National Aeronautics and Space Administration

DRAFT Entry, Descent, and Landing Roadmap Technology Area 09

Mark Adler, Chair Michael Wright, Chair Charles Campbell Ian Clark Walt Engelund Tommaso Rivellini

November • 2010 DRAFT This page is intentionally left blank

DRAFT Table of Contents Foreword Executive summary TA09-1 1. Detailed Portfolio Discussion TA09-6 1.1. Aerobraking, Aerocapture, and Entry Systems (AAES) TA09-6 1.1.1. Rigid Thermal Protection Systems (TPS) (cross cutting with TA05-Thermal Management) TA09-8 1.1.2. Flexible Thermal Protection Systems (cross cutting with TA05-Thermal Management) TA09-9 1.1.3. Rigid Aeroshells TA09-10 1.1.4. Deployable Aeroshells TA09-11 1.1.5. Entry Instrumentation and Health Monitoring TA09-11 1.1.6. Entry Modeling and Simulation TA09-12 1.2. Descent TA09-13 1.2.1. Attached Deployable Decelerators TA09-13 1.2.2. Trailing Deployable Decelerators TA09-14 1.2.3. Supersonic Retropropulsion (SRP) TA09-14 1.2.4. Guidance, Navigation, & Control Sensors TA09-15 1.2.5. Descent Device Modeling & Simulation TA09-15 1.3. Landing TA09-16 1.3.1. Touchdown Systems TA09-16 1.3.2. Egress and Deployment TA09-17 1.3.3. Landing Propulsion TA09-17 1.3.4. Landing Guidance, Navigation, and Control for Large Bodies TA09-17 1.3.5. Small Body Systems TA09-18 1.4. Vehicle Systems Technology TA09-18 1.4.1. Separation Systems TA09-19 1.4.2. Vehicle Technology TA09-20 1.4.3. Atmospheric Modeling and Surface Characterization TA09-22 2. Infrastructure TA09-23 3. Immediate Actions TA09-24 Acronyms TA09-25 Acknowledgements TA09-25 References TA09-26

DRAFT Foreword NASA’s integrated technology roadmap, including both technology pull and technology push strategies, considers a wide range of pathways to advance the nation’s current capabilities. The present state of this effort is documented in NASA’s DRAFT Space Technology Roadmap, an integrated set of fourteen technology area roadmaps, recommending the overall technology investment strategy and prioritization of NASA’s space technology activities. This document presents the DRAFT Technology Area 09 input: Entry, Descent, and Landing. NASA developed this DRAFT Space Technology Roadmap for use by the National Research Council (NRC) as an initial point of departure. Through an open process of community engagement, the NRC will gather input, integrate it within the Space Technology Roadmap and provide NASA with recommendations on potential future technology investments. Because it is difficult to predict the wide range of future advances possible in these areas, NASA plans updates to its integrated technology roadmap on a regular basis.

DRAFT Executive summary and reliable combination of systems. Even though For the purposes of this 20-year technology such a mission may be three decades away, there roadmap, shown in Figure 1, entry, descent and will need to be several flight tests in Earth’s atmo- landing (EDL) is defined to encompass the com- sphere and possibly at Mars in order to gain suf- ponents, systems, qualification, and operations to ficient confidence in these systems before relying safely and usefully bring a vehicle from approach on them for a human expedition. Tracing back the conditions to contact with the surface of a solar schedule from human landing on Mars, assuming system body, or to transit the atmosphere of the reasonable development times for the flight tests, body. In addition to landing from space on the and allowing for an occasional failure, it becomes surface of a body with an atmosphere, EDL in- clear that we need to begin such technology de- cludes those missions that enter and then exit the velopments within the next few years in order to atmosphere of a body for aerocapture or aero- enable an early 2040 decade pathway to human braking (just “E”), landing on small or large bod- landings on Mars. Alternate pathways should also ies with no substantial atmosphere (just “L”), be pursued for revolutionary capabilities, with and missions that end in the atmosphere, such as downselects occurring along the way. probes, or that deploy aircraft into the atmosphere New EDL technologies, both revolutionary and (“ED”). This roadmap does not address aircraft or evolutionary, will also enable future robotic mis- aircraft technologies, such as for balloons or pow- sions to other solar system destinations, includ- ered airplanes (see TA04), nor does it address in- ing asteroids, comets, Venus, Mercury, Mars, icy space propulsion preceding atmosphere entry (see moons, the gas giant planets, Titan, and others. TA02). Decadal Survey planning teams have previously EDL is an emergent behavior of a system aris- identified and are currently in the process of up- ing from the combination of hardware, software, dating the set of high priority science goals. These trajectory, operations (including site selection), goals will include missions that enter the Venus and the natural environments being encountered. atmosphere and deliver long lived surface landers, The components of the behavior are highly inter- Mars surface penetrator and network lander con- dependent and can only be validated in the con- cepts, Saturn probes, Titan airships and landers. text of the sequence of events. However the end- They will also