and the CheckOut Equipment (SCOE) used for the multiple simulation and stimulation functions of the respective simulations required during the test campaigns of IXV GNC components. and ExoMars Schiaparelli missions [1], and is currently Real Time Simulator. It implements the environ- developing a similar SCOE for ExoMars 2020. ment and spacecraft dynamics models that are used to The GNC-SCOE supports the GNC validation and coordinate the operation of the equipment simulation is also integrated in the ground support equipment used models so that the spacecraft under test receives credi- for the GNC subsystem verification campaign. It hosts ble, coherent data from all of the sensors during simu- the simulation models for the GNC units and for the lation of a given scenario. The GNC-SCOE solution Dynamics Kinematics and Environment (DKE). Simu- from SENER has demonstrated being capable of exe- lation models used for the GNC validation are adapted cuting any GNC unit model used in the Entry Descent and integrated into a real time simulator, which is exe- and Landing vehicles in real time. Regarding the Envi- cuted it in real time. The GNC-SCOE provides the ronment models, it has demonstrated to perform realis- adequate interfaces to use these models in the GNC tic simulations of atmospheric and aerodynamic mod- verification. The interfaces can be set-up in different els, gravity model and parachute model. For the Dy- ways, allowing the use of the GNC-SCOE at different namics, it included the computation of linear and angu- GNC verification stages. As an example of this flexi- lar components of the status vector based on instanta- bility, the GNC-SCOE can interface directly the On- neous forces and torques applied. Pre-defined perturba- Board Computer (OBC) providing simulation of the tion model can also be included in the simulation, in GNC units, and it can also stimulate electrically the open loop. units when they are in the loop. Hardware architecture: GNC-SCOE Hardware is General Design: The GNC-SCOE design is based based to the maximum possible extent on “commercial on a flexible architecture, which has been successfully off-the-shelf” (COTS) equipment minimizing the cus- adapted to different type of missions. Functions are tom-made hardware development. The main element is summarized in the following figure: a computer with real time operating system, which is

GNC SCOE able both to run the simulator and to access all the FEE interfaces. FEE interfaces implement digital and analog MMI inputs and outputs, specific buses such as MIL-1553, CCS I/F CANbus, reflective memories for high speed interface Controller with other SCOEs and even synchronization signals to Units allow a realistic simulation. EGSE Software architecture: GNC-SCOE software is Real Time Sensors Simulator deployed in two computers, one is a windows computer OBC I/F Actuators GNC Units I/F which hosts the MMI function and other functions not having real time requirements. The other is a real time controller, which hosts the real time simulator. Figure 1. GNC-SCOE Functions Real Time Simulator is structured in models, which

can be compiled generating Dynamic Link Libraries MMI. In addition to the interface with the CCS, (DLL). Updated versions of models can replace older which will allow integration with other SCOEs in the models by generating newer DLLs with any suitable Test Bench, the GNC SCOE can be operated locally compiler such as “Visual Studio” or “GCC”, capable to from its Man-Machine-Interface (MMI). generate Windows DLLs. The simulator models for the units include a math- a manner, that the GNC Software functionality can be ematical model, which performs a behavioural simula- verified. tion of the unit, and also an operational and interface On-board Computer in the loop. In a second stage, model, which simulates the unit modes and the behav- the GNC Software is integrated with the Basic Soft- ior of its interfaces using the dedicated GNC-SCOE ware and the rest of the Application Software functions FEE. to compose the On Board Software (OBSW). The Real Time Simulator GNC Software is thus executed in a representative target processor, so it runs in real time, and uses repre- Sensor Model sentative interfaces. Depending on the hardware avail- Operational Mathematical and Interface ability, Functional Models (FUMO) can also be used. model model FEE The test bench will reproduce as much as possible space flying conditions, allowing high fidelity emula- tion for the GNC. The GNC-SCOE is used in two modes, open loop to test individually all the different DKE interfaces with sensors and actuators and closed loop to verify the GNC functions. In this stage a connection with the other SCOEs is also implemented. Actuator Model

GNC Units and OBC with SW Units in the loop. In a third stage, the GNC will Operational Real Time Scheduler (data exchange) Mathematical provide representative hardware for some of the sen- and Interface model model FEE sors and actuators. In this configuration, the test bench uses the real sensors and/or actuators, for high repre- sentativeness. Two different new tasks come up in this Figure 2. Real Time Simulator scenario: the GNC-SCOE provides stimuli to the hardware sensors; the GNC-SCOE monitors the real hard- In a timed loop with high fidelity, the DKE gets da- ware actuators actuation, measuring it in a manner that ta from the actuator units models, changes the dynamic can be used as an input to the DKE running in the status of the vehicle calculating the corresponding ef- GNC-SCOE. fects to the satellite attitude and position status. As a Tailoring these steps is performed for each project result of this calculation, the simulator provides data to in order to meet the specific V&V requirements. In the the sensor units models. Sensor models will determine frame of Exomars, the GNC team uses the GNC-SCOE the FEE outputs that are generated in time so the On- as a valuable tool for the GNC subsystem verification Board Units and On-Board Computer can acquire in a staggered approach [2]. It is a low cost solution them. Up to 200Hz real time simulation has been suc- and it provides high flexibility in the design process as cessfully executed in the GNC-SCOE. This way the the verification and validation can happen in parallel. behavior of the GNC is verified with sufficient accura- Common aspects with other SCOEs: Sinergies cy. and differences between Entry, Descent and Landing GNC Verification and Validation: The GNC GNC SCOEs and other SCOEs have been identified subsystem validation and verification is an incremental and successfully applied for other mission (e.g. MTG process along the development lifecycles. It starts with and EXM20 Cruise). a local validation using the design & validation envi- Lessons learnt: Among the variety of lesson learnt ronments of each function; afterwards, all the functions during SENER experience in this equipment, it is im- are integrated in a formally verified simulation envi- portant to underline features such as the flexibility add- ronment to perform functional verification. This func- ed by the model replacement oriented architecture and tional verification can be extended embedding the cod- the electrical interface catalog availaibility. ed o/b software. The next step involves real time test- Conclusion: GNC SCOEs for Entry, Descent and ing, where the SCOE is deployed an utilized. It is per- Landing have been successfully designed and devel- formed in three stages: oped by SENER for IXV and ExoMars missions; such Software in the loop. In the first stage, The GNC heritage is being the basis for other SCOEs applied for has only available the GNC Software with the algo- other missions including Cruise phase and AOCS. rithms coded in it. For this purpose, “Matlab/Simulink” References: [1] E. Rodríguez et al. (2012) SCOE is used. In this stage, the GNC-SCOE must provide a for IXV and ExoMars GNC, SESP 2012. software interface that matches the GNC Software in- [2] E. Rodríguez et al. (2010) Integrated Test terface, and simulate the DKE, sensors and actuators in Bench For Rapid GNC Design, Verification And Vali- dation. MOLECULAR SIMULATION OF MATERIALS DURING HYPERSONIC PLANETARY ENTRY. T. E. Schwartzentruber1, A. D. Achambath1, S. Poovathingal1, E. C. Stern1, 1Department of Aerospace Engineering and Mechanics, University of Minnesota, Minneapolis, MN.

Introduction: There are two main classes of mate- of Fiberform is shown in Fig. 2. Specifically, the de- rials used for thermal protection systems (TPS) on hy- gree of in-depth diffusion of reactive oxygen atoms personic vehicles; re-useable ceramics (which form can be quantified (shown in Fig. 2). This is a compet- silica-based oxide coatings), and ablative carbon mate- ing process between diffusion of oxygen atoms from rials. For re-useable ceramics, the surface remains mi- the high-temperature boundary layer into the material, croscopically flat and chemical reactions occur at de- surface reactions which deplete O to form CO and fect sites. For ablative materials, complex changes in CO2, and the diffusion of these reaction products back microstructure occur that have a significant impact on into the boundary layer. Results will also be presented the overall gas surface interaction. It is now possible to where a pyrolysis gas is out-gassing into the boundary use x-ray micro-tomography to image a real TPS mate- layer, further affecting the diffusion of O atoms into rial in 3D, triangulate a surface geometry, and directly the material. It is ultimately surface reactions involving simulate shock-heated gases interacting with the mate- O atoms that degrade/ablate the fiber structures, so rial microstructure [1-3]. determination of the amount of in-depth diffusion of O atoms is the goal of such simulations. Reaction rates between atoms and molecules in the shock-heated gas and the carbon surface are based on recent molecular beam experimental results [4].

Figure 2: Molecular simulation of atomic oxygen dif- fusing into, and reacting with, a Fiberform sample. Contours show the depth of oxygen diffusion.

Acknowledgments: This research was supported by AFOSR grant FA9550-10-1-0563. E. C. Stern was Figure 1: SEM image of porous fiber-based carbon supported by a NASA NSTRF grant. S. Poovathingal material (top) and 3D scan using X-ray micro- was partially supported by a Doctoral Dissertation Fel- tomography (bottom). lowship from the University of Minnesota.

Results: Results will be presented where the References: [1] Poovathingal, et al. (2016), AIAA DSMC particle method (the MGDS code developed at Journal, Vol. 54, No. 3. [2] Stern et al. (2015), AIAA the University of Minnesota) is used to simulate react- Paper 2015-1450. [3] Stern et al. (2014), AIAA Paper ing flow through porous fiber-based carbon TPS mi- 2014-2247. [4] Murray et al. (2015), J. Phys. Chem. C, crostructure obtained from tomography (Fig. 1). An Vol. 119, No. 26. example DSMC simulation of flow through a sample Model Development of Parachute Dynamics During Planetary Descent. J. D. Sparta1†, G. A. Rossman2†, B. A. Sforzo3†, and K. A. Hart4†, 1Undergraduate Research Assistant, [email protected], 2Graduate Research Assis- tant, [email protected], 3Postdoctoral Fellow, [email protected], 4Graduate Research Assistant, [email protected], †Space Systems Design Lab, Daniel Guggenheim School of Aerospace Engineering, Georgia Institute of Technology, 270 Ferst Drive, Atlanta, GA 30331.

Introduction: Parachutes used during planetary cients to the state variables of the parachute using hid- descent, especially at Mars, experience vigorous and den layers of mathematical feed forward relationships. complex aerodynamics that lead to randomness in the These relationships were developed by training the dynamics of the descent vehicle that are not well un- neural network on the dynamic data collected in the derstood and present a challenge for modeling. Flight test. The neural network model’s predictive capability dynamics simulations for current EDL systems with was greatly improved by adding acceleration terms to parachutes either do not include models of dynamic the angular position and angular rates that are the state aerodynamics at all or use highly simplified models variables upon which the model was trained. such as in the Mars Science Laboratory mission [1]. A In order for the developed models to be validated, more sophisticated model is desired to better predict they had to be implemented in a simulation of the wind the dynamics for future missions involving parachutes. tunnel test and have their resulting dynamics compared The creation of a database of the dynamics of para- to those observed during the test. chutes from which dynamic aerodynamic models could A simulation of the wind tunnel test was construct- be developed was defined as a principal goal of ed in MATLAB which recreates the wind tunnel test NASA’s recent wind tunnel tests of Disk-Gap-Band conditions and implements the parachute’s equations of (DGB) and Supersonic Ringsail (SSRS) type para- motion in the wind tunnel. Several analyses were per- chutes [2]. formed to verify the physics of the simulation before Before models could be built from the dynamic da- implementing the new aerodynamic models. tabase created during the wind tunnel test, a data pre- Validating the new aerodynamic models in the sim- processing trade study was performed to determine the ulation has proved to be a challenging endeavor. Since best criteria for smoothing the experimental data. A real parachute dynamics can exhibit randomness, it is methodology was finalized for filtering the data, re- expected that the incorporation of these dynamic mod- moving erroneous measurements, and filling gaps in els should also be able to introduce similar behavior the data collection, allowing for smooth, differentiable into the simulation. This means that the models can be data from which models could be built. numerically sensitive, and therefore if they have imper- Model Development: Two approaches at model- fections in them, the models can become numerically ling the dynamics observed in the wind tunnel test were unstable. This is what has so far been observed in undertaken. The first approach assumes that the dy- simulations including the newly developed models, namics of parachutes can be determined with current where the parachute has experienced rapid dynamically understanding of the forms of the models. An apparent unstable motion until the simulation fails to solve. Fu- inertia model developed by Eaton was adopted to mod- ture investigations will be necessary to understand the el the dynamic aerodynamics [3]. This model accounts behavior of these models in the simulation. for the dynamic effects of the air moving in and around In the assessment of the fidelity of the aerodynamic the parachute canopy. A technique known as parameter models, it is unrealistic to expect the models to predict identification was used to extract the unknown inertia the parachutes state variables vs. time exactly. There is coefficients for the parachute system directly from the too much randomness in the dynamics of parachutes to dynamic data collected in the test. This approach was produce an exact simulation. Metrics for statistical able to produce aerodynamic coefficient predictions comparison between the dynamics of the simulation that were in good agreement with those calculated, for and the wind tunnel test are laid out as a basis for mod- given input states of the parachute and tunnel. el validation. It is possible that the current understanding of the forms of the models is not adequate for modelling the Acknowledgements: This work has been funded dynamics. That is, perhaps there are underlying physics through JPL subcontract number 1548418 under driving the dyamics that are not understood at this NASA contract NNN12AA01C. The authors would time. A second approach utilizing system identification also like to thank the collaborators at JPL; Ian Clark techniques was investigated in the case that this as- and Clara O’Farrell, and NASA; Juan Cruz and Justin sumption should be the truth. A neural network model Green, for their guidance, as well as provision of was implemented which maps the aerodynamic coeffi- experimental data, software tools, and analyis formulations.

References: [1] Cruz, J. R., et al. (2013). “Parachute Models Used in the Mars Science Laboratory Entry, Descent, and Landing Simulation,” AIAA Aerodynam- ic Decelerator Systems (ADS) Conference, Aerody- namic Decelerator Systems Technology Conferences, (AIAA 2013-1276). [2] Zumwalt C. H., et al., (2016). "Wind Tunnel Test of Subscale Ringsail and Disk-Gap-Band Para- chutes", 34th AIAA Applied Aerodynamics Confer- ence, AIAA AVIATION Forum, (AIAA 2016-3882). [3] Eaton, J.A., "Validation of a Computer Model of a Parachute," Ph. D. Dissertation, University of Leicester, Leicester, UK, 1982.

EXACT ANALYTIC CALCULATION OF AERODYNAMIC COEFFICIENTS IN THE HYPERSONIC REGIME AND SHAPE OPTIMIZATION FOR ENTRY VEHICLES P. J. S. Gil1 and A. Ferrero2 1IDMEC, Instituto Superior Técnico, Universidade de Lisboa, Av. Rovisco Pais, 1049-001 Lisboa, Portugal, email: [email protected] 2DIMEAS, Politecnico di Torino, Corso Duca degli Abruzzi, 24, 10129 Torino, Italia, email: [email protected]

Abstract: In this work, the aerodynamic coeffi- References cients of simple shapes are determined analytically for [1] Grant, M.J. and Braun, R.D. (2010) 48th AIAA the case of hypersonic flight conditions, where the Aerospace Sciences Meeting, Paper AIAA 2010-1212. Newtonian impact theory can be applied. The shapes [2] Grant, M.J. and Braun, R.D. (2012) 2012 AIAA are defined by only a few parameters but are sufficient- Atmospheric Flight Mechanics Conference, Paper ly interesting to be used as a good approximation of AIAA 2012-4580. real cases such as rocket noses or entry capsules. [3] Grant, M.J. and Braun, R.D. (2015) J. Space- The exact analytic expressions obtained for the aer- craf and Rockets, 52, 177-182. odynamic coefficients enable the possibility of an optimization procedure of the shape. Each expression is dependent from the characteristic geometrical parameters of each shape. The same parameters are then used in the optimization process as the variables to be optimized. Examples are shown where possible constraints are defined and the optimal geometric configurations are obtained for a minimal value of the drag, in the case of a launcher, or a minimal value of the ballistic coefficient, in the case of an entry vehicle. Although the analytical solutions and the optimization procedures are limited to the hypersonic regime and simple forms, they can be useful in the context of preliminary design and multidisciplinary design optimization, where speed can be a more important factor than accuracy, or as a benchmark for software testing. Our work was based on the work of Grant and Braun [1-3], which we were able to extend, by obtaining the exact analytic solutions for several simple but nevertheless useful shapes.

Figure: Example of a blunted bi-conic nose before and after shape optimization that minimizes the ballistic coefficient at constant internal volume. NUMERICAL MODELLING OF THE RECESSION OF AN ABLATOR TPS WITH AN EMBEDDED “ReGS” SENSOR. P. Georgiopoulos, G. Vekinis and A. Marinou, Insitute of Nanoscience and Nanotechnology, NCSR Demokritos, Agia Paraskevi Attikis, 15341, Greece, [email protected]

Introduction: The experimental development of was assumed to be incident uniformly throughout the the “ReGS” TPS recession sensor (discussed elsewhere surface and the flux was assumed to reach maximum in this conference) has been aided by numerical model- gradually as the top layer of elements burns away and ling of the recession that occurs when an ablator cou- a new layer takes its place. Element recession is pon containing a sensor is heated by a continuous achieved by “killing” the top layer of elements when thermal flux. The ANSYS package (thermal) has been they reach pyrolysis (start of ablation) temperature by utilised for this purpose and main aim of the this work suddenly increasing their thermal conductivity 100- was to develop and calibrate a simple model to predict fold enabling heat flux to reach the next layer. The recession behaviour of the sensor in an ablator under surface temperature of the ZURAM during ablation different incident heat fluxes. (about 1700oC) was determined by a pyrometer during The ReGS sensor: The “ReGS” (“Resistive Grid testing. TPS recession Sensor”) sensor is currently being de- The model was calibrated by carrying out ablation veloped with ESA funding [1] on the basis of the pre- tests at constant heat flux (approximately 3MW/m2 at a viously reported “ReWiG” sensor [2] and has been constant oxy-acetyelene flame-specimen distance) and used to monitor recession in the ZURAM [3] ablator as measuring the actual recession post-test. well as other ablating materials. Modelling results: The experimental and model- The ReGS sensor consists of a thin metallic grid ling results for the recession of the ZURAM ablator are whose resistance increases almost linearly for the first compared in Figure 2 where agreement can be seen to half of its length as the grid elements melt away when be very good. embedded in the ablator, as shown in Figure 1.

Figure 1. A recessed ReGS sensor in ZURAM after flame

ablation. Note the deeper recession around the sensor. Figure 2. Comparison of transcient numerical model of ZURAM recession with experiment. Modelling: The work was carried out using ANSYS Thermal with all materials modelled with the The model is able to predict the recession of PLANE55 (2D) element. Transcient calculations were ZURAM with a ReGS sensor, see Figures 3 and 4. carried out in all cases. An ablator disc containing a ReGS sensor (modelled as a simple thin metallic foil) in the centre coated with a silicate glue was used as the model system. A total of 77760 elements were used for the model with a finer mesh around the sensor and glue. The material properties (thermal heat capacity and thermal conductivity as a function of temperature, phenolic and carbon pyrolysis and melting point of the sensor foil and glue) were either obtained from literature or determined experimentally. Because of pyrolysis of the ablator, the actual heat flux experienced by the surface is a function of the temperature. Heating Figure 3. Modelling recession of Zuram with a ReGS [1] “ReGS: a Resistive Grid TPS recession Sensor”, ESA/ITI sensor (right) contract B000016985, 2016 The results in Figure 3 agree with the experimental [2] G. Vekinis, “ReWiG: a resistive grid TPS recession sen- observations (Figure 1) that inclusion of the sensor sor”, IPPW12, 15-19 June 2015, Cologne, Germany (thin vertical black line between the protective glue [3] A. S. Pagan, J. Rieser, B. MassutiBallester, G. Herdrich, “Characterisation of the Lightweight Ablative Heat Shield layers on the right in Figure 3 shows the recession of th the sensor) results in non-uniformity of heating and Material ZURAM® in HighEnthalpy Air Flows”, 8 Euro- pean Workshop on TPS and Hot Structures, ESA/ESTEC, 19-22 April 2015, The Netherlands.

Figure 4. Modelled recession as a function of distance from sensor and time for a 3MW/m2 heat flux

heavier recession close to the sensor (actuall y about 1.4mm from it). This is shown clearly in the results in Figure 4 which also indicates that the sensor correctly measures the maximum recession of the Zuram at all times. Further comparison between model and experimental recession sensor data is shown in Figire 5 where the model results (at 1.4mm from sensor) for two heat flux levels is compared with sensor data .

Figure 5. Comparison of sensor recession data with model results for two flux levels.

The first-order numerical model developed in this work has been able – after due calibration – to correctly simulate the actual experimental observations (ablation by oxy-acetylene flame burning), both in terms of temperature profiles and in terms of recession.

The work is continuing.

References: ADVANCES IN THE FIELD OF HIGH PERFORMANCE PARALLEL COMPUTING AND THE IMPACT ON SCIENTIFIC DISCOVERY. M. E. Pizzo1 and D. P. Hammond2, 1Old Dominion University (2300 Engineer- ing and Computational Sciences Building, Norfolk, VA 23529, [email protected]), 2NASA Langley Research Center (NASA Langley Research Center, Mail Stop 128, Hampton, VA 23666, [email protected]).

Introduction: The evolution of high performance academia, and industry, “[the] simulations are compu- computing (HPC) has seen several paradigm shifts tationally expensive, and even with optimized codes, since the introduction of supercomputing in the 1960s. HPC resources are critical [5].” Original supercomputers were driven by the total float- Even though the ESM Project already utilizes ing point operations per second (FLOPS) a machine HPC technologies, many researchers in the field still could deliver, and carried a large operational cost. By underutilize HPC. It is important to generate conversa- the 1970s and 1980s, the number of processors in su- tions within the scientific community to discuss key percomputers increased which shifted HPC from sym- concepts within HPC as the field moves towards in- metric multiprocessing systems to low cost compute creased levels of parallelism and the use of hybrid, clusters [1]. This shift followed the observation known heterogeneous architectures and co-designed systems. as Moore’s Law, which states that the number of tran- The objective of this presentation is to generate im- sistors in a dense integrated circuit doubles approxi- portant discussions necessary for advancing HPC tech- mately every two years. HPC has since continued to nologies within the larger EDL community to better evolve moving from systems with a relatively low impact scientific discovery. Key concepts that will be number of fast processing cores (multi-core) to ones discussed include (but are not limited to) the: with a vast number of cores and an increased level of 1. Differences between parallel architectures and parallelism (many-core). This extension of Moore’s programming models including both shared Law “transformed computing from a rare and expen- memory (OpenMP, OpenACC) and distributed sive venture into a pervasive and affordable necessity memory (MPI), [2].” 2. Scalability of programming models, Today’s leadership class machines can deliver pet- 3. Different capabilities of compute performance ascale (10e15 FLOPS) performance, with the potential and algorithm acceleration available through of reaching exascale (10e18 FLOPS) performance as CPU, GPU, Xeon Phi, and FPGA hardware in- early as year 2023 [3]. The effective use of exascale terfaces, and architectures will require a redesign in modeling, algo- 4. Application to EDL systems. rithms, and system components. This hardware, system References: [1] Eadline, D. (24/08/16). The software, and application co-design process [1] is driv- Evolution of HPC. Inside HPC. Web (accessed on en by the growing scale and complexity of computer 22/02/17). insidehpc.com/16/08/the-evolution-of-hpc. architectures and science applications; increased levels [2] 50 Years of Moore's Law. Intel© Corporation. Web of data production, storage, and consumption; sustain- (accessed on 15/02/17). www.intel.com/content/www/ ability of software; and training for the next generation us/en/silicon-innovations/moores-law-technology.html. of HPC systems. [3] Messina, P. and Lee, S. (20/12/16). Update on the Objective: Entry, descent, and landing (EDL) sys- Exascale Computing Project. American Geophysical tems benefit from HPC and will be impacted by the Union, ASCAC Meeting, Washington, D.C. Web impending exascale programming models and architec- (accessed on 27/02/17). science.energy.gov/~/media/ tural changes. Predictive modeling for NASA EDL ascr/ascac/pdf/meetings/201612/ECP_Update_ASCAC missions, for example, was most recently demonstrated __Dec2016.pdf. [4] Wright, M. (10/11/16). Predictive to audiences in attendance at Supercomputing 2016. As Modeling for NASA Entry, Descent, and Landing stated by Michael Wright, Senior Research Scientist at Missions. NASA Advanced Supercomputing Division. NASA Ames Research Center, “The ultimate goal of Web (accessed on 17/02/17). www.nas.nasa.gov/ NASA's Exploration Missions is to safely land humans SC16/backgrounder_wright_predictive_modeling.html. on the surface of Mars and return them safely to Earth. [5] Wright, M. (10/11/16). Predictive Modeling for This may well be our most challenging achievement as NASA Entry, Descent, and Landing Missions. NASA a nation – and it will not happen until we are able to Advanced Supercomputing Division. Web (accessed accurately simulate all of the complex steps involved on 17/02/17). www.nas.nasa.gov/SC16/demos/demo19. and demonstrate that our technology is up to the task html. [4].” Established by the Entry Systems Modeling (ESM) Project through partnerships within NASA, Blackout Analysis of Reentry Vehicles for Martian Missions. S. Ramjatan1, T. E. Magin1, J. Thoemel2, A. Lani1, J. B. Scoggins1, A. Turchi1, G. Chatzigeordis1, S. Boccelli1, B. Dias1, O. Karatekin3, and B. Hoeve1 1von Kármán Institute for Fluid Dynamics, 1640 Rhode-Saint-Genese, Belgium 2Royal Belgian Institute for Space Aeronomy, Avenue Circulaire 3, 1180 Brussels/Uccle, Belgium 3Royal Observatory of Belgium, Ringlaan 3, Brussels/Uccle 1180, Belgium

Introduction: Atmospheric entry of spacecrafts is planetary missions should have a more robust commu- one of the most challenging aspects of the design phase nication system including a blackout communication for space missions. Due to Martian atmospheric entry capability [3]. However, blackout was also recently ex- flight conditions, high velocities on the order of 5-8 perienced on ESA’s Schiaparelli module during its en- km/s imply strong shock waves and extreme aerody- try into Mars and therefore an improved blackout mod- namic heating of the vehicle. High temperatures lead to eling approach and prediction will benefit upcoming an increase in collisions of the molecules, which result Martian missions. Results were obtained on blackout in the disruption of the electronic structure producing analysis at the Von Karman Institute for Fluid Dynam- free electrons and ions. This production of electrons and ics (VKI) and involved using the plasma cut-off fre- ions creates a region of ionized flow or plasma around quency to assess the blackout period for small cone- the vehicle that causes communication blackout by at- shaped reentry vehicles [4]. tenuating radio frequency signals from the spacecraft, Approach: The current paper performs a compre- and transmission is possible only at RF frequencies hensive modeling of blackout for Martian missions by greater than the characteristic plasma frequency. During accounting for the plasma frequency and collision fre- blackout, vehicles cannot receive guidance or maneu- quency, and examines their effect on a ray tracing vering information from a control center and can travel model. This paper applies state of the art tools for pre- for hundreds of miles during this period of loss of com- dicting blackout duration on previous Martian space munication. The blackout period interferes with normal missions that have flight data readily available in the lit- performance of data telemetry, monitoring and com- erature. Blackout predictions are compared with flight mand, and control systems operating between the space- data to examine the difference between the predicted craft and the ground. As future Mars exploration mis- and actual blackout period. Maps of the plasma distri- sions require increased landing performance (higher bution and collision frequencies around the reentry ve- mass, smaller landing ellipse, higher elevation) the en- hicle are generated by non-equilibrium hypersonic CFD try, descent, and landing (EDL) phase in the Martian at- simulations and subsequently used to examine electro- mosphere will become increasingly challenging [1]. As magnetic wave interactions with the plasma using Ap- a result, radio blackout is a potential mission killer since pleton theory [5]. A ray-tracing program is implemented loss of communication with the spacecraft results in to compute the trajectory of the radio wave until it exits high risk to crew safety or mission success. the plasma allowing for a more realistic estimate of the With future space missions aimed towards Mars in- blackout phase [6]. cluding NASA’s Mars 2020 and ESA’s ExoMars 2020, Results: Blackout predictions are made for a Mar- there is a motivation and need to improve prediction of tian reentry at different flight conditions by applying ac- the blackout duration period. As evident, Morabito [2] curate CFD and thermodynamic modeling. This is examined the 30 second blackout of the X-Band antenna achieved by the coupling of the VKI’s MUTATION++ of the Mars Pathfinder mission by comparing the elec- library for reacting and plasma flows to the VKI’s tron density to the critical density of the spacecraft’s fre- COOLFluiD CFD code [7], [8]. A ray-tracing model is quency band. He found that the electron density in the applied for the first time for a Martian reentry. In addi- wake region exceeded the critical X-band electron den- tion, electron-heavy particle collisions are accounted for sity during the first 20 seconds of the 30-second black- resulting in an improved understanding of the physics out period. He explained that the difference was likely associated with the blackout phenomena. Significantly, due to High Doppler dynamics and low signal to noise the results acquired in this paper assists in designing the ratio, which may have contributed to the blackout dur- EDL phase for upcoming Martian missions and results ing the last 10 seconds. Thus, current blackout predic- in a greater understanding of the Martian thermo-chem- tion methods must be improved to accurately predict the ical modelling. loss of communication. Significantly, the European Space Agency’s Beagle 2 spacecraft was launched in 2003 on the Mars Express Spacecraft where communication was lost on the day it was supposed to land. Out of the recommendations, it was suggested that future References:

[1] M. J. Wright, C. Y. Tang, K. T. Edquist, B. R. Hollis, P. Krasa and C. A. Campbell, "A Review of Aerothermal Modeling for Mars Entry Missions," in 48th AIAA Aerospace Sciences Meeting Including the New Horizons Forum and Aerospace Exposition, 2010. [2] D. Morabito, "The Spacecraft Communications Blackout Problem Encountered During Passage or Entry of Planetary Atmospheres," Jet Propulsion Laboratory, IPN Progress Report 42- 150, Pasadena, 2002. [3] F. Bonacina, "Lessons Learnt from Beagle 2 and Plans to Implement Recommendations from the Commission of Inquiry," Ireland, 24 May 2004. [Online]. Available: http://www.esa.int/ESA_in_your_country/Irelan d/Lessons_learnt_from_Beagle_2_and_plans_to _implement_recommendations_from_the_Com mission_of_Inquiry. [Accessed 26 January 2014]. [4] S. Ramjatan, T. Magin, V. Van der Haegen and J. Thoemel, "Blackout Analysis of Small Cone- Shaped Reentry Vehicles," AIAA Journal of Thermophysics and Heat Transfer, http://dx.doi.org/10.2514/1.T4825, [Accepted for Publication]. [5] K. Davies, Ionospheric radio propagation, Washington: National Bureau of Standards, 1965. [6] A. Delfino, Modeling of the Antenna Radiation Pattern of a Re-entry Vehicle in the Presence of Plasma, Chicago: University of Illinois, 2004. [7] von Karman Institute for Fluid Dynamics , "Mutation++," von Karman Institute, [Online]. Available: https://www.mutationpp.org/docs/html/. [8] A. Lani, K. Bensassi, L. Kapa, M. Panesi and M. Yalim, "COOLFluiD: an open computational platform for multi-physics simulation," in 21st AIAA CFD Conference [AIAA 2013-2589], San Diego, 2013.

Parachute-Induced Wrist Mode Dynamics During Martian EDL. I. G. Clark, NASA Jet Propulsion Laboratory, California Institute of Technology, 4800 Oak Grove Dr., Pasadena, CA 91109, [email protected]

Introduction: The rotation of a body about its center of gravity while under descent on a parachute is often referred to as the wrist mode dynamics of the body. Concerns about the attitude and attitude rates associated with wrist mode dynamics often arise for planetary landers. For example, the Viking mission had stringent requirements to keep attitude rates to a peak of less than 100 deg/sec around the time of parachute inflation and to less than 30 deg/sec six seconds after mortar deployment [1]. Testing during Viking’s the Balloon Launched Decelerator Test series revealed rates as high as 148 deg/sec around peak inflation loads. However, a noted correlation of rate to peak load on the Viking capsule, in conjunction with an expected increase in the inflation time at Mars, helped convince Viking engineers that rates below 100 deg/sec would be achieved at Mars [1]. Later missions to Mars sought to reduce capsule dy- namcis due to the parachute by modifying the Viking parachute configuration to lengthen the band. This was done more to increase the stability of the parachute and limit the inflated parachute/capsule system dynamics, as opposed to a true wrist mode behavior. However, during the descent of both of the Mars Exploration Rover missions, an excitation in attitude rates was observed after the peak parachute inflation rates had appeared to dampen down. This led to an initial investigation into the nature and physics of capsule wrist mode while on parachute descent. Simple equations were developed that help provide insights into the correlation of attitude rates and attitude magnitude and how they may change in a decelerating system. More recently, concerns about capsule attitude rates after parachute deploy for the M2020 mission have led JPL to return to the prior wrist mode studies performed after MER and further the work. This presentation presents a smmary of the findings from that effort. Simple relations that cover capsule dynamics during inflation were developed to identify sensitivities to capsule size and mass properties, parachute attachment configuration, parachute inflation parameter, and initial conditions at the time of parachute inflation. A description of mechanisms that can exacerbate attitude rates after peak parachute loads is also provided in an effort to help bound some of the post-inflation dynamics that may occur. References: [1] Moog, R. D. and Michel, F. C., “Balloon Launched Viking Decelerator Test Program Summary Report,” NASA Contractor Report 112288, Mar. 1973. PROGRESS ON FREE-FLIGHT CFD SIMULATION FOR ENTRY CAPSULES IN THE SUPERSONIC REGIME. E. C. Stern1, J. M. Brock2, Alan M. Schwing3, Mark Schoenenberger4, Michael C. Wilder1, Brandon Smith1, and J. D. Hergert5, 1NASA Ames Research Center ([email protected]), 2Analytical Mechanics Associ- ates, Inc., 3NASA Johnson Space Center, 4NASA Langley Research Center, 5Stanford University

NASA, through the Entry System Modeling (ESM) project, has been developing a capability to simulate the super/transonic dynamics of planetary entry probes during descent using high-fidelity CFD [1]. At IPPW13 prelminary results from a study of a supersonic inflatable aerodynamic decelerator (SIAD) which showed promise in accurately predicting the dynamics observed in a series of ballistic range experiment. This presentation will highlight progress on maturing this technology over the past year, focusing on the three applications First, analysis of the SIAD ballistic range experiments has been completed. The methodology for ini- tializing a simulation for a free-flight ballistic range experiment has been refined, allowing for good agreement between experimental and simulated results for the total angle-of-attack, as seen in Fig. 2. The level Figure 1: Total angle of attack comparison for a agreement shown is consistent across the range of SIAD geometry compared with ballisitics range Mach numbers used in the experimentAdditionally, data. Free-flight CFD is the blue line, and experi- static and aerodynamic coefficients have been comput- ment is the red symbols. ed based on the CFD simulations, and also show good agreement with those derived from experiment. In addition, the free-flight CFD technology has been applied to the analysis of a recent ballisitic range campaign [3,4] in support of the Mars2020 mission. This experiment featured an instrumented model which has yielded unique data which can be used in the validation of the CFD tools. For the current work, we shall show comparisons between CFD predictied backshell pressures, and those measured in the experiment. Finally, an eventual goal of this effort is to be able to simulate atmospheric tractories at flight scale. De- velopments to this end are demonstrated by simulating segments from the upcoming ADEPT sounding rocket Figure 2: Flow visualization from a simulation of test. Using these demonstration, we shall comment on the Mars2020 ballistics range model. the feasibility of utilizing this technology to directly simulate capsule trajectories using high-fidelity CFD. [1] Stern E. C. et al. (2012) AIAA Appl. Aerody- namics Conf. [2] Brock J. M. and Stern E. C. (2016) IPPW, [3] Schoenenberger M. et al. (2016) IPPW [4] Stern E. C. et al. (2016) AIAA Appl. Aerodynamics Conf. Finite Rate Catalytic Models of Silicon Carbide based High-Temperature Ceramics. B. Massuti-Ballester and G. Herdrich, Institute of Space Systems, University of Stuttgart, Pfaffenwaldring 29, 70569, Stuttgart, Germany, [email protected].

Abstract: Planetary entries accompanioned with moderate heat loads facilitate the use of high temperature ceramics as part of re-usable Thermal Protection Systems (TPS). Thermochemical properties of silicon carbide (SiC) based ceramics have been widely investigated in flight and at ground test facilities, specifically the gas-surface interactions under higly dissociated flows [1, 2, 3, 4]. In the presence of atomic oxygen SiC oxidises forming a stable SiO2 glass layer associated to -quartz or -Cristobalite depending on the temperature regime. Catalytic activity of SiO2 surfaces has been investigated at many different facilities and conditions with a similar outcome, which is a reduced recombination coefficient . Available catalytic models for SiO2 are tipically based on global models where only depends on surface temperature leading to strong disagreements between studies. In this work, existing data in literature have been use to generate finite rate catalytic models accounting for both surface temperature and atomic partial pressure dependencies of . The model is based on two recombination type reactions following Eley-Rideal and Langmuir- Hinshelwood mechanisms and a simplified surface coverage computation, valid for low catalytic materials, in which is obtained from the reaction rate balance between adsorption and desorption mechanisms.