include new missions that enter the to-end sequence and associated environments are Gas Giants at high velocities, resulting in radiative difficult or impossible to test as a whole before and convective heating rates that exceed current launch, and so qualification of EDL systems and TPS material capabilities. Strategic EDL technol- operations depend strongly on computer simu- ogy investments, conducted in a coordinated and lations of the entire sequence. These simulations sustained manner, are needed to enable not only need to be grounded in component tests, analy- the current planned set of missions, but also the ses, and modeling. As a result, EDL technology mission sets and science goals that may not be re- developments must also be defined and execut- alizable based on current and near term evolving ed within the context of expected technology ap- technologies. plications, and careful consideration given to the A continuous backdrop of system design and qualification approach that will be used by the analyses will need to be funded for both the hu- mission that first infuses the technology. Further- man and robotic exploration mission sets in order more, different EDL technology developments to provide an evolving assessment framework for may interact strongly with each other in their ap- EDL technology development. As information is plications, and so as a technology matures, the gleaned from ground and flight testing, that in- interacting technologies and their requirements formation will feed back into the studies and in- must be reevaluated. fluence subsequent technology developments and These interactions are most evident for the de- flight demonstrations, and inform the science velopment and evolution of human class EDL for communities of mission feasibility and possibili- Mars. The approach that will be used to land tens ties for the future. of metric tons of payload on Mars is not known Earth testing on the ground and in flight is the today, and cannot be known purely through pa- foundation of EDL technology development and per studies. Only through technology develop- qualification. However, due to the expense of ment can the large uncertainties that exist today these developmental tests, EDL missions to date be reduced in the design of such a mission in or- have relied heavily on the technology developed der to begin to converge on the most cost effective and qualified in the 1960’s and 1970’s with very DRAFT TA09-1 few new developments. We have reached the lim- the natural environments. its of those technologies, and in some cases, have The key recommended areas of technology de- stretched their qualification limits. In order to velopment (in no particular order) are then: move forward, we will need to invest in a vigorous • Low mass TPS for higher entry speeds, and program of ground and flight testing as was com- qualification over a wider range of conditions mon in the 1960’s and 1970’s. • Resilience to and monitoring of TPS MMOD In addition to Earth testing, the science robot- and radiation damage ic, precursor robotic, and human missions to the Moon, Mars, and asteroids in addition to utiliza- • Deployed rigid or flexible drag devices for tion of the ISS can help lay the groundwork for both the entry and descent phases, including future technology developments. It is crucial to flexible TPS development for entry and high- acquire and analyze data on the performance of strength, high-temperature fabrics for descent, these technologies in their flight applications in and supersonic and subsonic parachutes order to enable further development and use in • Lightweight, high-temperature structures for later missions. entry vehicles The key performance characteristics that EDL • Improved entry and descent control authority technology developments will target are delivery/ through higher L/D, control surfaces, performance reliability, cost, delivered mass, land- C.G. modulation, and controllable descent ing site access, and landing precision. Like EDL decelerators subsystems, these characteristics interact with • Supersonic retropropulsion systems each other. Reliability is manifested in the com- • Deep-throttling terminal descent engines pleteness of the tests and analyses of component • Landing gear for very large payloads, and for technologies, such as thermal protection systems, small payloads on extremely hazardous terrain deployable decelerators, landing hazard tolerance, and separation systems. Reliability is also realized • High-G landed avionics, power systems, and with increased timeline to complete events as a re- sensors sult of larger drag devices applied earlier, preci- • Mitigation of surface modification and hazard sion landing (reliant on detailed site information creation by terminal descent propulsion for a priori hazard identification), hazard avoid- • Advanced sensing (terrain tracking and hazard ance, and the mitigation of site hazards created by detection) and guidance for terminal descent terminal descent propulsion. Low cost is enabled • Small body descent, touch, hover, ascent, fault by improved simulation and ground-to-flight ex- protection, and sustained contact methods trapolation, and by incorporating high-G land- • Staging systems