References: [1] D. A. Stewart, J. V. Rakich and M. J. Lanfranco, Catalytic Surface Effects Experiment on Space Shuttle, 1981. [2] A. Kolesnikov, M. Yakushin, S. Vasil’evskii, I. Pershin and A. Gordeev, Catalysis Heat Effects On Quartz Surfaces in High-Enthalpy Subsonic Oxygen and Carbon Dioxide Flows, Proc. 3rd European Symposium on Aerothermodynamics for Space Vehicles, Nov. 1998, Noordwijk, The Nether- lands. [3] F. Panerai and O. Chazot, Characterization of gas/surface interactions for ceramic matrix composites in high enthalpy, low pressure air flow, Materials Chemistry and Physics 134 (2012) 597-607. [4] B. Massuti-Ballester, S. Pidan, G. Herdrich and M. Fertig, Recent catalysis measurements at IRS (2015), in: Ad- vances in Space Research, 56:4(742--765). FLIGHT MECHANICS MODELING AND PREDICTIONS FOR THE ADEPT SR-1 MISSION. S. Dutta1, A.M. Korzun1, C. D. Karlgaard2, and B.P. Smith3, 1NASA Langley Research Center, M/S 489, Hampton, VA 23681 (corresponding author: [email protected]), 2Analytical Mechanics & Associates, M/S 489, Hampton, VA 23681, 3NASA Ames Research Center M/S 229-1, Moffett Field, CA 94035.

Abstract: The first flight test of the Adaptable current specifications of the on-board sensors. The Deployable Entry and Placement Technology (ADEPT) modeling and simulation tools have allowed the design is called Sounding Rocket One (SR-1) and is currently team to understand the end-to-end performance of the scheduled to take place in late 2017 [1]. The main vehicle and aid in critical design choices. objectives of the flight test are to demonstrate that the vehicle can be deployed exo-atmospherically and to Acknowledgments: characterize the stability of the vehicle in atmospheric The authors of this paper have greatly benefited the flight. The ADEPT vehicle used for SR-1 will be a 0.7 work of the rest of the ADEPT SR-1 team. The results m diameter 70 degree half-angle sphere-cone, which in this paper have been aided by their work. will be launched as the primary payload on an UP aerospace launch vehicle from the White Sands Missile References: Range (WSMR). [1] Wercinski et al., “ADEPT Sounding Rocket One In the build-up to the flight test, several activities (SR-1) Flight Experiment Design Summary,” IPPW14 and analyses have taken place to predict the flight (2017). performance. Free-flight ground testing has been [2] Smith et al., “Free Flight Ground Testing of conducted to characterize a priori the dynamic stability ADEPT in Advance of the Sounding Rocket One Flight of the vehicle at supersonic and subsonic conditions [2]. Experiment,” IPPW14 (2017). Additionally, drop tests and deployment tests have been conducted to quantify through experiment the vehicle’s response during certain dynamic phases of flight. However, in lieu of the actual flight, end-to-end analysis of the flight performance from ADEPT separation from the launch vehicle to touchdown have been done through six degree of freedom flight dynamics modeling and simulation tools. This paper will focus on the development of these modeling tools and how ground test data have been incorporated to improve the modeling fidelity. The current simulation starts with separation states estimated by UP Aerospace, models ADEPT separation from the launch vehicle and its ensuing pitch and yaw rates based on past UP Aerospace flight data, simulates deployment of the ADEPT shape from its stowed configuration based on ground test data, and finally models atmospheric entry and touchdown within the boundary of WSMR. The aerodatabase for the simulations transverse several different regimes, including free-molecular, supersonic continuum, and subsonic continuum, and have also incorporated information from ground test data, such as from the ballistic range tests for IRVE-3 and ADEPT. The flight dynamics modeling and simulation tools have been used to see how the current vehicle design meets requirements, predict flight performance, and characterize critical range safety metrics. Finally, another consideration during the end-to-end modeling and simulation has been to characterize the reconstruction accuracy of metrics of interest based on Quantitative guidelines on radiation model selection for the material response of ablative heat shields. V. Leroy1, J. R. Lachaud2 and T. E. Magin1, 1von Karman Institute for Fluid Dynamics, Sint-Genesius-Rode, Belgium ([email protected]; [email protected]), 2C la Vie, Nouméa, New Caledonia (jean.lachaud@c-la- vie.org).

Introduction: The high heat load to which thermal shows that the radiative power field yielded by the protection systems are submitted during the atmos- Fourier model becomes inaccurate for a significant part pheric entry of a spacecraft leads to a drastic tempera- of the simulation. However, the global effect on the ture increase in the heat shield. In porous ablative heat temperature field remains limited. A simulation run on shields, radiation transfer inside the protecting material the extended version of the Ablation Test Case #2 reaches levels equaling that of conduction or convec- leads to similar observations. tion heat transfer at about 800 K. Accounting for radia- Discussion: Although the structure of the radiation tion transfer with good accuracy therefore becomes field is qualitatively largely impacted by the change of essential for most applications. While most ablation radiation transfer model, the quantitative global effect simulation codes make use of a radiative conductivity is limited due to the temperature levels involved. In- model, the relevance of such an approach has not been deed, when the radiative power field yielded by the thoroughly questioned. This work aims at providing RTE significantly differs from that yielded by the radi- guidelines for the selection of a radiation model in ative Fourier law, radiation transfer has a low intensity material response codes, building on recent advances compared to conduction heat transfer due to the low in the modeling of radiation transfer in porous media. temperature levels. When the temperature levels be- Model selection criterion: Porous materials used come high enough for radiation transfer to have a sig- for the design of ablative thermal protection systems nificant intensity, the temperature field structure is are semitransparent scattering media. A study by such that the radiative Fourier model is relevant in Gomart and Taine [1] quantified the error brought by most parts of the sample. However, the radiative Fou- the use of a radiative conductivity model rather than a rier law remains inaccurate in the near-interface re- model based on the radiation transfer equation (RTE) gion, with effects difficult to quantify due to the nature and provided a validity criterion for the radiative Fou- of the models currently in use for the porous-flow in- rier law based on the temperature field in the material. terface and the near-interface region This criterion depends on the required accuracy on Conclusion: From these observations, it appears radiation transfer and on the material properties. that a radiative conductivity model is quantitatively Numerical methods: In this study, a comparison appropriate for the simulation of material response on of simulated material response using two radiation the considered test cases and in the current state of the models is made to quantify the impact of the selection art for transport in ablative materials. The qualitative of the radiation model on the global behavior of the disagreement in the near-interface region however sample. A radiation coupling library was added to the leads to conclude that significant modeling efforts are PATO material response simulation code [2], allowing still required; for that region, advanced radiation mod- radiation transfer simulation using either a radiative els are required. Fourier law or a fully reciprocal Monte Carlo ray trac- References: [1] Gomart, H. and Taine, J. (2011) ing RTE solver. The Ablation Test Case #2 [3] specifi- Phys. Rev. E, 83(2), 1–8. [2] Lachaud, J. R. and cation was extended to include dedicated radiation Mansour, N. N. (2014). J. Therm. Heat Transf., 28(2), modeling. Material radiative properties were estimated 191–202. [3] Lachaud, J. R., Martin, A., van Eekelen, from the theoretical benchmarking material TACOT T. and Cozmuta, I. (2012). 5th Ablation Workshop. used in the Ablation Test Case series. The RTE solver, Lexington, Kentucky. written in C++ and Python with a modern architecture for easy extensibility and maintenance, was thoroughly tested on a variety of test cases. The implementation of energy partitioning and a fully reciprocal formulation contribute to reducing the variance. Reference simulations are run using PATO coupled to the Monte Carlo solver, and the results are compared with simulations performed using a radiative conductivity model. Results: A purely conducting-radiating benchmark test case is used to demonstrate the capabilities of the solver. The detailed analysis of the temperature profile MODELING CRYOSPHERIC CONTACT AND PHASE-CHANGE PROCESSES FOR APPLICATIONS IN PLANETARY EXPLORATION, J. Kowalski1, K. Schüller, and A.G. Zimmerman, 1Aachen Institute for Advanced Study in Computational Engineering Science, RWTH Aachen University, Schinkelstr. 2, 52062 Aachen, Germany, [email protected].

Introduction: In our Solar System we are facing a Test case A – Self-sinking RTGs Our close-contact wealth of different cryo-environments, some of which model is applied to studying the self-burial potential of are prime targets for future exploration missions, e.g. RTGs and other common power sources for various the Jovian and Saturnian satellites Europa and Encela- environmental conditions. dus, where an icy surface is believed to sit on top of a Test case B – Thermal melting probes The model is subglacial ocean. Sophisticated simulation models ca- also used to study translational and rotational melting pable of quantifying cryospheric processes at high spa- modes of maneurverable melting probe technologies, tio-temporal resolution are key to both scientific pro- such as the IceMole [2], see figure 2 and also [3]. gress in understanding these cryo-environment at the planetary scale, as well as to the design and implementation of any exploration mission, and to the subsequent intrepretation of its collected data. In this contribution, we will describe our progress in developing innovative computational models for applications in cryosphere geophysics by means of two generic situations, namely x thermal contact involving phase-change, and Figure 2: Computed heat-flux driven contact melting (left) and x phase-change coupled to convection processes. phase-change evolution in the presence of convection (right) [4]. Phase-change coupled to convection is relevant when wanting to understand the spatio-temporal evolution of the phase interface, e.g. the ocean-ice interface, or the evolution of englacial water reservoirs. Physical model In the absence of convection, this situation is known as the Stefan problem. It consists of Figure 1: Schematic overview of thermal contact involving phase- two free-boundary second order partial differential change (left) and phase-change coupled to convection (right). equations. Closure is again provided by the Stefan Thermal contact involving phase-change A condition. In reality, the situation often is more com- lander mission naturally implies contact with the plex and further intrigued by natural convection which ground. If the ground consists of ice and contact tem- can be modeled by a Boussinesq approximation. peratures are high enough, we will observe so-called Computational approach We utilize a phase-field close-contact melting. While thermal melting probes method based on an enthalpy formulation that provides exploit this and intentionally melt into the ground, one a flexible approach to modeling complex solid-liquid could also think of a broken power source that uninten- phase-change problems coupled to flow. The software tionally self-sinks into the ice. used for the presented test case C is implemented into Physical model In order to model close-contact the finite volume framework OpenFOAM. melting we compute the thermo-dynamic processes in Test case C – Impact of natural convection Re- the melt film. Even under low gravity conditions such analysis study of a lab experiment (see figure 2). as on Enceladus, this film will be narrow and can be Acknowledgements: The project is supported by modelled by a Reynolds equation. It is important to the Federal Ministry of Economics and Technology, consider local energy balance at the interface, a heat Germany, on the basis of a decision by the German flux jump condition referred to as the Stefan condition. Bundestag (FKZ: 50NA 1502). Computational approach We solve the governing References: [1] Schüller K., Kowalski J. (2016), equations iteratively. Starting from an initial guess for Int J Heat Mass Trans 92, 884–892. [2] Dachwald B. et the film thickness, we at first integrate the Reynolds al. (2014) Ann. Glaciol., 55, 14–22. [3] Schüller K., equation and then use its result to solve for the temper- Kowalski J. (2016), Proc. 66th IAC. [4] Schüller K., ature field. Next, the melt film thickness is systemati- Kowalski J. (2017), poster at SIAM CSE 2017. cally adjusted until the Stefan condition is fulfilled [1]. CRYOGENIC CHARACTERIZATION OF ALGAN/GAN DEVICES FOR ICY WORLD EXPLORATION. K. M. Dowling1 and D. G. Senesky12, 1Department of Electrical Engineering, Stanford University (496 Lomita Mall Stanford CA, 94305 [email protected]), 2 Department of Aeronautics and Astronautics (496 Lomita Mall Stan- ford CA, 94305 [email protected]).

Introduction Europa is the 4th largest moon of Ju- to create Ohmic contacts. Platinum was evaporated for piter, and is a leading candidate for supporting life in Schottky contacts. Finally, the devices were passivated our solar system [1] due to evidence of geologic activi- with alumina through thermal atomic layer deposition. ty below the icy surface. However, the surface of Eu- The changes in current-voltage and capacitive-voltage ropa has only been studied through previous fly-by characteristics are measured to evaluate the stability of missions such as New Horizons in 2007. NASA is this 2DEG platform from room temperature down to - planning a lander mission, with requirements for multi- 175°C using a pressure and temperature controlled directional seismometers and magnetometers for explo- environment (Figure 2). ration and analysis of the icy crust [2]. The Europa [1] Pappalardo, R. T., et al. (2013) Astrobiology, surface has an average temperature of -190°C, and sees 13.8, 740-773. [2] Europa Study Team. (2012) Nation- radiation exposure of 20 krad/day (typical silicon elec- al Aeronautics and Space Administration. [3] Spieler, tronics can tolerate 1 krad, hardened chips 1 Mrad) [3]. Helmuth. (1997) AIP Conference Proceedings. IOP Background Current strategies to enable electron- Institute of Physics Publishing Ltd. [4] Son, Kyung-ah, ics to operate in this environment include thick alumi- et al. (2010) Nanoscience and Nanotechnology Letters num shielding from radiation and active heating to 2.2, 89-95. [5] Smorchkova, I. P., et al. (1999) Journal keep silicon electronics thermally operational in cold of Applied Physics, 86.8, 4520-4526. environments. This requires more power and adds to the payload of spacecraft. New robust materials are needed for electronics to have lighter payload and lower power consumption. Gallium nitride (GaN) and silicon carbide (SiC) are wide bandgap semiconductors; they are radiation hard (>2Mrad) [4] and operate in extreme temperatures. In particular, GaN transistors (known as high electron mobility transistors) operate with a channel of a thin electron sheet known as a two dimensional electron gas (2DEG). It has been shown that this mobility is extremely high (51700 cm2/Vs) at low temperatures (13 K) [5], and can operate in regimes where silicon transistors freeze-out (< 100 K). Methods Here, we present the cryogenic operation of AlGaN/GaN devices (diodes and transistors) with Figure 1: Preliminary data of Ig-Vg curves of commer- the 2DEG structure. These devices are starting plat- cial HEMT device at cooler temperatures forms that can be adapted to create photodetectors, strain sensors, chemical sensors, and magnetic field sensors, as well as electronic circuits. Preliminary data on a commercially available transistor is shown in Fig- ure 1. A current-voltage measurement was taken between a GaN HEMT gate and source terminal to measure the Schottky contact characteristics from room temperature down to -175 ºC. The forward bias curves show an increased Schottky barrier at lower temperatures and thus a higher threshold voltage. This procedure can be applied to devices developed in house such as diodes and HEMTS. Devices were manufactured with the following mi- Figure 2: Test set up (a) Linkam Chamber and connec- crofabrication process. AlGaN/GaN was grown by tions (b) Set up with liquid nitrogen (LN2) cooling metal organic chemical vapor deposition (MOCVD) onto a (111) silicon wafer. A stack of titanium, aluminum, platinum, and gold was evaporated and annealed PLANETANS: OCEANIC PLANETS L.V. Ksanfomality Space Research Institute, Moscow 84/32 Profsoyuznaya str., Moscow 117997, Russia. e-mail: [email protected]

Abstract: The analysis of experimental data obtained bounded by its adiabatic gradient can be predicted us- in studies of extrasolar low-mass planets indicates that ing the mass and thermodynamic characteristics of the there is one more class of celestial bodies— planet; a model of outer spheres of an oceanic planet planetans—oceanic planets with global water oceans can be constructed based on the comparison between that have high, but subcritical, temperatures. A con- experimental data and theoretical studies. Major char- venient method of their analysis is using of entropy- acteristics of planetans should be as follows. Their entalphy diagram. The atmospheres of planetans masses should be lower than 6 - 9 ME (Earth masses). should be composed mainly of water vapor under high Massive hydrogen - helium atmospheres cannot be pressure. The number of detected planetans will grow held by these masses, especially if a planet is located at as new exoplanets with masses of 1–5 Earth masses are an orbital distance within the habitable zone, in which discovered. The properties of some low-mass objects they differ dramatically from neptunes. Their composi- that were determined using different methods, includ- tion may be formed beyond the ice/snow line and mi- ing KEPLER low-mass planets, GJ 1214b, and Gl grate later on. The planetans should have comparative- 581g, differ appreciably. Properties of a planetan may ly massive silicate cores, extended ice mantles, and deep oceans. The cores of planetans are differentiated have the exoplanet GL 581g, if it spherical albedo and they can include metallic fractions. The endoge- reaches a value of 0.86 (like of some of Jupiter and nous heat flux that is generated due to the decay of Saturn satellites). radioactive elements in the core of a planetan.A possible habitability of planetans is considered.