for deployment, separation, ed systems into mission architectures. Delivered and recontact avoidance specific to each mass can be increased or enabled through the use component, and the development of advanced of more capable TPS for the more difficult envi- pyrotechnic initiation devices ronments presented by larger entry vehicles, larg- • Flight instrumentation of key performance er drag devices applied at higher speeds, descent- aspects of EDL systems to guide subsequent phase (supersonic) retropropulsion, and more applications and technology developments efficient terminal descent propulsion. Landing site access can be increased through TPS that per- • Improved modeling of all EDL systems to mits higher entry speeds (allowing a wider range reduce the number of parallel technology paths, of targets), small body proximity operations, in- project qualification costs, reduce margins, and creased altitude performance, again by virtue of increase reliability higher drag earlier than can be provided current- • Increased knowledge and improved models of ly, and higher precision allowing a wider range of the atmospheric and surface environments, both safe sites. Greater control authority also enables a priori and during EDL (using instruments on higher precision in the entry phase. Both preci- the vehicle or on leading probes) sion landing and hazard avoidance are enabled by In order to support these technology develop- a combination of more advanced terrain sensing ments, we will need: and algorithms with more capable terminal de- • More refined design concepts and architectures scent propulsion and guidance to divert the land- for large scale exploration missions to provide a er to the desired target. All of the objectives bene- framework for related technology developments fit from improved modeling of the systems and of • Maintenance and enhancement of key ground

TA09-2 DRAFT Figure 1: Entry, Descent, and Landing Technology Area Strategic Roadmap (TASR)

DRAFT TA09–3/4 This page is intentionally left blank test facilities in NASA and other government mented by the SkyCrane touchdown delivery Agencies, and the development of new test system, and will deliver approximately 1 metric capabilities at low and high altitudes and ton (t) of surface payload. Current estimates on speeds in Earth’s atmosphere the extensibility of the MSL architecture indicate • A steady pace of EDL technology developments that it is limited to roughly 1.5 t delivered mass, and flight tests as well as missions employing without significant investments in new technol- EDL segments to develop and grow the ogies. Human scale Mars missions, the ultimate experienced workforce that will be needed to goal in NASA’s human space exploration plans, successfully land humans on Mars and other will require 20-60 t of landed payload mass [1]. destinations beyond LEO This represents more than an order of magnitude • A robust STEM student outreach program increase in mass compared to MSL derived ca- that attracts America’s youth and engages them pabilities, an increase in capability on the order actively in EDL technology development of the jump between Apollo and the Space Shut- This roadmap’s most direct impact will be on ef- tle Orbiter. Human Mars missions of 20-60 t of forts initiated in the next five years, and it is antici- landed payload mass would necessitate Mars ar- pated that the roadmap will be regularly reviewed/ rival masses on order of 50-150 t. If NASA wants revised to reflect changing agency priorities. Sec- to send humans to the surface of Mars, sustained tion 3 of this document lists a set of recommend- and coordinated investments over a period of de- ed immediate actions. cades in new EDL system technologies must be made. Given that the probability of loss of mis- Benefits and Impacts sion during EDL tends to be comparable to that NASA investments in fundamental atmospher- during launch, it is imperative that technology ic flight and AEDL technologies in the 1960’s and developments in EDL be motivated by a mind- 1970’s have enabled many of our current EDL ca- set of enabling a mission by providing robust, re- pabilities of today. For example, the current state liable, and Earth-testable solutions. of the art for human scale Earth entry is defined Mission enabling EDL technology require- by the Shuttle Orbiter (1970’s). NASA’s ability to ments are not limited to human or Mars ex- land robotic payloads on the surface of Mars is ploration. Large-scale Earth return from hyper- largely reliant on the EDL technology set devel- bolic velocities will require thermal protection oped for the Mars Viking Program (1970’s) and materials well beyond those under development utilized in large part on all of the robotic Mars for Orion. Hyperbolic velocities may exceed 13 landers since. TPS technologies developed for km/s for some NEO and Mars return opportu- Apollo and the Space Shuttle in that same time- nities, and will include significant radiative and frame are being recycled or requalified today for convective heating. The availability of suitable current human spacecraft concepts. NASA’s pio- TPS materials has limited robotic mission plan- neering entry missions to Venus (Pioneer) and the ning for Giant Planet