References

Ksanfomality L.V. (2014). Solar System Research, 48, 81-91

The radiation of the star Gl 581 itself is mainly con- centrated in the IR range, making the photolysis of water vapor in the upper atmospheric layers of Gl 581g inefficient. For this reason, the exoplanet Gl 581g does not loss appreciable water on a cosmogonic timescale. A model of the structure of GJ 1214b is proposed. The middle layer includes a mixture of volatile substances, mostly water, with traces of methane and ammonia. Its dense atmosphere corresponds to the observed diameter of the exoplanet, extending to 7500 km. In terms of their masses, planetans (oceanic planets) should be between Earth-like rock planets and gas-and- liquid ice giants, including Neptune-like planets and “hot neptunes”. Physical properties of oceanic planets can be predicted theoretically under definite specified conditions. For example, the depth of the ocean PRELIMINARY TRAJECTORY DESIGN OF A MISSION TO ENCELADUS. D. F. Palma1, T. Hormigo2, J.F.L. Seabra3 and P. J. S. Gil4 1Instituto Superior Técnico, Av. Rovisco Pais, 1049-001 Lisboa, Portugal, email: [email protected] 2Spin.Works S.A., Av. Da Igreja 42-6, 1700-239 Lisboa, Portugal, email: [email protected] 3Spin.Works S.A., Av. Da Igreja 42-6, 1700-239 Lisboa, Portugal, email: [email protected] 4ACMAA, IDMEC, Instituto Superior Técnico, Universidade de Lisboa, Av. Rovisco Pais, 1049-001 Lisboa, Portu- gal, email: [email protected]

Abstract: The will to explore the ocean worlds [1] Strange N.J. and Longuski J.M (2002), “Graph- such as Enceladus and Europa, that exist within the ical Method for Gravity-Assist Trajectory Design”, dynamic environment dominated by giant planets such Journal of Spacecraft and Rockets, 39, 9-16. [2] Izzo as Jupiter and Saturn, motivated the study of a new D. (2012), “PyGMO and PyKEP: open source tools for class of trajectories – complex moon tours with the massively parallel optimization in astrodynamics”, 5th objective of drastically cutting down on the cost of ICATT. [3] Strange N.J., Campagnola S. and Russell orbital capture at these moons. R.P. (2010), “Leveraging flybys of low mass moons to This work tackles the specific challenge of design- enable an Enceladus Orbiter”, Advances in the Astro- ing a preliminary spacecraft trajectory capable of de- nautical Sciences, 135,2207-2225. [4] Russell R.P. and livering as much mass as possible to Enceladus' orbit. Lara M. (2009), “On the Design of an Enceladus Sci- This is achieved by integrating different trajectory de- ence Orbit”, Acta Astronautica, 65, 27-39. sign methodologies. A framework for the mission is established, with launch windows starting as early as 2022. The mission to Enceladus is divided in three phases, each with a specific problem to be solved. The first phase handles the interplanetary transfer, with the objective of obtaining the lowest possible ΔV, while minimising the cost of the Saturn Orbit Insertion (SOI) manoeuvre. The trajectory design starts with the optimisation of several interplanetary flight sequences, obtained through a sequencing process based on the Tisserand plots [1]. To achieve that, a global optimisation procedure with a Figure 1 – Complete moon tour solution. cooperative stochastic topology is outlined and implemented using ESA’s PyGMO [2] toolbox for Python. The second phase, Saturn Orbit Insertion, employs another optimisation methodology with the purpose of determining and optimising the spacecraft's arrival geometry, including the Periapsis Raising Manoeuvre (PRM) which brings the spacecraft to the last phase of the mission – the so-called moon tour. To obtain candidate solutions for the moon tour, a modern V -Leveraging technique [3], based on a branch and bound algorithm, is used. Finally, the effect of the addition of inter-moon transfer orbits to the moon tour's solution is discussed and found to be non- negligible, within the assumptions of this class of solu- Figure 2 – Leveraging solution of the Enceladus phase tions. The probe is inserted into a pre-determined sci- of the moon tour, overlaid on a Ra vs Rp Tisserand ence orbit [4]. Plot. The outcome of this work is an example of an integrated preliminary trajectory that enables an Encelaus orbiter spacecraft of up to nearly two tonnes, within a mission time of 13.9 years. References: Trajectory Tradespace Design for Robotic Entry at Titan. E. Roelke1 and R. D. Braun2 1 Georgia Institute of Technology, 270 Ferst Drive, Atlanta, GA 30313. [email protected] 2University of Colorado Boulder, 2598 Colorado Ave, Boulder, CO, 80302. [email protected]

Introduction: In recent years, scientific focus has emphasized other ocean worlds such as Europa, Encel- adus, and Titan, due to their potential for harboring life [1]. The only spacecraft ever to land on these moons was the Huygens Probe in 2005; however, this probe’s main purpose was to study the atmosphere and surface of Titan, with no real landing target [2]. Future missions to other ocean worlds would likely require a science target and thus add several constraints to the mission such as arrival time, entry state, and aeroshell geometry, among others. Of the three ocean worlds previously mentioned, Titan is an optimal target for initial mission concepts for several reasons. The atmospheric composition, winds, and surface features are well studied by Cassini and the Huygens Probe. Additionally, of the aforementioned moons, Titan does not have a thick ice sheet to penetrate in order to sample the surface and/or liquid seas, enabling such mission to double as a stepping stone for missions to other ocean worlds. Finally, Titan exhibits a myriad of interesting planetary features that, if studied, could further the understanding of both Titan’s and the solar system’s geologic history [3]. In this paper we analyze the trade-spaces of various important parameters involved in Entry, Descent, and Landing (EDL) as it pertains to robotic missions for Titan. This will provide a guideline for optimizing a mission’s system parameters while minimizing both system complexity and the landing footprint.

References: [1] Fortes A. D. (2000) Icarus, 146, 444-452. [2] Kazeminijad B. (2007) Journal of Space- craft and Rockets, 55, 1845-1876. [3] Jaumann R. et al. (2010) Titan from Casini-Huygens, 75-140. STUDIES OF ORGANICS WITH THE COSMORBITRAP, A HRMS ANALYZER FOR FUTURE MISSIONS TO OCEAN WORLDS. L. Selliez1,2, C. Briois1, N. Carrasco2, J-P. Lebreton1, L. Thirkell1 and the Cosmorbitrap team1,2,3,4,5,6 1 LPC2E, Orléans, France, 2 LATMOS, Guyancourt, France, 3 LISA, Créteil, France, 4 IPAG, Grenoble, France, 5 CSNSM, Orsay, France, 6 ThermoFisher Scientific, Bremen, Germany; Corresponding author: laura.selliez@cnrs- orleans.fr

The in situ exploration of organic-rich environments of Solar System bodies with High Resolution Mass Spectrometry (HRMS) requires the development of a new generation of space instruments. For instance, the Cassini-Huygens mission has highlighted, among many other discoveries, that the complex organic chemistry occurring in Titan’s upper atmosphere produces both positive and negative ions of very high masses [1]. This unexpected detection, although it has allowed to make significant advances in the understanding of Titan’s environment, has shown the limitation of the Cassini-Huygens mass spectrometers for the full identification of the detected species. For this reason, future missions to Titan would take a great analytical benefit to onboard more performant mass spectrometers than those used in the Cassini-Huygens probes.

Our team, the Cosmorbitrap Consortium, developes for several years a new type of HRMS analyzer based on the use of the OrbitrapTM technology [2]. In the laboratory and for commercial instruments, this analyzer provides mass resolutions higher than 105, and up to 106 at m/z=200 [3], [4]. The Cosmorbitrap development is described in a companion presentation [5]. In this presentation, we will present our work on the study of organics with the Cosmorbitrap coupled to a laser ablation/ionization front-end. The analytical performances obtained with various organics will be presented and discussed. We will give a progress report on our capability to identify unknown large heavy organic molecules.

References:

[1] Waite et al, (2007) Science, 316, 870–875. [2] Briois et al. (2016) Planetary and Space Science, 131, 33-45. [3] Makarov (2000) Analytical chemistry, 72, 1156-1162. [4] Denisov et al. (2012) International journal of Mass Spectrometry, 325-327, 80-85. [5] Briois et al, Cosmorbitrap: R&T development of a new HRMS analyzer for future space missions, Submitted to IPPW14. INVESTIGATING TITAN’S GEOLOGY AS A LABORATORY FOR FUTURE OUTER SOLAR ICY MOONS EXPLORATIONS. A. Solomonidou1,2, A. Coustenis2, R.M.C. Lopes1, M.J. Malaska1, P. Drossart2, B. Schmitt3, S. Philippe3, C. Matsoukas4. 1Jet Propulsion Laboratory, California Institute of Technology, California, USA, 2LESIA - Observatoire de Paris, PSL, CNRS, UPMC Univ. Paris 06, Univ. Paris-Diderot, Meudon, France, 3Institut de Planétologie et d’Astrophysique de Grenoble, France, 4Department of Physics, University of Athens, Greece.

Introduction: Many different geologic units have variability. Additional units that require extensive in- been identified on Titan’s surface and are currently vestigation are the dunes, the labyrinths, the alluvial being mapped and characterized using Cassini fans, and the hummocky terrains [7,8,9]. The study of RADAR data. A key factor that allows fully character- their organic nature and their link to aeolian/fluvial and ization of Titan’s geologic evolution is to constrain the or internal processes will significantly contribute to the composition of these different geomorphologic units. identification of the nature of the surface of Titan. Data from the Visible and Infrared Mapping Spec- Characterization of terrain types: The geological trometer (VIMS) on Cassini can provide such infor- objectives request landing site measurements of the in mation, as well as insights on the atmospheric opacity situ geological context, chemical composition by sev- which is mainly due to methane [e.g. 1]. Our study eral types of spectroscopy, mineralogy provided by focuses on understanding Titan’s geologic evolution infrared data and petrological properties such as poros- and placing constraints on the interactions between ity, grain size, permeability and more. The nature interior, surface, and atmosphere by first determining (composition, texture, homogeneity) of the landing site the atmospheric contribution via radiative transfer cal- is of crucial importance especially for the safety of the culations [4;5] and then by extracting the surface albe- landing and the sampling of surface material. do which is then compared to various surface candi- Engineering considerations: Among the most im- date components. Based on these findings, combined portant engineering constrains are the precision of with other Cassini investigations, our knowledge is landing, the landing ellipse/warning track, the surface now enriched and more focused concerning future Ti- hardness, the uniformity, the inclination and the at- tan and other moons investigations. New landing mis- mospheric conditions. In situ compositional analysis sions with in situ capabilities are required for atmos- will require the ability to sample a wide range of pos- pheric and surface sampling and high-resolution imag- sible materials, from loose organic sand grains to fro- ing. One crucial parameter for a future lander mission zen aqueous cryolavas. on Titan will be the selection of an appropriate landing Conclusion: Titan, Enceladus, Europa and Gany- mede share many common characteristics with the site. In selecting future landing sites one should take main one being the possible presence of an under- into account the science objectives of the mission, the ground ocean, for which measurements of the surface physical properties, the nature of the surface, and the composition can deliver key information. The ‘lessons engineering constraints. The goal of such studies, other learned’ using Cassini/Huygens instruments and the than the deeper understanding of Titan, is to use that methods we develop here, as well as the laboratory knowledge to interpret the geology of Sarturn’s largest component databases constituted serve as a laboratory satellite and by extension that of Europa, Callisto, and for the preparation of the upcoming exploration of Ju- Ganymede, which are the next targets of ESA’s future piter’s system and for future kronian system investiga- outer solar system mission JUICE. tions. Future exploration of the icy moons could possi- Potential sites of interest: Titan’s mid-latitude bly involve measurements taken from close range, at zone consists of morphologically and compositionally these scientifically interesting areas, with landers or complex features as shown by a number of studies [e.g. with aerial elements that could explore the surfaces in 2-7]. Sotra Patera has been proposed as a potential close-up ranges. source of methane supplied into the atmosphere and/or References: [1]Tomasko et al. 2005, Nature, 438, 765-778. as areas where photolysis products such as tholin mate- [2]Lopes et al. 2010, JGR 118, 416-435. [3]Lopes et al. 2013, JGR 118, 416-435. [4]Solomonidou et al. 2014, JGR 119, 1729-1747. rial accumulate, showing a profound connection be- [5]Solomonidou et al. 2016, Icarus 270, 85-99. [6]Lopes et al. 2016, tween the surface and the atmosphere [3;4;5]. In [5] we Icarus 270, 162-182. [7]Malaska et al. 2016, Icarus 270, 130-161. reported Sotra Patera and Tui Regio as ‘changing with [8]Radebaugh et al. 2008, Icarus 194, 690-703. [9]Radebaugh et al. time’ in terms of surface albedo areas from 2005 to 2016, Geological Society, London, Special Publications 440. 2006 and 2005 to 2009 respectively. Both Sotra and Tui are of high interest for future exploration. Other interesting geological units are the geological types called ‘plains’ [6,7]. [6] recently studied several types of these ‘plains’ and identified their morphological TRAVERSING TITAN’S SEAS: SIMULATING AN AUTONOMOUS SUBMARINE OPERATING IN EXTRATERRESTRIAL SEAS. S.R. Carberry Mogan1§, I. Sahin2§, J.W. Hartwig3¥, S.R. Oleson4¥, A. Tafuni2§‡. 1Graduate Student, 2Professor 3Senior Research Aerospace Engineer, 4COMPASS Lead, §New York University Tandon School of Engineering, 6 MetroTech Center, Brooklyn, New York 11201, ¥NASA Glenn Research Center, 21000 Brookpark Road, Cleveland, Ohio 44135, ‡Corresponding author: [email protected].

Introduction: The Cassini-Huygens has been a Phase I in 2014[4]. Concepts of vehicles capable of successful mission for several reasons, including the landing in the cryogenic, hydrocarbon seas of Titan international collaboration between NASA and ESA have also been explored in the past as part of a on such a large-scale, revolutionary project. The abil- NASA–ESA Flagship mission, Titan Saturn System ity of constructing elaborate orbital trajectories has Mission (TSSM), and a NASA Discovery solicitation allowed getting close-ups on outer solar system phe- study, Titan Mare Explorer (TiME)[5]. nomena and eventually the landing of a probe on the The NIAC Phase 1 design of an autonomous sub- farthest planetary object from Earth to date, Titan. An- marine operating on Titan would have direct to Earth other important and groundbreaking discovery result- (DTE) communication for up to 16 hours per day and ing from this mission has been the discovery of stable, spend the remainder of its time conducting submerged accessible bodies of surface liquid across Titan’s polar science. While communicating with Earth on the sur- regions (Fig. 1). With these recent findings, Titan is face, the submarine would be capable of conducting now the only known planetary object in the solar sys- various experiments, such as measuring sea-surface tem other than Earth to harbor such features. Data meteorology, observing larger-scale weather activity from Cassini has indicated that these bodies of liquid and investigating tidal waves, as portrayed in Fig. 2. maintain a surface temperature of around 93 K and Conversely, while submerged, the broad range of po- contain ethane, methane, and nitrogen at a concentra- tential science would include mapping using side- tion that vary with depth. The size of such hydrocar- looking sonar, imaging and spectroscopy of the sea at bon seas ranges from hundreds of meters to 1,000 kil- all depths and sampling of the sea bottom and shallow ometers across, such as the largest Kraken Mare[3]. shoreline[4], as illustrated in Fig. 3. Indeed, the Huygens probe was also designed with the ability to float in the event that the landing was to occur in a liquid rather than on land, a possibility still unknown during the design of the probe.

Figure 2. Artist’s impression of the Titan Subma- rine conducting science on the free-surface[4].

Figure 1. Cassini’s radar mapping of Titan’s polar regions[4].

Since the finding of large bodies of liquid on Ti- tan’s surface, NASA has supported research towards developing the capabilities of conducting in-situ science. A conceptual design of an autonomous submarine to navigate Titan’s cryogenic seas has been fund- Figure 3. Artist’s impression of the Titan Sub- ed by NASA’s Innovative Advanced Concepts (NIAC) marine conducting submerged science[4]. Methodology: The design of a submarine that op- ing which of the two surfaced operating conditions erates in extreme extraterrestrial conditions requires an shows the ideal free-surface navigation technique. accurate prediction of the effects that the fluid environment will have on the submarine throughout the duration of the mission. In light of this, a collaboration involving New York University Tandon School of Engineering and NASA Glenn Research Center has been initiated to optimize the conceptual design of a submarine that will navigate the cryogenic, hydrocarbon seas of Titan using numerical techniques. Particu- larly, Computational Fluid Dynamics (CFD) is used to analyze both modes of operation for the submarine: deeply submerged and near the free-surface. Through CFD it is also able to alter the liquid properties within the simulations of the various seas that change with depth and location, such as the property values from [4]. Future investigations may involve new data re- Figure 4. A rendering of an SPH simulation of the ceived from Cassini as it continues to orbit and analyze submarine maneuvering while surfaced. Titan. The CFD studies implement a hull geometry of the Future investigations will include increasing velocities submarine maneuvering in a liquid at rest. The hull to determine the “speed-limit” of the submarine while geometry of the submarine design is imported into the operating on the surface, further comparisons between computational domain and encompassed by a liquid surfaced and slightly submerged operating conditions domain with or without a free surface. To replicate the as well as varying surface properties based on potential interactions between the liquid and the submarine, locations of the mission. flow is then simulated around the hull. While the submarine traverses through the liquid, drag and buoyancy References: forces are quantified. Results offer a prediction of the [1] Lorenz, R.D., Mitton, J., 2002, “Lifting Titan’s power requirements. They also lead to alterations of Veil: Exploring the Giant Moon of Saturn,” Cam- the design in order to increase efficiency as well im- bridge UP, Cambridge. proving navigation techniques to optimize the overall mission. [2] Lorenz, R.D., Mitton, J., 2008, “Titan Unveiled: Deeply Submerged. When analyzing the deeply Saturn’s Mysterious Moon Explored,” Princeton UP, submerged condition, Navier–Stokes based code is Princeton. utilized to represent velocity and pressure contours as well as to calculate total drag force on the submarine. [3] Hartwig, J.W., Colozza, A., Lorenz, R.D., Oleson, The absence of a free-surface allows the simple use of S., Landis, G. Schmitz, P., Paul, M., Walsh, J., 2016. a grid-based commercial software such as ANSYS© “Exploring the depths of Kraken Mare – Power, ther- FLUENT©[5]. Other investigations include varying the mal analysis, and ballast control for the Saturn Titan velocity of the submarine as well as the properties of submarine,” Cryogenics, 74:31–46. the liquid the submarine is operating in based on depth and location. [4] Oleson, S.R., Lorenz. R.D., Paul, M.V., 2015, Surfaced Condition. When examining the subma- “Phase 1 Final Report: Titan Submarine,” NASA rine navigating at the free-surface, the particle method Tech. Rep. Server, 20150014581. Smoothed Particle Hydrodynamics[6] is used to yield information about wave generation and fluid forces. [5] ANSYS© Academic Research, R. 17.1. Within the free-surface operating conditions, there is an analysis made by comparing the submarine operat- [6] Crespo A.J.C., Domínguez J.M., Rogers B.D., ing while it is “surfaced,” when the ballast tanks are Gómez Gesteira M., Longshaw S., Canelas R., Vacon- aligned with the free-surfaced, as illustrated in Fig. 4, dio R., Barreiro A., García-Feal O., 2015. “Du- and “slightly submerged,” when the submarine is sub- alSPHysics: Open-Source Parallel CFD Solver on merged to the point that only the mast and communica- Smoothed Particle Hydrodynamics (SPH),” Comp. tions array are exposed to the atmosphere. Determin- Phys. Commun., 187:204–216.

Surface Phase of the Europa Lander Mission Concept Tejas Kulkarni1, Miles Smith1, Steve Sell1, Dara Sabahi1, Jennifer Dooley, Sam Thurman1, 1Jet Propulsion Laboratory, California Institute of Technology.