probes (due to high heat- Giant Planets (Galileo Jupiter) were also designed ing and pressure encountered) and Mars Sample in the 1970’s, and current SMD proposal con- Return (due to planetary protection reliability re- cepts are largely reliant on derivative technologies quirements). Autonomous landing operations at from those missions for current mission planning. or near the surface of small bodies with low grav- Currently, advances in EDL capabilities are gen- ity (asteroids and comets) will require advanced erally driven by individual mission performance capabilities in terrain recognition, guidance and requirements, constrained budgets, and near-term surface anchoring, particularly for small bodies schedule demands, and often require high TRL, where the surface may be icy or highly fractured. low-risk technologies for mission infusion. There Human lunar or NEO exploration will require is a large reliance on heritage technology, even to hazard avoidance, precision landing, deep throt- the point where technology limitations are effec- tle descent propulsion and lightweight landing tively constraining science objectives for many attenuation systems. Giant planet (and possibly SMD missions. Venus) exploration will require, at a minimum, The Mars Science Laboratory (MSL), NASA’s requalification of out-of-production TPS mate- flagship Mars mission scheduled to launch in rials. New classes of advanced science missions, 2011, defines the current SOA for Mars EDL sys- including long-lived Venus landers, robotic air- tems. MSL is using much of the same Viking-de- ships or airplanes at Mars, Venus, or Titan, new rived EDL technologies and architecture, aug- missions that enter and explore the atmospheres

DRAFT TA09-5 of the gas giants, and Europa landers, may be en- er agency planning documents, the current SMD abled by coordinated and strategic investments in decadal survey [4], and other sources. SMD is in “push” technologies. Key push technologies in- the process of updating their decadal survey now; clude dramatic reductions in TPS mass (Giant this roadmap will be amended once those data be- Planet), and low ballistic coefficient deployable come available. decelerators (all destinations). Revolutionary ad- Technology Area Breakdown Structure (TABS) vances in propulsion (e.g. nanoenergetic fuels) The Technology Breakdown Structure is shown may enable new EDL concepts to augment aero- in Figure 2. dynamic drag, and potentially to harvest drag en- ergy during entry. 1. Detailed Portfolio Discussion In general the benefits of focused EDL technol- The recommended technology investment port- ogy activities include: folio is discussed by TABS element, including a • Increased mass delivery to a planet surface (or brief assessment of the current SOA and recom- deployment altitude), mended mission pull and technology push invest- • Increased planet surface access (both higher ments. Push investments are highlighted in blue. elevation and latitudes), 1.1. Aerobraking, Aerocapture, and Entry • Increased delivery precision to the planet’s Systems (AAES) surface, AAES are defined as the intra-atmospheric tech- • Expanded EDL timeline to accomplish critical nologies that decelerate a spacecraft from hyper- events, bolic arrival through the hypersonic phase of en- • Increased robustness of landing system to try. Over the next 20+ years, NASA mission surface hazards, objectives will require significant advances to the • Enhanced safety and probability of mission state-of-the-art in AAES in the following areas: success for EDL phases of atmospheric flight, higher entry speeds (crew/sample return from be- • Human safety during return from missions yond LEO), larger entry systems for human ex- beyond LEO, and ploration, extreme environment systems for Ve- • Sample return reliability and planetary nus and Giant Planet exploration, high reliability protection. systems for human and sample return missions, Low TRL EDL technology advancements, from and low cost COTS access to space. The unique the laboratory through development, qualification challenges of entering large systems (>1 t) pay- and flight test, also provide fertile training ground loads at Mars will require revolutionary chang- for young systems engineers and technical work- es to the current SOA. Other mission classes will force. In the 1950’s and 1960’s, the plethora of benefit, or in some cases be enabled, by evolution- aircraft and aerospace flight projects served to de- ary or revolutionary improvements to the current velop a cadre of engineering and technical talent SOA. In order to support the required advances in that largely enabled the successful Apollo program this area, the team has identified six key technolo- to land men on the Moon. In addition, other na- gy investment areas (product lines) in AAES. Each tions are making strategic and concerted invest- product line will be discussed below. ments in EDL technologies which will dimin- Major challenges for all AAES technologies in- ish our technological advantage without focused clude a wide range of mission requirements that NASA investment. For our nation to continue to necessitate multiple parallel investments. These in- lead, future human solar system exploration and clude