Introduction: The Surface Phase represents the primary segment of a Europa Lander mission concept’s science activities. The surface phase is characterized by constrained energy, thermal and data resources. The Lander would be expected to complete between three and five sample acquisitions and perform in-situ scientific analysis of those samples, as well as additional environmental surface monitoring. The data colletected would be relayed back to Earth via the Carrier Relay Orbiter (CRO). The surface mission concept is broken into “Tals” which represent a period of performance on Europa defined by the Orbital period of the CRO, nominally 24 Earth hours. A tal is generally composed of Lander activities followed by a ground cycle. The ground cycle gives the science and engineering teams time to plan and generate commands, including making the decision on the Lander activities for the next tal. There are two primary types of Tals: Sample Tals and Monitoring Tals. A monitoring tal represents the default operational tal of the Lander where the primary activity is to monitor Europa. A sample tal represents tals during which the primary activity is sampling. The surface phase is planned to be 20 Earth days in duration, which is constrained by the total available energy. References: [1] Hand,K.P., Murray, A.E., Garvin, J.B., Brink- erhoff, W.B., Christner, B.C, Edgett, K.S, Ehlmann, B.L., German, C.R., Hayes, A.G., Hoehler, T.M., Horst, S.M., Lunine, J.I., Nealson, K.H., Paranicas, C., Schmidt, B.E., Smith, D.E., Rhoden, A.R., Russell, M.J., Templeton, A.S., Willis, P.A., Yingst, R.A., Phil- lips, C.B., Cable, M.L., Craft., K.L., Hofmann, A.E., Nordheim, T.A., Pappalardo, R.P., and the Project En- gineering Team (2017): Report of the Europa Lander Science Definition Team. Posted February, 2017 COMPARING THE PLUMES OF EUROPA AND ENCELADUS. L.W. Esposito,1 C.J. Hansen,2 G. Portyankina1 and D.E. Shemansky3, [email protected], LASP, University of Colorado, Boulder, 3665 Discovery Drive, Boulder, CO 80303-7820, [email protected], Planetzry Science Institute, 109 S. Puerto Dr., Ivins, Utah 84738, 3 [email protected], [email protected], Space Environment Technologies, Plane- tary and Space Science Division, 650 Alameda St., Altadena, CA 91001

Introduction. Evidence for Europa plumes comes from a variety of HST observations. Enceladus erup- tions have been investigated by multiple Cassini experiments for more than a decade. The observations indicate significant differences between the phenomena at the two moons. Total eruption mass. The amount of erupted water molecules in the plume is 10 times larger at Europa. This may explain why a similar search to detect the Enceladus plume in absorption against the disk of Sat- urn did not yield any detections. See Figure on right. Variability. Enceladus has a nearly constant eruption rate of 200 kg/sec, whereas the Europa plumes are epi- sodic with duty cycle of 30% or less. Height. The Enceladus plume has a scale height of 80 km; Europa plumes may extend to 200km, although a Figure: UVIS observation of Enceladus backlit by Gaussian model with z=90km can fit Sharp’s HST Saturn on Cassini orbit 120. observation. These similar heights imply a much larger ejection velocity, because of Europa’s higher surface gravity. Neutral cloud around the planet. Enceladus erup- tions create a cloud of neutral oxygen atoms around Saturn that was discovered on Cassini’s approach. No such cloud was evident at Jupiter when Cassini flew by in 2000. Ice grain fraction. The mass loading factor for icy grains at Enceladus has been estimated at 0.1-0.7 by various methods, perhaps varying with orbital longitude, time and individual eruption site. Cassini ISS, VIMS and UVIS see sunlight reflected by these grains, which is not yet observed at Europa.

CONVECTIVE HEAT FLUX AND SPECIES COMPOSITION IN THE WAKE OF A GENERIC CAPSULE DURING AN AEROBRAKING MANEUVER AT TITAN. P. Nizenkov1, A. K. Zimmer2, J. O. Elliott2, D. T. Lyons2, and S. Fasoulas1, 1Institute of Space Systems, University of Stuttgart (Pfaffenwaldring 29, 70569 Stuttgart, Germany, [email protected]), 2NASA Jet Propulsion Laboratory (4800 Oak Grove Dr, Pasadena, CA 91109, USA, [email protected]).

Abstract: Future missions to Titan are of great interest to the scientific community due to the diversity of organic chemistry that occurs in its atmosphere as well as the possibility of liquid water beneath the surface. The thick atmosphere mostly comprised of nitrogen presents a challenging aerothermodynamic environment for atmospheric entry maneuvers such as aerocapture, aerobraking or entry, descent and landing. The thermal protection system, which can be divided in the main part and the backshell, protects the payload from the incident heat. The sizing of the main heat shield can be accomplished by computational fluid dynamics (CFD) such as Navier-Stokes solvers as the flow is well within the continuum regime. However, in the wake of the capsule the continuum assumption can break down, making conventional CFD not applicable anymore. A different approach has therefore to be applied for the design of the backshell that can account from 5% up to 30% of the total entry mass, where savings directly translate in an increase in payload mass. The Direct Simulation Monte Carlo (DSMC) method [1] is utilized to simulate the rarefied gas flow in the wake of the main heatshield. For this purpose, the in- house DSMC code PICLas was verified and validated for the simulation of atmospheric entry maneuvers [2] and further developed to include polyatomic species such as methane [3]. In a first step, the convective heat flux on the backshell and species composition in the wake of the vehicle shall be investigated at the peak heating point of an exemplary aerobraking trajectory. The geometry of the generic capsule is based on the Huygens probe [4]. The presented simulation results will give insight into the rarefied gas flow behind the capsule and identify the presence of strong radiators such as CN.

References: [1] G. A. Bird, Molecular Gas Dynamics and the Direct Simulation of Gas Flows, Oxford University Press, 1994. [2] P. Nizenkov et al. CEAS Space J 9(1):127. [3] M. Pfeiffer et al. Phys. Fluids 28(2), 027103. [4] M. J. Wright et al., “Post-flight Aerother- mal Analysis of Huygens Probe”, 3rd International Planetary Probe Workshop, 2006, ESA SP-607. Dragonfly: A Rotorcraft Lander to Enable Titan Exploration. D. S. Adams,1 T. G. McGee,1 K. E. Hibbard,1 R. D. Lorenz,1 and E. P. Turtle,1 1Johns Hopkins University Applied Physics Laboratory (11100 Johns Hopkins Road, Laurel, MD 20723; Doug- [email protected]).

Introduction: Titan offers abundant, complex, di- with the abundant photochemical products that litter verse carbon-rich chemistry on an ocean world [1,2], the surface [2]. making it an ideal destination to study prebiotic chem- The basic elements and features that support the istry [3] and document habitability of an extraterrestri- lander flight, navigation, and landing capabilities are al environment. Moreover, Titan's thick atmosphere discussed. An overview of the flight performance and low gravity provides the means to access different characteristics including range, speed, power, and agil- geologic settings 10s - 100s of kilometers apart via ity, is provided, and mapped to the concept of opera- exploration by an aerial vehicle. The Dragonfly relo- tions for traverse and exploration. The entry, descent, catable lander (Fig. 1) is being proposed by APL to the and landing, sequence is discussed, including the safe- NASA New Frontiers program to explore Titan using a landing-site determination for the first landing. Finally, multi-rotor mobility system. mission specific highlights of the MMRTG powered Scientific Relevance: It has long been recognized lander are provided including: thermal, power, tele- that Titan's rich organic environment provides a unique com, & GNC. opportunity to explore prebiotic chemistry (e.g., Summary: Dragonfly is a revolutionary concept CSWG on Prebiotic Chemistry in the Outer Solar Sys- providing the capability to explore diverse locations to tem [4,5]), and development of mobile aerial explora- characterize the habitability of Titan's environment, tion was considered a next step after Cassini-Huygens. investigate the progression of prebiotic chemistry, and Studies include airships, balloons, and fixed-wing ve- search for chemical signatures indicative of water- hicles [6,7], but access to surface materials for analysis based and/or hydrocarbon-based life. presents a challenge. While multiple in situ landers could address Titan's surface chemical diversity, multiple copies of instrumentation and sample acquisition equipment would be necessary. Pinpoint targeting of a lander is infeasible, so there is no means available to ensure a direct landing at a site of scientific interest. Rotorcraft Lander: A more efficient approach is to convey a single instrument suite to multiple locations using a lander with aerial mobility. Given Titan's thick atmosphere (density 4.4× Earth's) and low gravity (1.35 m/s2), heavier-than-air mobility at Titan is highly efficient [8,9], and improvements in autonomous aircraft make such exploration a realizable prospect. A multi-rotor vehicle [10] is mechanically Figure 1. Artist's impression of the Dragonfly lander. straightforward and has simple control laws, as the proliferation of terrestrial quadcopter drones attests. Aerial mobility also allows the lander to easily overfly impassable terrain and to cover distances that would be prohibitive to a wheeled platform. There are no consumables so the lander can be relocated many times throughout its surface mission. Thus, the Dragonfly mission concept is a lander designed to take advantage of Titan's environment in order to sample materials in different geologic settings through the use of a multi- rotor mobility system. Areas of particular interest are impact-melt sheets [11] and potential cryovolcanic flows where transient liquid water may have interacted References: [1] Raulin F. et al. (2010) Titan's As- trobiology, in Titan from Cassini-Huygens Brown et al. Eds. [2] Thompson W. R. and Sagan C. (1992), C. Organic chemistry on Titan: Surface interactions, Symposium on Titan, ESA SP-338, 167-176. [3] Neish C. D. et al. (2010) Astrobiology 10, 337-347. [4] Chy- ba, C. et al. (1999) LPSC 30, Abstract #1537. [5] Lo- renz, R. D. (2000) Journal of the British Interplanetary Society 53, 218-234. [6] Leary J. C. et al. (2008) Titan Flagship study https://solarsystem.nasa.gov/ multime- dia/downloads/Titan_Explorer_Public_Report_FC_opt .pdf. [7] Barnes J. W. et al. (2012) Experimental As- tronomy 33, 55-127. [8] Lorenz, R. D. (2000) Journal of the British Interplanetary Society 53, 218-234. [9] Lorenz R. D. (2001) Journal of Aircraft 38, 208-214. [10] Langelaan J. W. et al. (2017) Proc. Aerospace Conf. IEEE. [11] Neish C. D. et al. (2017) LPSC 48. JOINT EUROPA MISSION (JEM). A MULTISCALE STUDY OF EUROPA TO CHARACTERIZE ITS HABITABILITY AND SEARCH FOR EXTANT LIFE M. Blanc1, O. Prieto-Ballesteros2, N. André1, J. Gómez- Elvira, G. Jones,V. Sterken, D. Mimoun, A. Masters, Z. Martins, E. Bunce, W. Desprats, D. Gaudin, P. Garnier, G. Choblet, V. Lainey, F. Westall, K. Szegő, M. Volwerk, J. Cooper, B. Bills, S. Vance, R. Lorenz, K. Khurana, S. Kempf, G. Colins, E. C. Sittler, T. van Hoolst, A. Jäggi, L. Iess, A. Longobardo, F. Tosi, P. Hartogh, K. Stephan, R. Wagner, N. Krupp and the JEM team. 1IRAP, France ([email protected]).

Introduction: There is a consensus in the plane- of its potential biosphere, and search for life in its sur- tary community that Europa is the closest and probably face, sub-surface and exosphere.” the most promising target to search for extant life in our solar system. The Galileo discovery of a sub- We suggest to address these goals by a combination of surface ocean likely in direct contact with a silicate five Priority Scientific Objectives, each with focused floor that could be a source of the key chemical species measurement objectives providing detailed constraints needed for the build-up of biomolecules, the many on the science payloads and on the platforms used by indications that the ice shell is active and may be partly the mission. Our observation strategy to address them permeable to transfer materials, including elementary will combine three types of scientific measurement forms of life, and the identification of candidate ther- sequences: measurements on a high-latitude, low- mal and chemical energy sources necessary to drive a latitude Europan orbit providing a continuous and metabolic activity, have raised great hopes that Europa global mapping of planetary fields (magnetic and grav- is likely habitable, and strongly support a scientific ity) and of the neutral and charged environment during plan to go there and see if it is indeed inhabited. a period of three months; in-situ measurements to be performed at the surface, using a soft lander operating We propose that ESA works with NASA, which pres- during 35 days, focusing on the search for bio- ently leads the way towards in situ exploration of Eu- signatures at the surface and sub-surface by analytical ropa, to design and fly jointly an ambitious and excit- techniques in the solid and liquid phases, and on the ing planetary mission to reach this objective. In doing operation of a surface geophysical station whose so, we aim at characterizing biosignatures in the envi- measurements will ideally complement those of the ronment of Europa (surface, subsurface and exo- orbiter; and measurements of the chemical composi- sphere), while we also want to address a more general tion of the very low exosphere in search for biomole- question: how does life develop in a specific habitable cules originating from the surface or sub-surface, to be environment, and what are the evolutional properties performed near the end of the mission during the final of a habitable planet or satellite and of its host plane- descent phase. tary/satellite system which make the development of life possible. The implementation of these three observation sequences will rest entirely on the combination of two JEM proposal was submitted to the ESA M5 call last science platforms equipped with the most advanced October 2016. instrumentation: a soft lander to perform all scientific measurements at the surface and sub-surface at a se- Scientific goals of JEM: Our search for life there lected landing site, and an orbiter to perform the orbital will build on the advanced understanding of this sys- survey and descent sequences. In this concept, the or- tem which the missions preceding JEM in its explora- biter will also provide for the lander the vital functions tion will provide: improved understanding of its origin of carrier, with the objective of carrying the lander and formation (JUNO), of its evolutionary mechanisms stack from the Earth to a Europan orbit on which it will (JUICE) and even a preliminary comparative under- release it before its descent, and of data relay during standing of its habitability: while JUICE will charac- the 35 days of lander operations. Using its own instru- terize a “type IV habitat” at Ganymede, NASA’s ment platform, it will in perform science operations EMFM mission will provide a first characterization of during the relay phase on a carefully optimized halo a “type III habitat” at Europa, using a multiple fly-by orbit of the Europa-Jupiter system, before moving to strategy. Building on these invaluable assets, the over- its final Europan science orbit for three months. arching goals for JEM is: Payload proposed for JEM: We derived from our “Understand Europa as a complex system responding science objectives a carefully selected science payload to Jupiter system forcing, characterize the habitability for the lander and for the orbiter.

Our proposed orbiter payload suite includes six well- search for extant life outside our own planet becomes proven instruments provided by European institutes in both fully credible and extremely appealing. an international collaboration framework to characterize the planetary fields and the plasma, neutrals and This way, JEM can be the next major exciting joint dust environment, fitting within the allocated mass, venture of NASA and ESA to the outer solar system, and one additional instrument that will be considered inspired by and following the unique success of Cassi- depending on the mass margin to be identified after the ni-Huygens. It will provide an outstanding opportunity assessment study. To efficiently address the radiation to preserve and develop the unique spirit of collabora- issue, we propose to decouple the sensor heads from tion and friendship which links the European and the other parts of the electronics, and to group these American planetary science communities, by propos- parts in a dedicated vault, or a well-shielded location ing to these two communities to work together toward within the platform, that will facilitate radiation miti- one of the most exciting scientific endeavours of the gation. Appropriate planetary protection measures cor- XXIst century: to search for life beyond our own plan- responding to at least Planetary Protection Category et. IVb will be implemented to all subsystems, including the payload and the spacecraft element.

Our lander science platform is composed of a geophysical station and of two complementary astrobiology facilities dedicated to biosignature characterization experiments operating respectively in the solid and in the liquid phases. The design and development of the liquid phase laboratory, called AWL for “Astrobiology Wet Laboratory”, will be a specific European contribution to the surface science platform. The two astrobiology facilities will be fed by a common articulating arm operating at the platform level that will collect the samples at the surface or sub-surface and will deliver them to the analytical facilities. We are proposing two alternative options for the deployment of AWL: inside the main platform, where it would benefit from all its infrastructure and services, or outside of it as an independent sub-platform, to be deployed with the help of the articulated arm. Further discussions between NASA and ESA will be needed to identify the best option.

Mission configuration: To fly the JEM mission, while making it affordable to the two Agencies and making JEM an appealing joint exploration venture for the two of them, we propose an innovative distribution of roles; while NASA will provide an SLS launcher, the lander stack and will cover most of the mission operations, ESA will design and provide the carrier- orbiter-relay platform. This delivery is technically possible using a safe technical approach, taking advantage of a double heritage of European developments for space exploration: the JUICE spacecraft for the JEM orbiter avionics, and an adaptation of the ORION ESM bus to the specific needs of JEM for its structure. This approach to the provision of the carrier makes it possible to propose a total contribution of ESA to JEM that fits well within the limits of an M-class mission, as required. Thanks to this approach, a joint venture of NASA and ESA to fly the first mission that will go and THE AKON EUROPA PENETRATOR. G. H. Jones1,2 , Z. Martins3, M. Blanc4, J. C. Bridges5, M. T. Capria6, P. Church7, J. Gomez-Elvira8, J. M. C. Holt5, C. Howe9, J.-C. Josset10, A. Longobardo6, R. D. Lorenz11, A. Masters3, K. Middleton9, E. Palomba6, M.-C. Perkinson12, J. D. Piercy5, W. T. Pike3, O. Prieto Ballasteros8, P. Reiss13, S. Sheri- dan14, C. S. Wedlund15, F. Westall16, R. Wimmer-Schweingruber17, V. Sterken18, and the Europa Initiative consortium. 1UCL Mullard Space Science Laboratory, UK ([email protected]), 2The Centre for Planetary Sciences at UCL/Birkbeck, UK 3Imperial College London, UK ([email protected]), 4IRAP, France, 5University of Leicester, UK, 6 INAF - IAPS Roma, Italy, 7QinetiQ Ltd., UK, 8Centro de Astrobiología, Spain, 9Rutherford Apple- ton Laboratory, UK, 10Space Exploration Institute, Switzerland, 11Johns Hopkins University Applied Physics Labor- atory, USA, 12Airbus Defence and Space Ltd., UK, 13Technical University Munich, Germany, 14The Open Universi- ty, UK, 15University of Oslo, Norway, 16Centre de biophysique moléculaire, France, 17University of Kiel, Germany, 18University of Bern, Switzerland.

Introduction: Jupiter’s icy moon Europa is Context: The Akon Europa Penetrator was pro- thought to have a subsurface ocean, which is a poten- posed to the ESA M5 mission proposal call, as part of tially habitable environment that may host life. The the Europa Initiative: an international effort to identify seafloor of Europa is thought to be habitable due to the and pursue the most promising collaborative interna- existence of ecosystems in analogue environments on tional missions to explore Europa. Akon would have Earth, such as in deep sea hydrothermal vents. Hydro- been transported to near its target body by the planned thermal vents could provide the necessary energy NASA Europa Lander Mission. source for organisms likely living in such an extreme habitat. Furthermore, the salty subsurface ocean may be in contact with the rocky mantle, leading to several complex chemical reactions, and generating chemical disequilibrium necessary for life. Europa is therefore the ideal location to determine whether extra-terrestrial life currently exists in the solar system beyond the Earth. Biosignatures from any potential extra-terrestrial life forms present in the subsurface ocean of Europa could be detected and analysed in several ways; they could be incorporated into the ice, and transported to the surface of this icy moon, which would then be detected. However, UV and ionizing radiation driven reactions would modify and destroy these biomarkers. It is therefore unlikely that biomarkers are present on the surface of Europa, and one therefore needs to penetrate the icy surface of Eu- ropa to sample material shielded from the harsh radiation environment. This may be achieved by the Akon Europa Penetrator Mission Mission – “Akon” (Άκων) is the ancient Greek word for a javelin, as presented to Europa by Zeus. The Akon Europa Penetrator Mission would enhance and complement the science return from the Artist’s impression of Akon descending to the surface NASA Europa Clipper and ESA JUpiter ICy moons of Europa. It would measure approximately 390mm Explorer (JUICE), providing ground-breaking data and in length, and 190mm in width. science. The instrument suite of the Akon would obtain data to address several key scientific areas; not only Unlike the NASA soft lander which it would com- the topics of habitability and biosignatures, but also plement, Akon’s kinetic penetrator design would pro- our understanding of Europa’s subsurface ocean, crus- vide direct access to Europa’s subsurface. Such in- tal dynamics, magnetospheric environment, surface strumented impactors provide an efficient way of de- morphology, and general chemical compositon. Akon livering instruments to planetary sub-surfaces. Penetra- would require the transport and delivery to the surface tors have been studied and tested extensively, and they of Europa by the NASA Europa Lander mission. are designed for impacts at 100-500 m/s, resulting in penetration into the surface of 0.5 to a few metres. ESA has successfully funded technology development • Descent Imager studies into the feasibility of delivering instrumented • Thermogravity Analyser penetrators into the subsurface of icy satellites such as • Wet Chemistry Package Europa. The consortium funded by ESA to conduct • Habitability Conditions Package these studies have been led by Airbus Defence & • Exosphere and Regolith Mass Spectrometer Space. These studies have led to a mature design, • Sample Imager which allows operation at extremely low temperatures • Magnetometer (~80K), effective insulation of the penetrator interior, • Seismometer and long-term power requirements through the applica- • Energetic Particle Instrument tion of advanced battery technology. The payload would have been provided by instrument The four key science objectives of the Akon Mis- consortia primarily funded by the ESA member states. sion are the following: 1. Search for biosignatures in near-surface material. Summary: We shall provide an overview of the Akon 2. Determination of the internal structure of Europa penetrator design, its proposed mission, and its instru- and its dynamics. ment suite. Planetary kinetic penetrators hold great 3. Determination of the existence and characteris- promise for the direct delivery of rugged instruments tics of a subsurface ocean. to planetary subsurfaces. Akon is an excellent example 4. Characterization of the physical (e.g. radiation, of the scientifically valuable potential missions that thermal, magnetic, electrical, mechanical) and chemi- could employ this technology. cal environment of the near-surface region.