the availability and suitability of ground fa- EDL systems capabilities will require investment cilities for testing and model validation, our abili- in and development of the young technical work- ty to human rate the developed systems, long term force, many of whom may only be in grade school retention of critical EDL-unique skills (particular- today. ly for ablative TPS) in the NASA workforce, and the requirement of system level validation of many Mission Pull technologies via flight test. Flight demonstration The following missions (shown in Table 1) rel- of aerocapture may facilitate adoption by mission evant to EDL and associated dates were assumed planners, but was not considered to be a major for this roadmap. They were derived from the technical challenge. Aerobraking, while a critical Agency Mission Planning Manifest [2] a Human capability for some NASA missions, was not de- Exploration Framework Team briefing [3], oth- termined to have significant technical capability

TA09-6 DRAFT Table 1. Potential missions related to EDL technologies and capabilities

Mission Launch Critical EDL Capabilities Comments Crewed orbital velocity return 2015 Large-scale Earth EDL NASA and/or commercial vehicles Lunar sample return Competed opportunities in Lunar landing / Earth EDL of sample New Frontiers 2017/2022/2027 Asteroid sample return Competed opportunities in Asteroid touch and go, proximity New Frontiers 2017/2022/2027 operations / Earth EDL of sample Venus lander Competed opportunities in Extreme environment TPS New Frontiers 2017/2022/2027 Saturn probes Competed opportunities in Low Mass Extreme environment TPS New Frontiers 2017/2022/2027 Mars sample return sample 2018 1-2 t class Mars EDL Advanced from MSL acquisition ISS down-mass capability 2018 Low cost TPS, deployable HIAD testing and application decelerators Crewed high-velocity Earth return 2020 Low mass TPS From HEO in preparation for asteroid mission Mars sample return orbiter 2022 High reliability TPS and SRC Planetary protection requirements (planetary protection) Mars sample return surface 2024 Precision landing, deployable Meets up with sample cache left on rendezvous decelerators surface Crewed asteroid rendezvous and 2025 High-q low mass reliable TPS Low gravity proximity ops, hover, return touch, land Mars network 2029 Guidance, small low cost SRC Seismology Titan aerial vehicle, landers/ 2032 Titan entry, descent and deploy, Flagship mission splashers Titan EDL Crewed Mars orbiter 2035 Aerocapture, possible crewed Phobos landing: Mid L/D or Large Deployable Decelerator for Aerocapture Crewed Mars surface 2041 Mars large EDL: SRP, Mid L/D or large ~30 metric tons lander Deployable Decelerator Icy moon lander 2042 Icy moon EDL e.g., Europa, Enceladus

Figure 2. Entry, Descent, and Landing Systems Technology Breakdown Structure DRAFT TA09-7 gaps other than higher temperature capable solar nus and Galileo missions as a result of a lack of panels, an advance carried in the In-Space Power suppliers for the critical precursor Rayon mate- (TA02) roadmap. A roadmap has been developed rials and the absence of a U.S. based carboniza- which will overcome most challenges through sus- tion processing capability that is required to pro- tained investment in the technology area, with duce the entry grade TPS. Advances are required time-phased milestones to support mission driv- to significantly lower the areal mass of TPS con- ers, as well as development of key push technolo- cepts, demonstrate extreme environment capabil- gies that may significantly enhance future NASA ity, demonstrate high reliability, demonstrate im- mission capabilities or enable new mission class- proved manufacturing consistency and lower cost, es. The limitations of existing ground test capa- manufacture larger integrated aeroshells in a cost- bilities can only be addressed by the extension of effective manner, and demonstrate dual-heat pulse current arc-jet capabilities for Venus, Giant Plan- (aerocapture plus entry) capability. These advanc- et, and human hyperbolic return capability. Tech- es will have long lead times due to the need to nical challenges specific to each AAES investment understand complex nonlinear performance and area will be discussed below. failure modes. Experience has shown that a large Finally, it should be noted that TPS technology fraction of aeroshell risk and DDT&E costs are is also integral to Thermal Management Systems. due to integration and closeout issues. However, Entry TPS technologies identified under this TA these issues are implementation-specific engineer- are consistent with key technologies identified ing problems, not technologies per se. Therefore, during the Thermal Management (TA05) tech- while very important, they are not included in this nology roadmapping process and have been ful- roadmap. Current agency investments include ab- ly coordinated with that team. Several cross-cut- lative materials development within ARMD (Hy- ting technology areas, including reusable acreage personics) and ESMD (Orion & ETDD), primar- and WLE TPS, heat pipes, in-space repair, self- ily in support of crewed return from the Moon healing TPS, and multifunctional combined ther- and Exploration missions to Mars. Investments