Mission Overview: Although other delivery options exist, the nominal proposed scenario involved the release of Akon from the NASA Europa Lander at an altitude of ~35km. The penetrator would accelerate during freefall to strike the surface at ~300m/s, embedding itself in the moon’s near-surface layers. To achieve the four key science objectives described above, data would be gathered during three mission phases (descent, impact and sub-surface analysis). Most of the experiments would be carried by the penetrator itself, while the remaining would be included on the delivery module. The instruments are split into two groups: those mounted on the delivery module to make measurements during the descent phase, and those included in the penetrator which would operate under the subsurface of Europa. The Akon penetrator instruments will be further divided in two groups: short-term instruments (i.e. the thermally-isolated for in situ analysis of subsurface material), and the long- term instruments (i.e. the back-end mounted geophysics instruments for long-term observations and regolith imaging). Akon would be buried up to several metres in the subsurface of the moon. Samples would be gathered from the subsurface material adjacent to the penetrator, and studied onboard by a suite of instruments. Other instruments would image the surface material, while the geophysics instruments would take long-term measurements of the magnetic, gravity, and radiation conditions. Instrumentation: The proposed scientific instruments are based on flight-proven technologies, and the nominal payload would have consisted of the following: EARLY EVOLUTION OF THE MISSION ARCHITECTURE FOR NASA’S EUROPA LANDER CONCEPT. A. K. Zimmer1, S. W. Sell1, A. Frick1, T. P. Kulkarni1, M. D. Spaulding1, E. D. Skulsky1, A. M. San Martin, and D. M. Kipp1, 1Jet Propulsion Laboratory, California Institute of Technology, (4800 Oak Grove Dr, Pas- adena, CA 91109, USA, [email protected], [email protected], [email protected], [email protected], [email protected], [email protected], [email protected], [email protected])

Abstract: NASA has extended the scope of the The third architecture would provide a dedicated potential exploration of Europa beyond the planned launch for the Lander spacecraft a few years after Europa Clipper mission and initiated a Pre-Phase A Clipper’s projected launch. Shortly before Europa final mission concept study of a Europa Lander. With a approach, the Lander spacecraft would separate from massive liquid water ocean beneath its icy crust, Euro- its carrier, after which the carrier would function as a pa is one of the Solar System’s prime candidates for telecom asset to relay the lander signal to Earth. While hosting life. the carrier transfers to its relay orbit, the lander would The objective of this paper is to provide an over- proceed to its target. This architecture concept would view of several different Europa Lander mission archi- be characterized by full independence between the tectures and mission design concepts that have been Clipper and Lander missions with exceptions for using studied to assess their relative science, engineering, Clipper data for the landing site selection and the po- and operations impacts as well their synergies with the tential Clipper use for contingency relay of Lander planned Europa Clipper mission. Since our current surface data to Earth. This approach would allow post- maps of Europa from the Voyager and Galileo space- landing telecom relay geometry optimization between craft do not provide sufficient resolution for determin- the Lander on the surface and the carrier in orbit, opening the most scientifically interesting site that is also ing up the Jupiter-Europa low-energy regime and ena- safe for landing, the landing site selection would be bling the carrier to linger in the vicinity of Europa for based on reconnaissance data acquired by Clipper’s the full duration of the surface mission. This cannot be onboard cameras while the notional Lander is en route achieved with Clipper in the Piggyback or Free-Flyer to Europa. Consequently, it is a requirement that the architectures because the radiation dose associated trajectory as well as the flight system design be robust with such trajectories would be prohibitive for Clipper to enable landing in a wide range of locations on Euro- after its prime mission. pa and a variety of terrains. Three options were assessed that cover a wide range of potential mission architectures. In the first architecture, the Piggyback scenario, the Lander would be mounted to the Clipper spacecraft and launched on a single launch vehicle. Following its prime mission, the Clipper spacecraft would pump down its orbit to reduce its velocity with respect to Europa and deliver the Lander after which the Lander would rely on Clipper as a telecom asset to relay the Lander signal to Earth. The second architecture, the Free Flyer scenario, assumes that the Lander and Clipper spacecraft would be co-manifested for launch but they would separate after Jupiter orbit insertion. Similar to the Piggyback scenario, Clipper would relay the Lander signal to Earth after landing, which would occur after Clipper finishes its prime mission. De-coupling the trajectories of Clipper and the Lander after arrival at Jupiter would allow (1) the Lander trajectory to be optimized for minimum radiation exposure while Clipper acquires the required reconnaissance data, as well as (2) the Clipper trajectory to be optimized post-landing for maximum telecom relay visibility. Site Selection for a Europa Lander Concept: Approach on Collaborative Mission Planning and Site Certifi- cation. G.H. Tan-Wang1, M. Heverly1, M. Cable1, B. Buffington. 1Jet Propulsion Laboratory, California Institute of Technology, 4800 Oak Grove Drive, Pasadena, CA 91109

Selecting a landing site for the Europa Lander Concept is challenging in that the best information currently available is from the Galileo mission, including a single image at 6 m/pixel. There are another handful of images at <50 m/s, and rest of Europa images are at 500-2000 m/pixel resolution. Less than 10 areas have coverage in stereo. What the current concept for a safe Descent and Landing needs are more on the order of 0.5 m/pixel, stereo images for a multitude of candidate landing sites. The Europa lander mission concept would use the reconnaissance data from the planned Clipper mission which is planning for over 40 close flybys of Europa. The approach for a successful landing is to both 1) have a robust landing design tested against a range of surface terrains and 2) selecting a site that best matches to the tested surface terrains. This paper will focus on the initial concepts of the latter step on how Clipper’s data could identify sites of interest and the site certification process that would narrow down which can be landed safely. Overview of the Europa Lander Flight System Concept. A. Frick1, S. W. Sell1, J. S. Stehly1, , D. M. Kipp1, B. T. Cook1, M. E. Greco1, D. M. Kipp1, A. K. Zimmer1, T. P. Kulkarni1, E. D. Skulsky1, M. D. Spaulding1 1Jet Propulsion Laboratory, California Institute of Technology, 4800 Oak Grove Drive, Pasadena, CA 91109

NASA is actively studying a Europa Lander mission concept that would follow the planned Europa Clipper (multiple flyby) mission and conduct an in-situ search for life on the surface of Europa. The baseline flight elements of the Lander mission would include the following elements: A primary battery-powered Lander; a Descent Stage (DS) employing a “sky crane” maneuver similar to the one first used on the Mars Sci- ence Laboratory; a DeOrbit Stage (DOS) using a solid rocket motor (SRM); and a Carrier & Relay Stage (CRS) responsible for interplanetary cruise, Jovian tour, and subsequent data relay operations during the Lander’s Deorbit, Descent, & Landing (DDL) and Sur- face Mission phases. This paper primarily addresses the conceptual flight system architecture of the Lander, Descent Stage, and DeOrbit Stage. The design of these flight system elements is driven by several unique challenges of the Europan environment and mission objectives. While landing site selection would be facilitated by images provided by the Europa Clipper mission, the landing system must account for yet uncertain and possibly challenging terrain. To this end, the Descent Stage would employ Terrain-Relative Navigation (TRN) based on capabilities developed for the Mars 2020 mission, while the lander employs a touchdown system of conforming stabilizers to provide a level platform, even on challenging terrain. Furthermore, both the Descent Stage and Lander would feature aluminum “vaults” to protect avionics from Europa’s harsh radiation environment. Each vault may include a novel in-situ sterilization concept to comply with stringent planetary protection requirements that aim to protect Europa’s subsurface ocean from terrestrial microbial contamination. The Lander’s sampling system would collect samples up to a depth of at least 10 cm to gather material that should be largely unaltered by radiation processing on the surface, while accounting for a broad range of possible surface compositions and topographies. Sam- ples would be delivered to a suite of instruments and analyzed over the course of an approximately 20-day surface mission.

EUROPA LANDER: THE CRUISE AND RELAY STAGE ARCHETECTURE S.R. Vernon (Steven. R. [email protected]), C. Sheldon ([email protected]), D. Adams ([email protected]) The Johns Hopkins University, Applied Physics Laboratory, Laurel MD.

The Europa Lander Mission Concept Study is de- countered during the mission. The launch, cruise, Jupi- signed to deliver a lander to the surface of Europa and ter orbit insertion (JOI), and the pump-down to Europa conduct scientific investigations on the surface chemis- phases of the mission are also introduced. The mission try and other features of Jupiter’s moon in the search design description proceeds with the deployment of the for life. The Jet Propulsion Laboratory (JPL) and the DOV/Lander event as the CRS transitions from a de- Johns Hopkins University/Applied Physics Laboratory livery vehicle into the relay phase of the Mission. The (JHU/APL) are implementing this mission for NASA. separation dynamic events are described including the The flight system, consisiting of the Cruise and Re- jettison of propulsion tanks, deployment of the lander lay Stage (CRS) and Deorbit Vehicle (DOV) [DOV and the divert maneuvers. includes Descent Stage (DS) and Lander], is planned to launch in October of 2025 aboard a NASA Space Launch System (SLS) Block 1-B launch vehicle. The SLS Launch Configuration is depicted in Figure 1. The CRS will deploy the Lander to the Europan surface after arrival in 2032.

Fig. 3: Comparison of an Apollo Lunar Module with the CRS [1].

The CRS is in the same class as the largest spacecraft designed for an interplanetary mission [1], [2] (see figure 3). A technical description and overview of the CRS is provided. The very large and unique pro-

Fig 1: SLS Launch Configuration for the CRS pulsion system, including the jettisoned fuel and oxidizer tanks is described along with a largely passive Figure 2 depicts several of the critical deployments thermal system design. The power system is introduced required to deliver the Lander to the surface, which including the largest solar array ever flown. will be explained in detail. Figure 2 also depicts the The CRS mission imposes many challenges upon DOV bio-barrier left behind in the CRS, showing it as the spacecraft design including the radiation environ- if it was deployed. ment that must be endured. The radiation environment encountered at Europa by the CRS is severe and unique. The impacts to the CRS system design are coupled to solutions that will be presented. Several unique approaches to autonomy and fault protection, associated with deployment of the Lander and transition into the relay phase are described. Given that Europa is considered to potentially have an ecosystem, the mission must prevent forward contamination of Europa and comply with NASA’s planetary protection requirements [3]. An exhaustive micro- Fig2: DOV Deployment Sequence w/Bio-Barrier Il- bial reduction campaign will be run against all hard- lustrated ware on each flight system. The methods used for this A technical description of the CRS Spacecraft and reduction drive parts and materials selections for hard- mission design, technical challenges and telecom archi- ware components. tecture are the primary focus of this paper. Accommodations to the Lander payload and the The unique and challenging mission design is brief- lander deployment system and design are presented. ly introduced including a description of the phases en- Prominent features and several selected key and driving requirements are illustrated where appropriate and described. The Europa Clipper hardware development leveraged for the CRS is presented.

+Z

+Y

+X

Fig. 4: CRS antenna fields of view: low gain antennas (purple), fan beam antennas (green), medium gain antenna (red) and high gain antenna (blue).

The CRS telecom system (see figure 4) performs the dual functions of telemetry, tracking, and command (TT&C) with the ground and enables Lander relay communications. The Lander does not have a direct-to- Earth communications capability, and relies upon the CRS for communications with Earth. For the first time in deep space, an X-band proximity link will be used to return data from a Lander. The telecom relay phase of the mission begins after Lander separation from the CRS. The Lander will transmit telemetry to the CRS before, during and after deorbit, descent and landing (DDL). Following DDL, the Lander surface mission will last approximately 20 days. Contacts will be scheduled with the capability for either the CRS or the Lander to initiate an unscheduled contact. After completion of the Lander’s surface mission the CRS will enter the disposal phase. Options include placing the CRS in a permanent or temporary parking orbit or disposal by impact into Io, Ganymede, or Calisto. This phase of the mission poses several challenges due to the need to maintain an operational spacecraft for up to 30 days after the accumulated radiation dose from earlier mission phases

References: [1] Jimenez, S. et al (1967) “Apollo Training: Apollo Spacecraft & Systems Familiarization”. [2] Orloff, R. (2000) “Apollo by the Numbers: A Statistical Reference”. [3] Hand, K.P. et al (2016) “Europa Lander Study 2016 Report”. Touchdown System for a Europa Lander Concept. S. W. Sell1, B. T. Cook1, M. E. Greco1, 1Jet Propulsion La- boratory, California Institute of Technology, 4800 Oak Grove Drive, Pasadena, CA 91109

The task of landing on the Surface of Europa has many challenges. One of the of the largest is designing a touchdown system for this relatively unknown surface. While reconnaissance photos exist of the surface, most have resolutions on the order of 500-2000 m per pixel, a few are <50 m per pixel, and only a sungle image at 6 m per pixel. In addition, the composition of the surface is expected to vary from a soft snow-like texture to granite-hardness ice. This work presents the challenge of designing the touchdown system for a Europa Lander concept which would be launched prior to any additional reconnaissance from the currently planned NASA Europa mission. A brief history of the concepts considered is discussed along with the mechanical design of the final concept as well as its conceptual operations. The design provides robustness to up to 0.5 m of terrain relief within the lander footprint while providing a level platform for the main body of the lander. Beyond 0.5 m of relief a graceful degradation of performance is achieved with up to 1 m of terrain relief. E3P: Europe's New Vision for Space Exploration. D. Parker1, 1Director of Human Spaceflight and Robotic Ex- ploration, European Space Agency ([email protected], ESA ESTEC, Keplerlaan 1, PO Box 299 NL-2200 AG Noordwijk, The Netherlands).

Since the late nineties, ESA has worked intensively to build a sound strategic vision for Europe’s space exploration efforts. Through strong engagement with the international community, notably including the International Space Exploration Coordination Group, European thinking matured into the European Space Exploration Strategy, adopted at the ESA Council at ministerial level in Luxembourg (2014). This presents a vision where four strategic objectives are guiding Europe’s exploration efforts: science, economics, global cooperation, and inspiration. While the long- term ambition is one day to have Europeans living and working on Mars, it recognises the value of LEO and the Moon as important stepping stones supporting all four objectives, and emphasises the symbiotic role of robotic and human exploration.

In December 2016, the ESA Council at ministerial level in Luzern adopted the resolution "Towards Space 4.0 for a United Space in Europe". This resolution encourages ESA to "pursue and further strengthen Euro- pean cooperation in the space sector for the benefit of European citizens". Helping translate this vision into reality, Member States also approved the European Space Exploration Envelope Programme (E3P). E3P integrates existing robotic and human space exploration projects into a single compelling programme. In the first period to 2020, exciting missions targeting all the three exploration destinations. Thus, successive European astronauts will perform excellent science on the ISS including preparation for beyond LEO human exploration. The first two European Service Modules for NASA’s Orion vehicle will enable humans to travel to the lunar vicinity for the first time since Apollo. Meanwhile, ESA will also contribute to the Russian Luna 25, 26 and 27 robotic Moon missions alongside ExoMars 2020, which will deliver European drilling and life-search technology to Mars. Furthermore, study and technology work will prepare future ESA missions for implementation while a partnership initiative will help European industry establish commercialised exploration services.

Taken together, this unified European programme will secure more discovery, excitement and international cooperation in the years that lie ahead. The Saturn PRobe Interior and aTmosphere Explorer (SPRITE) Entry Probe Science D. Atkinson1, A. A. Simon2, D. Banfield3, S.Atreya4, J. Blacksberg1, W. Brinckerhoff2, A. Colaprete5, A. Coustenis6, L. Fletcher7, T. Guillot8, M. Hofstadter9, J. Lunine3, P.Mahaffy2, M. Marley5, O. Mousis9, T. Spilker10, M. Trainer2, C. Webster1, 1Jet Propulsion Laboratory, California Institute of Technology ([email protected]). 2NASA Goddard Space Flight Center, 3Cornell University, 4Univ. Michigan, 5NASA Ames Research Center, 6LESIA, Observ. Paris-Meudon, CNRS, Paris Univ., France, 7Univ. Leicester, 8Observatoire de la Cote d'Azur CNRS / Laboratoire Cassiopée, 9Laboratoire d'Astrophysique de Marseille, 10Independent Consultant

Introduction: The Vision and Voyages Planetary Decadal Survey [1] identified Saturn as a target of high priority for a New Frontiers probe mission. To better constrain models of Solar System formation, as well as to provide an improved context for exoplanet systems, fundamental measurements of noble gas abundances and isotope ratios of hydrogen, carbon, oxygen, and nitrogen, as well as the interior structure of Saturn are necessary. The New Frontiers SPRITE mission would achieve the science goals outlined in the Decadal Sur- vey and would provide ground truth for remote sensing to improve understanding of Saturn’s interior structure and composition, and (by proxy) those of extrasolar giant planets.

Many key questions regarding the structure and composition of Saturn’s atmosphere remain unanswered, including the abundance of noble gases, helium in particular needed to understand the formation history and evolution of Saturn, key isotopic ratios, and the abundance of water in the deep atmosphere, a key diagnos- tic of Saturn’s formation since it is thought that the heavy elements were delivered to Saturn by water- bearing planetesimals. Additionally, the structure of Saturn’s deep interior including the presence of a core and any layered structure would test instability models in the protosolar nebula.

SPRITE would make measurements that address these key questions through delivery of an atmospheric entry probe, as well as remote sensing from the carrier spacecraft. SPRITE would provide direct measurement of composition and atmospheric structure (including dynamics) along the probe descent path, providing science that is not accessible to remote sensing measurements, as well as providing ground truth for tropo- spheric measurements from carrier remote sensing. In addition to the deep atmospheric composition, SPRITE would measure Saturn’s profile of temperature and pressure, and vertical profile of zonal winds along the probe descent path.