in mal and integrated load bearing structural TPS SMD (In-Space Propulsion) advanced the TRL of are discussed in detail in the Thermal Manage- several new ablators to 3-4 and are beginning to ment Roadmap, while the core Entry technolo- explore the challenges of Carbon Phenolic rede- gies of rigid and flexible TPS are discussed in both velopment. These investments have already dem- roadmaps. onstrated the potential to reduce TPS areal mass 1.1.1. Rigid Thermal Protection Systems by ~40% for these applications via dual-layer or (TPS) (cross cutting with TA05- graded design [5]. Larger mass savings may be Thermal Management) possible via push investment in tailored materials that reject a component of incident heating (radi- Previous vehicles used TPS installed on a rig- ation or catalysis), tailor material properties as a id aeroshell/structure, ranging from the reusable function of depth, or employ new reinforcement tiles on the Orbiter to ablative systems employed or impregnant concepts. However, there is cur- for planetary entry and Earth return from be- rently minimal investment in materials concepts yond LEO. For many exploration missions, such for other mission classes and highly innovative as near-Earth asteroid and Mars missions, abla- multifunctional materials. Notably, as a result of tive materials are an enabling technology need- the lack of Carbon Phenolic material supply chain ed for dual heat pulse reentries and for high ve- issues noted above, missions that involve Earth re- locity entries (>8 km/s). However, the current turn of crew from beyond the Moon, or robotic selection of high TRL rigid TPS materials is in- entry to Venus or the Giant Planets, have no avail- adequate for future mission objectives, due to ei- able qualified TPS solution at this time. Further ther insufficient thermal performance or high ar- exploration of Jupiter will require significant re- eal mass. Only a limited number of materials, duction in areal mass over Carbon Phenolic in or- PICA (Stardust, MSL), Avcoat (Apollo, Orion), der to avoid entering probes with a TPS mass frac- Carbon Phenolic (Pioneer Venus, Galileo), SLA- tion approaching 1.0. Other mission classes must 561V (MER, Phoenix) and ACC (Genesis), have resort to capable, but extremely heavy, TPS solu- been used for previous or upcoming, exploration tions. Entry to Titan, robotic entry to Mars, and missions. The U.S. no longer possesses the capa- LEO return missions can be accomplished with bility to manufacture the flight qualified Carbon existing high TRL TPS, although evolutionary ad- Phenolic TPS material used for the Pioneer-Ve- vances may have mission benefit. Push technol-

TA09-8 DRAFT ogies, such as TPS materials that reflect incident • Multifunctional materials designed for thermal shock layer radiation, or materials that are chem- protection as well as improved MMOD and ically designed to self-heal or affect boundary lay- radiation protection (covered in the Thermal er modification (such as delay of transition), are Management Roadmap) currently very low TRL, but have the potential to • Improved thermal response models, including significantly enhance future mission performance. high fidelity ab initio in-depth ablation Recent efforts under Constellation have revived response, gas-surface interactions, and ablation analysis capabilities. TPS development computational materials design capability. efforts should be expanded to include material re- (overlap with Materials & Structures) sponse/flow field coupling codes (for both engi- • Improved processes for quantification of neering calculations and high-fidelity solutions), TPS margin and system reliability, including integration of ablation models into standard 3-di- statistical analysis, testing techniques, and mensional thermal modeling codes, and ground archival storage of agency thermal test data, testing to generate data for code correlation and as required for COTS and NASA crewed validation. The parallel development of high-fi- vehicles as well as high reliability sample return delity material response models is particularly im- missions. portant for entry TPS, where a “test as you fly” approach to system validation cannot be applied, • Improved manufacturing and inspection and therefore the designer is dependent on sim- processes, which should include the use of ulation capability in order to define the baseline automated, robotic systems especially for system as well as the necessary margin in order coating/primer application and material to meet reliability requirements. There is some installation on structure. TPS investment within other government Agen- 1.1.2. Flexible Thermal Protection Systems cies and industry, with a primary focus is on re- (cross cutting with TA05-Thermal usable systems to support hypersonic cruise. Ma- Management) jor technical challenges include the development System studies have shown that large entry de- of fundamentally new material concepts, fidelity celerators provide a potentially enabling means of current response models, availability of suitable to increase landed mass on the Martian surface, ground test facilities, ground to flight traceability, as well as the consideration of a whole new class high uncertainties in input