Reference: [1] “Vision & Voyages for Planetary Science in the Decade 2013-2022,” National Academies Press, Mar. 7, 2011. Two-Staged Saturn Probe to Explore 60+ Bar Atmosphere. K. M. Sayanagi1, R. A. Dillman2, D. Goggin2, K. Vipavetz2, and K. Gough2, M. Darrach3, T. R. Spilker4, N. Tosoc2, R. Fairbairn2, A. Bowes2, J. Brady2, S. Bowen2, F. Stillwagen2, S. Horan2, W. Bruce2, C. Thames2, N. Ryan2, 1Hampton University (Atmospheric and Planetary Sci- ences Department, 23 E Tyler St. Hampton, VA, 23668, USA, [email protected]), 2NASA Langley Research Center, 3Jet Propulsion Lab, 4Planetary Flight Architect.

Introduction: We present results of an engineer- large instrumented probe opens a parachute at ~100- ing design study for a two-staged Saturn probe in mbar, jettisons the heatshield, and then releases the which we examined the feasibility of releasing a de- parachute for a rapid descent. Our study shows that, tachable rapid-descent deep-stage probe from a para- due to its aerodynamically unoptimized shape, a large chuted shallow-stage that relays data to the carrier- single-stage probe takes 104 minutes to reach 20-bar. relay spacecraft (CRSC). Our study demonstrates the Two-Stage Probe – Deep-Stage Design: The viability of configuring a two-staged probe that could Deep-Stage delivers a 10-kg mass spectrometer (JPL’s reach 63-bar in 70 minutes after atmospheric entry, and Atmospheric Composition Explorer (ACE) design is return data at a rate of at least 4.2 kbps. adopted as a notional instrument) and a 1-kg Atmos- Scientific Objectives: Our design enables in-situ pheric Structure Instrument (ASI) which measures at- measurement of deep water abundance of Saturn, mospheric temperature, pressure and probe accelera- which is crutial for understanding the formation pro- tion as functions of depth. It carries a 1-W transmitter cess of Saturn and the evolution of the outer solar sys- and an omni-directional antenna to send data to the tem. To measure the deep water abundance, the probe Shallow-Stage. In our design, the Deep-Stage mass is must reach below the water condensation level, which estimated to be 40 kg in a cylindrical package 56-cm is expected to be at 20-bar (~300km below 1-bar) [1]. long, 30-cm diameter with a nose cone. Our study Reaching below the water cloud would also enable demonstrates that the deep stage could reach 63 bar in determination of atmospheric stratification in the layers 70 minutes after atmospheric entry. affected by moist convection, revealing the role of la- Two-Stage Probe – Shallow-Stage Design: In our tent heat release in shaping the atmospheric structure. design, the Shallow-Stage contains a communication Challenges: We aim to address two challenges in relay system, Deep-Stage separation mechanism, and a performing in-situ measurements of water abundance parachute. To relay data to CRSC, it carries a 20-W below the 20-bar level. First, a probe that could reach transmitter and a 5-turn helical antenna with a 7.5db 20-bar must descend rapidly because the CRSC in a directional gain with HPBW of 69°. We found that this flyby trajectory could remain in communication with configuration allows significant margin to carry extra the descent probe for only ~70 minutes. Second, due to batteries or more instruments before all of the 216.4-kg the microwave absorption by ammonia vapor, Saturn’s mass budget for the probe is allocated. Both the Deep- atmosphere attenuates the radio signal significantly, Stage and the Shallow-Stage are powered by batteries and increases the transmission power requirements for for at least 70 minutes plus margins. any data return [2]. A two-stage probe could solve Link Analysis: Our radio opacity analysis shows these challenges, in which a fast-descending deep-stage that, for 401-MHz frequency we considered, the signal would deliver a mass spectrometer to 20-bar and be- attenuates by ~12-dB between 60-bar and 1-bar. The low, sends data to a parachuted shallow-stage, which Our link analysis shows that the Deep-Stage can trans- in turn would relay data to CRSC. mit data to the Shallow-Stage at a rate of 4.2 kbs from Design Constraints: We designed a two-stage 63-bar with 9.4 dB margin. The Shallow-Stage to probe system that could be carried on the Saturn Probe CRSC link with a 61,000-km range (farthest consid- mission designed during the 2013-2022 Planetary Sci- ered for the 70-min window) can be closed with 4.2 ence Decadal Survey, which delivered a probe with kbps and 7.1 dB margin. 216.4 kg mass to Saturn [3, 4]. To perform our link Cost Analysis: Our parametric cost analysis esti- analysis, we adopted the Saturn arrival conditions, re- mates that a single-stage reference probe should cost lay geometry (carrier-probe range and carrier off-zenith $112.7M, which is comparable to the $95.6M probe angle as seen by the probe), and CRSC trajectory de- cost estimated by the decadal mission study, in which rived by the decadal mission study. We performed the total mission cost estimate was $990M [3, 4]. In thermal analysis using standard tools. We used SEER- comparison, our cost estimate for a two-stage probe is H cost model to perform parametric const analysis. $137.3M, $17M more than our single-stage reference Single-Stage Referene Design: To evaluate the design. The $17M cost difference represents a 2.5- performance of our two-stage design, a single-stage percent increase from the total mission cost of $990M. reference design was also studied. In this scenario, a Conclusion: Our study demonstrates that a two- stage design is a viable approach to reach below the water cloud level of Saturn. Our analysis demonstrates that the Deep-Stage could reach the 63-bar in the 70- minute communication window, and return data at a rate of at least 4.2 kbps to CRSC via the Shallow- Stage. References: [1] Weidenschilling, S. J. and Lewis, J. S. (1973) Icarus. [2] Spilker T. R. (2007) IPPW-5. [3] Spilker, T. et al. (2010a). Saturn Atmospheric Entry Probe Trade Study. Mission Concept Study Final Re- port, Planetary Science Decadal Survey. [4] Spilker, T. et al. (2010b). Saturn Atmospheric Entry Probe Mis- sion Study. Mission Concept Study Final Report, Plan- etary Science Decadal Survey.

RETURN TO THE ICE GIANTS. Kim Reh1, Mark Hofstdater1, Amy Simon2, and John Elliott1 1Jet Propulsion Laboratory/Caltech, Pasadena, CA, United States, 2NASA Goddard Space Flight Center, Greenbelt, Md, United States

Introduction: Voyager 2 mission flew by Uranus in 1986 and Neptune in 1989 resulting in stunning remote observations not previously accessible from the Earth. There have been no follow-up space flight missions to examine ice giants and, as a result there are significant gaps in our understanding of planetary formation and evolution. This gap limits our understanding of the solar system and that of exoplanets; the majority of planets discovered around other stars are thought to be ice giants. Ice giants are likely to be far more abundant in our galaxy than previously thought.

The U.S. 2011 Planetary Science Decadal Survey committee recognized the importance of Uranus and Neptune, and prioritized the exploration of the ice giants. Following from this, NASA and ESA have recently completed a study of candidate missions to Ura- nus and Neptune, also known as the ice giant planets. The intent was to examine what could be accomplished within the budget realities of the predictable future. This "Pre-Decadal Study," focused on opportunities for missions launching in the 2020’s and early 2030’s.

This paper presents results from the Ice Giants study (science, architectures and technologies) and concludes that compelling and affordable missions to the Ice Giants are within our reach.

References: [1] Mark Hofstdater, Amy Simon, John Elliott, Kim Reh (2017) Ice Giants, Pre-Decadal Survey Mis- sion Study Report. THE HERA SATURN ENTRY PROBE MISSION: A PROPOSAL IN RESPONSE TO THE ESA M5 CALL. O. J. Mousis1, D. H. Atkinson2, and the Hera team,1Aix Marseille Université, CNRS, Laboratoire d’Astrophysique de Marseille, UMR 7326, 13388, Marseille, France ([email protected]), 2Jet Propulsion Laboratory, 4800 Oak Grove Drive, Pasadena, CA 91109, USA.

Introduction: The Hera Saturn entry probe mis- tists and engineers from both agencies and many affili- sion is proposed as an ESA M-class mission to be pig- ates participating in all aspects of mission development gybacked on a NASA spacecraft sent to or past the and implementation. A Saturn probe is one of the six Saturn system. Hera consists of an atmospheric probe identified desired themes by the Planetary Science built by ESA and released into the atmosphere of Sat- Decadal Survey committee on the NASA New Fron- urn by its NASA companion Saturn Carrier-Relay tier’s list, providing additional indication that a Saturn spacecraft. Hera will perform in situ measurements of probe is of extremely high interest and a very high the chemical and isotopic composition as well as the priority for the international community. structure and dynamics of Saturn’s atmosphere using a The Hera Saturn probe mission will begin its flight single probe, with the goal of improving our under- phase as an element of a NASA Saturn mission (likely standing of the origin, formation, and evolution of Sat- a NASA New Frontiers mission) launch to place both urn, the giant planets and their satellite systems, with the NASA spacecraft, which functions also as the Hera extrapolation to extrasolar planets. Hera will probe probe's CRSC, and the Hera probe on a transfer trajec- well into and possibly beneath the cloud-forming re- tory to Saturn. The NASA CRSC releases the probe on gion of the troposphere, below the region accessible to a ballistic trajectory that will carry the probe into remote sensing, to locations where certain cosmogeni- Saturn's atmosphere several weeks later. During the cally abundant species are expected to be well mixed. ~70-90 minute Hera descent, the overflying CRSC will maintain the data relay link with the descent mod- Deciphering the origin of the outer solar system ule, storing multiple copies of each channel of the and characterizing the atmospheres of giants: In probe's science data in redundant onboard storage me- situ measurements of Saturn’s well-mixed atmospheric dia for later downlink to Earth. After the data reception gases will provide a vital comparison to the Galileo window ends the CRSC will turn its high gain antenna probe measurements at Jupiter, and a crucial “ground to Earth, downlink each the entire dataset multiple truth” for the remote sensing investigations by the times, and then begin its NASA science mission. Cassini orbiter. Hera will improve our understanding of the physical processes responsible for the formation Costing: A preliminary cost study realized by Air- of giant planets (contribution of the local solar nebula, bus Defence and Space indicates that the total budget accretion of icy planetesimals, and nature and for- of the Hera probe is about 293 million euros, excluding mation temperature of the latter), and will shed light on the Science payload and the TPS supplied by NASA. the composition of giant planet precursors and the dy- The total estimated budget for a single Hera probe is namical evolution of the early solar system. Hera will significantly lower than the cost of the Huygens probe, also address the question as to why Jupiter and Saturn estimated to be ~455 M€ by ESA in 2016 economic are so different in size, density and core dimension by conditions. Airbus Defence and Space estimates that investigating different pathways to planetary for- the industrial cost corresponding to the manufacture of mation. Hera will thus provide new insights on the a second Saturn probe would be ~20% lower than a mechanisms that led to the stunning diversity of giant single one, i.e. ~148 million euros, implying that send- planets in our solar system and in exoplanetary solar ing two ESA probes in Saturn’s atmosphere would still systems. Hera will investigate Saturn’s atmospheric match the M5 envelope. A multi-probe contribution dynamics along its descent trajectory, from (1) the from ESA could be of the highest interest for NASA in vertical distribution of the pressure, temperature, the case of the selection of a Uranus mission in the clouds and wind speeds, and (2) deep wind speeds, framework of the “Roadmaps to Ocean Worlds” differential rotation and convection, by combining (ROW) program. A Uranus orbiter could use the gravi- probe, gravity and radiometric measurements. This is ty assist of Saturn to fly toward Uranus. Given the sim- the next logical step in our exploration of the gas gi- ilarities for the entry conditions in the two giants (T. ants beyond the Galileo and Cassini missions. Spilker, personal communication), two identical ESA Hera probes could be released in the atmospheres Sat- Mission design: Following the highly successful urn and Uranus. example of the Cassini-Huygens mission, Hera will carry European and American instruments, with scien- Hera website: http://hera.lam.fr Saturn PRobe Interior and aTmosphere Explorer (SPRITE): Mission Implementation Overview. M. A. Lobbia,1 R. M. Danner1, and A. Simon; 1Jet Propulsion Laboratory, California Institute of Technology, 4800 Oak Grove Dr., MS 301-490, Pasadena, CA 91109-8099; 2NASA Goddard Space Flight Center

Introduction: The most recent Planetary Science Therefore, SPRITE’s primary payload would consist of Decadal Survey [1] provided a recommended set of two spectrometers (Quadrupole Mass Spectrometer – mission concepts that NASA should consider for future QMS, and Tunable Laser Spectrometer – TLS) and an planetary investigations. One of the six mission themes atmosphere structure investigation (ASI). The spec- suggested for consideration under the NASA New trometers would target the first science goal of measur- Frontiers Announcement of Opportunity is a Saturn ing atmospheric composition including noble gas and Probe mission. This mission would deliver an entry elemental abundances, and isotopic ratios; the ASI probe into Saturn’s atmosphere (similar to the Galileo would provide characterization of the entry accelera- Probe [2]), and provide atmospheric composition tions, and atmospheric sampling throughout descent measurements, most notably noble gas and elemental via pressure and temperature sensors. Finally, a Dop- abundances, and isotopic ratios for hydrogen, carbon, pler wind experiment is also planned. nitrogen, and oxygen. The mission would also charac- Mission Architecture Overview: SPRITE is terize the atmospheric structure (e.g., pressure, temper- planned for launch in November 2024, with a 10-year ature, clouds, and dynamics) at the probe descent loca- cruise to Saturn using an Earth-Venus-Earth-Earth- tion. Saturn gravity assist trajectory. The baseline Carrier- Saturn PRobe Interior and aTmosphere Explorer Relay Spacecraft (CRSC) would be solar-powered and, (SPRITE) [2] is a concept being proposed to fulfill in addition to the Probe, would carry a Multi-Channel these high-priority Decadal Survey science objectives. Imager for context imaging prior to entry. While remote sensing (e.g. Cassini proximal orbits) On November, 2034, SPRITE’s Probe would enter has provided considerable insight into Saturn’s upper- into Saturn’s atmosphere. At a relative entry velocity level atmosphere, in situ measurements are required to of 27 km/s, the Probe experiences a peak heat flux provide ground truth and connection to the observed near 3000 W/cm2, and a peak deceleration up to 45 cloud-top motions. Additionally, the composition of Earth-g. After the entry phase, the Probe jettisons its the deep well-mixed atmosphere including noble gases aeroshell and deploys a parachute to slow the Descent cannot be detected by remote sensing techniques. Vehicle (DV) containing the science insturments to accommodate the science measurement phase. The DV would perform a two-hour science operations phase, in which the QMS, TLS, and ASI take measurements to at least 10 bars pressure. The design of the DV, powered by primary batteries, is fully redundant to ensure low-risk data return. The science data would be relayed to the CRSC, which performs a flyby of Saturn and then transmits the data to Earth via the Deep Space Network. Further details of the mission concept implementation will be discussed in the final paper, including subsystem requirements, design implementation, and predicted mission performance. References: [1] Squyres, et. al., “Visions and Voyages for Planetary Science in the Decade 2013- 2022,” National Academies Press, (2011). Fig. 1 SPRITE would follow a similar entry concept to the Galileo Probe [2] Simon, A., et al, “SPRITE: Saturn mission, and provide in situ sampling of the Saturn atmosphere using two spectrometers and an atmosphere structure instrument package. PRobe Interior and aTmosphere Explorer,” 13th International Planetary Probe Workshop, (2016).

Small Next-generation Atmospheric Probe (SNAP) Concept. K. M. Sayanagi1, R. A. Dillman2, A. A. Simon3, D. H. Atkinson4, M. H. Wong5, T. R. Spilker6, S. Saikia7, J. Li8 D. Hope2, 1Hampton University (Atmospheric and Planetary Sciences Department, 23 E Tyler St. Hampton, VA, 23668, USA, [email protected]), 2NASA Langley Research Center, 3NASA Goddard Space Flight Center, 4Jet Propulsion Laboratory, 5University of California, Berkeley, 6Independent Consultant, 7Purdue University, 8NASA Ames Research Center.

Introduction: We present the Small Next- to multiple planets, we will examine the feasibility, generation Atmospheric Probe (SNAP) as a secondary benefits and impacts of adding SNAP to the Uranus payload concept for future missions to giant planets. As Orbiter and Probe flagship mission. a case study, we examine the advantages, cost and risk Science Targets: Although the current study tar- of adding SNAP to the future Uranus Orbiter and gets the future Uranus Orbiter and Probe flagship mis- Probe flagship mission; in combination with the mission, a probe that is built for Uranus should be viable at sion’s main probe, SNAP would perform atmospheric Neptune and Saturn with few changes, and enables in-situ measurements at a second location. missions such as; Scientific Objectives: The primary objectives to be • Adding SNAP as a second probe to a flagship mis- addressed by the SNAP mission concept are: sion with probe to Uranus or Neptune. • Vertical distribution of cloud-forming molecules. • Adding SNAP as a second probe to the Saturn • Thermal stratification. Probe New Frontiers mission. • Wind speed as a function of depth. • Deploying SNAP at Saturn during a gravity-assist These objectives originate from the recommenda- flyby en route to Uranus or Neptune. tions in the 2013 Planetary Science Decadal Survey for Nature of Expected Science Advancement: The the Uranus Orbiter and Probe flagship mission and the SNAP mission will focus on the physical and chemical Saturn Probe New Frontiers mission. The decadal sur- processes governing the planetary energy balance by vey also prioritized noble gas abundance and isotopic measuring the vertical cloud structure, stratification, ratios; our objectives assume that a future SNAP mis- and winds. Understanding these processes will inform sion will be a secondary probe that accompanies a us how the giant planets have evolved. The atmospher- large probe mission; the noble gas and isotopic ratios ic composition measurements will constrain the abun- are to be measured by the main probe. dance of volatile species on the target planets, which The SNAP objectives represent the unique ad- may clarify how, where, and when the giant planets vances enabled by a SNAP mission: in-situ detection of formed [1, 2, 3]. spatial variabilities in the vertical structure of clouds, A major source of spatial variability is seasonal thermal stratification, and atmospheric dynamics. Plan- forcing. We envision that the SNAP concept will ex- etary Science Decadal Surveys have expressed desires plore a different location than the main probe, so the for multi-probe missions; the 2003 survey advocated mission can examine two hemispheres in different sea- for a Jupiter Multi-Probe mission; and the 2013 survey sons. Uranus represents an especially interesting target emphasized that a second probe that takes measure- to study seasonal variability because the planet’s rota- ment at a separate location can significantly enhance tion axis is tilted ~98º to the orbital plane, imposing a the scientific value of the mission by providing data on strong summer-winter hemispheric dichotomy [4]. If a atmospheric variability. Such missions have not been Uranus mission launches around 2030, the spacecraft realized because their costs are perceived to be pro- should arrive at Uranus around 2040; by then, the north hibitive. We advocate the SNAP concept as a path to- pole will have been basking in continuous sunshine for ward giant-planet multi-probe missions. over 30 years since the equinox of 2007, while the Design Goals:We envision that the science objec- south pole will have been in winter darkness for the tives can be achieved with a 30-kg entry probe ~0.5m same period. Deploying an atmospheric probe into in diameter (less than half the diameter of the Galileo each hemisphere will reveal the effects of seasonal probe) that reaches 5-bar pressure-altitude and returns forcing on the clouds, thermal stratification, and winds. data to Earth via the carrier spacecraft. As the baseline Furthermore, as the winter hemisphere of Uranus al- instruments, the probe will carry an Atmospheric Struc- ways faces away from Earth, the winter side of the ture Instrument (ASI) that measures the temperature, planet can be observed only by visiting spacecraft; this pressure and acceleration, a carbon nanotube-based valuable remote-sensing opportunity can be significant- NanoChem atmospheric composition sensor, and an ly enhanced by an in-situ probe that establishes the Ultra-Stable Oscillator (USO) to conduct a Doppler ground-truth. Wind Experiment (DWE). While SNAP is applicable Another source of spatial variability is atmospheric convection. Uranus has cumulus outbreaks that cause localized clouds that evolve rapidly [5], and meridional circulation that is predicted to vary with seasons [6]; however, the vertical structures of these clouds remain debated. While models of Uranus predict that CH4 and H2S ice clouds condense between 1 and 5 bars [7], remote-sensing retrievals do not agree on the vertical structure of the observed clouds. Karkoschka and To- masko [8] present a diffuse cloud layer across 1-2 bars, while Sromovsky et al. [6] show three compact layers at around 1, 1.5, and 5 bars, which is consistent with Voyager 2 radio occultation data [10]. Furthermore, retrieved thermal stratification of Uranus depends on the poorly known CH4 concentration [11]. An atmospheric probe could resolve these open issues on the atmospheric vertical structure. SNAP exploration of a second location will provide data on spatial variability, and also reduce the risk of sampling an unrepresenta- tive site. The instruments that are available today for the determination of isotopic ratios and noble gas abundances are beyond the payload mass and power appropriate for the proposed SNAP concept. Our proposal will challenge the current state-of-the-art for atmospheric entry probe designs, enable new mission concepts, and identify new instrument and spacecraft systems technologies. By doing so, we will define the performance goals and technology development requirements to achieve decadal objectives with a small probe. References: [1] Mousis, O. et al., (2012). ApJ Let. 751, L7. [2] Walsh, K. J., et al., (2011). Nature 475, 206–209. [3] Hueso, R., and Guillot, T., (2005). A&A 442, 703–725. [4] Friedson, J., and Ingersoll, A. P., (1987). Icarus 69, 135–156. [5] Sayanagi, K. M., et al., (2015). AGU, P41B–2055. [6] Sromovsky, L. A., et al., (2011). Icarus 215, 292–312. [7] de Pater, I., et al., (1991). Icarus 91, 220–233. [8] Karkoschka, E., and Tomasko, M., (2009). Icarus 202, 287–309. [10] Lindal, G. F., et al., (1987). GRL 92, 14987–15001. [11] Orton, G. S., et al. (2014). Icarus 243, 494–513.

DECODING HUYGENS' DESCENT DYNAMICS : SELF-GENERATED AND TURBULENCE-FORCED MOTIONS. R. D. Lorenz1, 1Johns Hopkins Applied Physics Laboratory, Laurel, MD 20723, USA ([email protected])

Introduction: The Huygens probe was the first not explicitly resolved. However, some statistical tech- deep-atmosphere probe with descent imaging, and thus niques [3] allow elucidation of distinctive motions – encountered novel (compared with e.g. Pioneer Venus, e.g. the kurtosis of a sparsely-sampled signal is much Galileo) demands on attitude stability on a parachute- higher for a sinusoidal signal than a Gaussian one, and borne vehicle. A further factor was that the antenna on this basis the tilt sensors indicate that a period of design assumed that tilts of 10o could be allowed, and descent appears to have characteristics of pendulum some basis for asserting that this specification would motion that on terrestrial weather balloons are associ- not be often exceeded was needed. Limited simulation ated with turbulence and sometimes cloud. capability in the 1990s, and a near-total absence of So far, the tilt, Doppler and accelerometer data information on Titan's winds challenged these efforts. have not been analyzed together. Additional datasets Wind Shear: Doppler tracking of the probe yield- have emerged in the years since the probe descent, ed a zonal wind profile. This profile had several re- including an independent Doppler record with higher gions (several km thick) of wind shear. The strongest time resolution, and pointing and angular rate (smear) shear [1] in the free atmosphere was of the order of 5 measurements derived from images. The present effort m/s/km : this may be compared with the pre-mission aims to interpret these datasets to understand the self- specification (95%) of 9.7 m/s/km, suggesting the ra- excited and environmentally-forced probe motions. tionale ([2]a critical Richardson number, derived from Probe Spin: The probe was intended to spin slow- the Voyager 1 radio occultation profile) for specifying ly to pan optical instruments in azimuth, and was this shear was reasonable. equipped with small vanes for this purpose. In fact, the The near-surface Doppler measurements (0-2 m/s) spin behaviour was unexpected : the probe spin rapidly can be interpreted as a time series – the fluctuation declined and reversed. At least part of the story ap- characteristics are well-captured [1] with a simple pears to be previously-unmodelled torque due to the AR(1) autoregressive model (similar to the filtered spin-eject mechanism brackets (hardware that was not, white-noise implementation of Von Karman gust mod- perhaps adequately represented in early drawings to els in terrestrial aviation). appreciate the possible impact). It was observed post- Self-Induced Motions: The generation of sus- hoc in data from the SM2 parachute drop test that a tained motions was suspected from the immediate in- reverse torque was encountered, although the lack of crease in motion associated with the main chute jettis- an initial spin and the approximate equivalence of the son and fast descent under the 'stabilizer' – a 'rough predicted and observed rates meant that the incorrect ride' [3]. This motion died down in the lower atmos- spin direction was not noticed at the time. phere, such that the probe attitude was very stable in Recent wind-tunnel tests with a special low-friction the lowest 10km. air bearing to measure spin torques [5] confirm that the Strong motions were also observed under the stabi- SEPS brackets could counter the spin vanes. However, lizer chute in the SM2 test on Earth, but were dis- this torque alone does not completely explain the spin missed at the time as due to ambient turbulence that rate history, which shows some discontinuous features. was not to be expected on Titan [4]. However, the It is hypothesized that one of the two HASI booms emergence in recent years of sophisticated fluid- failed to completely deploy at first, leading to a strong structure interaction models [5] able to simulate the unbalanced reverse spin torque, as well as unphysical aerodynamic wake of the probe and its effect on the HASI results. flexible parachute has demonstrated that strong oscilla- Conclusions: Several aspects of Huygens' dynam- tions which excite the 'wrist-mode' (or 'scissors mode', ics were not as anticipated; some lessons can be drawn. with the parachute swivel moving antiphase with the References: [1] Lorenz, R., Icarus (submitted). chute and probe) could develop at around 90km alti- [2] Strobel, D. and B. Sicardy (1997), ESA SP-1177. tude, whereas motions in the lower troposphere were [3] Lorenz, R. (2007) Planetary and Space Science, 55, much more subdued – as observed. 1936-1948 . [4] Jaekel, E. et al. (1998) Adv. Space. Decoding the Motion: Because the prominence of Res., 21, 1033-1039 [5] Lingard, S. et al. (2015) 8th wrist-mode motions (~1 Hz) was not anticipated, the European Conf. on Aerothermodynamics for Space sampling rate of tilt sensors (~1 Hz) and accelerome- Vehicles, Lisbon, Portugal. ters was chosen to characterize simple pendulum motions (~0.05-0.1 Hz) so the dynamics of the probe are SYSTEM DESIGN FOR PROXIMITY OPERATIONS OF THE ASTEROID IMPACT MISSION (AIM). I. Gerth1, B. Burmann2, M. Rohrbeck3, M. Scheper4, I. Carnelli5, K. Mellab6, B. Garcia Gutierrez7, V. Pesquita8, A. Pellacani9, and P. Kicman10 [email protected], OHB System AG. Universitätsalle 27-29, 28359 Bremen, Germany [email protected], OHB System AG. Universitätsalle 27-29, 28359 Bremen, Germany [email protected], OHB System AG. Universitätsalle 27-29, 28359 Bremen, Germany [email protected], OHB System AG. Universitätsalle 27-29, 28359 Bremen, Germany [email protected], ESA HQ. 8-10 Rue Mario Nikis, 75015 Paris, France [email protected], ESA ESTEC. Keplerlaan 1, 2201 AZ Noordwijk, The Netherlands [email protected], ESA ESTEC. Keplerlaan 1, 2201 AZ Noordwijk, The Netherlands [email protected], ESA ESTEC. Keplerlaan 1, 2201 AZ Noordwijk, The Netherlands [email protected], GMV ES. Isaac Newton, 11, P.T.M. Tres Cantos, E-28760 Madrid, Spain [email protected], GMV PL. ul. Hrubieszowska 2, 01-209 Varsovia, Poland

Abstract: The Asteroid Impact & Deflection As- The design of AIM strictly follows a low-cost ap- sessment (AIDA) mission is an international collabora- proach and is scheduled for launch no later than 2020. tion of ESA and NASA, with the primary goal to To save costs and make use of flight-proven equip- demonstrate asteroid deflection for the first time by ment, AIM’s primary payload will be the two “AIM determining the impulse transferred by a kinetic im- Framing Cameras” (AFC), which are flight spares of pactor [1]. The impact will be performed by NASA’s the DAWN Framing Cameras. In an additional effort Double-Asteroid Redirect Test (DART) spacecraft, to simplify the spacecraft and thus costs, these cameras currently under investigation at Johns Hopkins Univer- will fulfil a dual role as science and navigation camer- sity Applied Physics laboratory (JHU/APL) with KDP- as, which is unusual for a mission performing close- B in March 2017. The transfer of momentum will be proximity operations. A detailed operational scenario measured by both, ground-based means using radar based on these cameras has been worked out for a observations, as well as a monitoring spacecraft in the number of different operational phases, including: vicinity of the asteroid. The latter will be accomplished • Initial characterization of the asteroid system by ESA’s contribution to AIDA, the Asteroid Impact • Detailed characterization from closer distances Mission (AIM) [2]. This paper presents the key results • MASCOT-2 lander deployment from the ongoing industrial work on AIM, which is • CubeSat deployment currently performed by a consortium led by OHB Sys- As such, AIM will have to fulfil similar operational tem in scope of a bridging phase until the formal ap- requirements as Rosetta (see Figure 2), while being proval of the mission. Figure 1 shows an artist’s im- more constrained by the use of a single camera and pression of the current design. The focus of this contri- flying at the binary asteroid system 65803 Didymos bution will be on the system design elements that are (1996 GT) as compared to the singular comet Chur- required to perform the proximity operations at the yumov Gerasimenko. mission’s target asteroid system “Didymos” [3, 4], which potentially includes the deployment of the MASCOT-2 lander and the release of CubeSats.

Figure 2: AIM spacecraft compared to Rosetta.

This submission will first outline the mission in general and summarize the low-cost design of the spacecraft and operations. After that, the system impacts derived from these aspects will be presented.

Figure 1: Artist’s impression of the DART impact These imapcts derived are in large parts from detailed as monitored by the AIM spacecraft, the GNC analyses that have been performed by GMV in MASCOT-2 lander, and the COPINS CubeSats. scope of the ongoing study and supporting acitvities. Prototype navigation algorithms based on feature • Fixed high-gain antenna, requiring additional tracking have been implemented and tested in a labora- slews and impacting the navigation solution tory environment using a qualification model of the • Limitations for achieving different phase an- AFC. Moreover, simpler algorithms that do not require gles for performing scientific operations an FPGA to enable on-board autonomy have been • Operational concept and system impact of per- studied, such as center-of-brightness or limb-fitting forming a radio-science experiment algorithms that can be used for autonomous attitude These points are considered to be of great interest for pointing. These algorithms have been applied to differ- future and ongoing mission studies and proposals to ent operational aspects, including nominal payload small bodies that might include autonomous opera- operations from different altitudes, lander deployment tions, for instance for deploying micro landers. activities, and for releasing CubeSats. Figure 3 pre- AIM is currently in a re-definition phase after a sents an example of the output generated by the de- negative funding-decision at ESA’s council meeting at tailed simulations of the proximity operations, based ministerial level in December 2016. A descoped base- on ground-in-the loop navigation for both the line is currently being elaborated by the Agency and knowledge and dispersion error. The results show that the industrial team, with hopes of achieving approval for operations below 20 km distance to the asteroid, for this iterated concept by ESA member states until autonomy is mandatory to keep the asteroid in the June 2017. This presentation will therefore also pro- cameras field-of-view. Therefore, and for multiple vide a brief status update on these activities and pre- additional reasons, autonomy is also a necessity for sent the latest concept to the community. any lander-deployment operations, or well-targeted CubeSat releases. References: [1] Cheng A.F. et al. (2015) Acta As- tronautica, 115, 262–269. [2] Michel, P. et al. (2016) Advances in Space Research 57.12, 2529-2547. [3] Kueppers, M. and P. Michel (2016) Lunar and Plane- tary Science Conference 47, 1204. [4] Barnouin, O.S. et al. LPSC 2016 47th Lunar and Planetary Science Conference 47 1427.

Acknowledgments: The work presented in this paper has been carried out in scope of the AIM System Consolidation Phase Part 1 (ESA AO / 1-8790 / 16 / NL / GLC). The consortium led by OHB System AG also comprises its partners QinetiQ Space, GMV, AntwerpSpace, Astronika, and Spin.Works.

Figure 3: Example for achievable ground navigation performance during proximity operations.

A number of key design drivers have been identified as a result of the programmatic and technical constraints in this activity, which will be detailed in this presentation, including: • Omitting a wide-angle navigation-camera, which leads either to the requirement of onboard autonomy or restricts the maximum distance to the asteroid, while at the same time significantly reducing costs, mass, and imaging resolution for observations • Need for autonomous GNC, including FDIR, for releasing a lander • System impacts resulting from release requirements of CubeSats HEDGEHOG HOPPING ROVERS FOR THE EXPLORATION OF SMALL SOLAR SYSTEM BODIES. B. Hockman1, A. Frick2, R. G. Reid2, J. Castillo-Rogez2, I. A. D. Nesnas2, M. Pavone1, 1Stanford University (496 Lomita Mall, Stanford, CA. 94305, {bhockman, pavone}@stanford.edu), 2NASA Jet Propulsion Laboratory.

Introduction: The in-situ exploration of small Solar System bodies, such as asteroids, comets, and irregular moons, has become a central objective for planetary exploration. While some measurements can be obtained remotely, measurements that constrain composition and physical properties require contact with the surface at multiple sites for extended periods of time. Accordingly, controlled surface mobility on small bodies has been identified as a high priority for NASA’s technology development [1]. Several mobility concepts have been proposed for negotiating the microgravity environment on small bodies. Specifically, hopping has been recognized by Autonomous Exploration: Targeted mobility on agencies such as NASA [2], ESA [3], RKA [4], and distant bodies presents additional challenges beyond JAXA [5], as having many advantages over alternative hopping control. It also requires an autonomy stack techniques such as wheels, legs, and thrusters. In fact, that allows the rover operate independently without two hoppers are currently en route to Asteroid Ryugu continuous communication with ground stations— aboard JAXA’s Hayabusa 2 spacecraft: a MASCOT specifically, the ability to localize itself on the surface lander developed by DLR [4] and three MINERVA and plan safe trajectories in a potentially hazardous landers [5], which are both designed to perform small environment. For localization, we propose a hybrid hops, albeit with minimal control. approach that utilizes an onboard IMU and star tracker to coarsely constrain the global position of the rover Hedgehog Rover: At Stanford University, and on-board cameras to provide more precise local NASA’s Jet Propulsion Laboratory, and MIT, we have visual odometry. been developing a minimalistic, hopping rover called The motion planning for hopping rovers is inherent- “Hedgehog,” which aims to extend the concept of in- ly different than planning for wheeled rovers, as they ternal actuation to an architecture that is capable of must contend with uncertain hopping trajectories, ir- controlled mobility (see Fig. 1, and overview video at: regular gravity fields, and chaotic bouncing dynamics. http://youtu.be/bDmoqjNQAu8). Specifically, Hedge- Thus, instead of traditional path planning techniques, hog consists of a (10-50 cm) cube-shaped chassis hous- we cast the planning problem as a sequential (Markov) ing three mutually-orthogonal flywheels and is sur- decision process and utilize high fidelity simulations to rounded by eight “spikes” on the corners to grip the surface. By slowly accelerating the flywheels and then learn safe and efficient control policies using tools abruptly braking them, angular momentum in trans- from model-free reinforcement learning. This “data- ferred to the rover, causing it to rotate and push off driven” approach explicitly accounds for many forms from the surface. By controlling the total angular mo- of uncertainty in the dynamics and has been shown (in mentum of the flywheels, Hedgehog is able to perform simulation) to significantly outperform heuristic con- various maneuvers such as directional hopping (up to trol policies (e.g. hopping directly towards the goal). hundreds of meters), precise “tumbling” (90o pivoting This concept has the potential to lead to small, qua- rotation), and twisting for pose adjustments. si-expendable, and maneuverable rovers that enable a The dynamics and control of these maneuvers has focused, yet compelling set of science objectives been studied in detail, from simplified analytical mod- aligned with interests in planetary science and human els to high fidelity simulations [6]. Experiments in var- exploration. Moreover, this new paradigm of mobility ious test beds have also validated these control laws, for “nanorovers” is highly scalable within typical Cu- including a custom-built 6 DoF gravity-offloading test beSat sizes from 1U to 27U, allowing many of the sub- bed [6] and a successful parabolic flight campaign [7]. systems to be leveraged from interplanetary CubeSats Hopping experiments in reduced gravity test beds sug- being developed at JPL (e.g., C&DH/avionics boards gest a motion accuracy of about 10% (with respect to from NEA Scout, UHF telecom system from INSPIRE, some desired velocity vector) on various types of sur- and electrical power system from MarCO). We present faces, both rigid and granular. a notional mission architecture to Phobos that addresses both high-priority science identified for Mars' moons and strategic knowledge gaps for the future Human exploration in the Martian system. References: [1] National Research Council, Tech. Rep. (2011). [2] P. Fiorini and J. Burdick, Autonomous Robots, 14-2, pp. 239–254 (2003). [3] C. Dietze, S. et al. IAC (2010). [4] R. Z. Sagdeev and A. V. Zakharov, Nature, 341-6243, pp. 581–585 (1989). [5] JAXA, Tech. Rep. (2011). [6] B. Hockman et al., JFR (2016). [7] B. Hockman, et al. ISER (2016).