aerothermal environ- of missions. Flexible TPS is an enabling compo- ments (covered in Section 1.1.6) and the inher- nent for deployable entry systems (Section 1.1.4). ent conflict between low mass and robust perfor- In addition, current research [5] indicates that mance. Concept maturation to TRL 5 will require high performance flexible materials may provide extensive ground testing, while maturation to strongly enhancing benefits for rigid systems as TRL 6 may require a small-scale component level well (in terms of reduced life cycle cost and ease flight test. Primary areas of recommended NASA of manufacturing, as evidenced by Orbiter LCC investment are summarized below: data). In addition, flexible TPS applied to rigid • Advanced Ablator materials aeroshell structure should significantly mitigate »»Mid-density ablators for Earth return design issues such as cracking, thermostructur- from beyond LEO, Venus, and Mars entry al performance, bond verification, and scalabil- missions, including dual heat pulse capable ity to large aeroshells. These advances may pro- materials with at least 40 percent lower vide game-changing benefits to multiple current areal mass than the current SOA (overlap and future NASA missions, as well as COTS Car- with Thermal Management, Materials & go and/or Crew return concepts. The current Structures). SOA is the AFRSI blankets employed on the lee- »»Extreme environment (q > 2 kW/cm2, p > ward side of the Orbiter, which are reusable sys- tems designed for ~5 W/cm2 of maximum heat- 1 atm) materials, including redevelopment ing. Future deployable entry systems will require of extremely high-reliability (using heritage TPS concepts that can be stowed for months in Rayon) Carbon-Phenolic for Mars Sample space and then deployed into an entry configura- Return, and alternate materials for future tion that can withstand 20-150 W/cm2 of heat- missions to Venus and the Giant Planets. ing at Mars or Earth. These are envisioned as sin- »»Development of radiation reflective and gle or dual (entry plus aerocapture) use systems. smart or self-healing ablator systems Both non-ablating and ablating concepts may be

DRAFT TA09-9 suitable, where the key trade is TPS development Mars, 60° at Earth, 45° at Venus and Giant Plan- complexity versus system scalability and control- ets) and the truncated sphere (Apollo and Orion). lability. Non ablating concepts will either be mul- Typically these aeroshells are either metallic or tilayer insulative systems, or possibly transpiration composite, with the TPS bonded to the structure cooled fabrics. Ablative systems may include or- via an adhesive. The carrier structure is designed ganic resins as impregnants or as woven fibers. Ad- to bear all aerodynamic loading (without reliance vanced fiber weaving techniques can be employed on the TPS). Control of these vehicles has been ef- to tailor material properties for given mission re- fected via Reaction Control Thrusters in most cas- quirements. The current NASA portfolio includes es (aerodynamic control surfaces are of course em- small investments in q<40 W/cm2 non-ablative ployed on the Orbiter). There has been very little materials (ARMD Hypersonics), and q<150 W/ effort within NASA to develop new entry aero- cm2 ablative materials (ESMD ETDD). All of shells or aeroshell technologies beyond the con- these materials are low TRL at this time, although ceptual stage, primarily due to the lack of a mis- a TRL 4 insulative 20+ W/cm2 system is planned sion driver. However, one class of mission, human to be flight tested in 2012 on IRVE-3. As a push exploration of Mars, will require a fundamentally technology, the extension of flexible systems to new aeroshell design due to large landed mass re- higher heat flux (q > 400 W/cm2) may have game quirements. In addition, optimized aeroshells for changing benefits for a variety of proposed NASA specific mission classes to other destinations may missions. Finally, it should be possible to tai- provide evolutionary advances in current mission lor the surface chemistry of these systems via im- capabilities, and self-righting highly stable designs pregnants or additives in order to reject heating are desirable for sample return missions. Current from surface catalysis and or shock layer radiation, NASA investment in alternate aeroshell designs which may provide strongly enhancing mass sav- is minimal beyond the conceptual stage, and in- ings for all NASA missions. Major technical chal- cludes an Apollo derived aeroshell for Mars 2018, lenges include maintaining thermal and structur- and a stable parachuteless capsule for Mars Sam- al properties after long duration storage in space, ple Return (SMD – ISPT). Previous studies in- packaging efficiency, performance under aeroelas- cluded mid lift/drag (L/D) (biconic or ellipsled) tic and shear loading, and ease of handling. Con- designs for Mars entry and Neptune aerocapture, cept maturation to TRL 5 will require extensive and raked blunt cones for crew return and orbit- ground testing, while maturation to TRL 6 may al transfer at Earth. Mid L/D vehicles could be require a small-scale component level flight test. designed for dual purpose use as a payload fair- Areas of recommended NASA investment are ing during ascent. Even for “heritage” aeroshell summarized below: shapes, the current state of the art in several sup- • Non-ablative (insulative or transpiration porting disciplines, notably aerothermodynam- cooled) concepts with high flexibility & ics, has large uncertainties leading to large design stowability, margins and mass-inefficient entry systems (cov- • Ablative concepts, including systems that ered in Section 1.1.6). Non-NASA investment in rigidize in-space or during entry. Includes such systems has been restricted to slender cones high-heat flux semi-flexibles (q > 150 W/cm2) for RV applications and mid to high lift hyper- for low-cost application to rigid aeroshells, sonic cruise vehicles, which are minimally applica- • Multifunctional materials that provide ble to proposed NASA missions. Major technical MMOD/radiation protection and/or carry challenges include developing lightweight struc- structural loads, tures, effecting non-propulsive control for low to mid L/D bodies, and the development of high fi- • Non-catalytic coatings to reduce aeroheating delity aero/aerothermal databases, including dy- environments and additives to reflect radiant namic stability. In addition, push investment in heat, advanced guidance algorithms (such as numeri- • Improved thermal response and reliability cal predictor/corrector) is required for aerocapture quantification models. and could enable precision landing of entry vehi- 1.1.3. Rigid Aeroshells cles, including optimal divert for proximity oper- All NASA entry vehicles to date have employed ations of multiple landed assets. Concept matu- rigid aeroshells, with almost all concepts in one ration to TRL 6 will require a mix of ground and of two classes: low-L/D, and mid-L/D (Orbiter). subscale system level flight tests. Areas of recom- The current SOA is the classic sphere-cone (70° at mended NASA investment include: TA09-10 DRAFT • Highly reliable sample return capsules that ly despite having been identified as attractive in meet planetary protection requirements, low multiple system analysis studies. Major techni- cost aeroshells (50% reduction over SOA), cal challenges include scalability, reliable deploy- and performance optimized low lift aeroshell ment, aeroelastic and aerothermoelastic effects, designs, and aerodynamic stability and controllability of • Performance optimized lifting bodies (L/D ~ large deployable or flexible structures. Concept 0.6) for high mass entry and aerocapture, maturation to TRL 6 will require a mix of ground • High lift (L/D > 2.0) maneuverable bodies for and subscale system level flight tests. Primary ar- aerogravity assist, eas of recommended NASA investment include: • Dual use launch shroud/payload fairings and • Attached or towed inflatable entry systems entry aeroshell systems employing advanced flexible TPS technology, • Advanced lightweight structures, including • Mechanically deployed and/or on-orbit warm/hot structures/adhesives and integrated assembled entry systems, structural/ TPS systems that reduce total • Transformable or morphable entry systems for aeroshell mass by at least 40% over current use during descent and or landing, SOA for multiple mission classes (overlap with • Advanced high temperature flexible structural Materials & Structures), materials, including bladders, ribs, and • Non-propulsive flight control effectors, rigidizable concepts that can reduce structural including control surfaces and active c.g. mass over current SOA (overlap with materials modulation, & structures), • Advanced guidance and navigation systems • Non-propulsive flight control effectors, for aerocapture and entry, including precision including control surfaces and active c.g., guidance for multi-probe missions. modulation; methods must account for and be consistent with potential system flexibility, 1.1.4. Deployable Aeroshells Deployable entry systems provide a means by • Advanced guidance and navigation systems which the ballistic coefficient at entry is relatively adapted to deployable system controllers, unconstrained by launch shroud limitations. This • Advanced fluid structure interaction modeling revolutionary advance has several potentially en- sufficient to predict and mitigate aeroelastic abling benefits, particularly for large payload de- and aerothermoelastic effects, including livery to the Martian surface, as well as the po- dynamic buckling, on material and system tential to significantly enhance a variety of NASA performance, and and COTS missions ranging from International • Advanced aerobraking concepts including Space Station (ISS) downmass to crewed Earth re- autonomous control methods that can greatly turn from beyond LEO. No truly deployable en- reduce “human in the loop” costs and reduce try system has been tested at flight-relevant scale the risk of planetary aerobraking. (other than for aerobraking, which is performed 1.1.5. Entry Instrumentation and Health using deployed solar panels), but multiple small- Monitoring scale flight tests have been conducted (most re- Entry instrumentation for both engineering data cently IRVE-II in 2009) with varying degrees of and vehicle health monitoring provides a critical success. NASA has significant investment plans link between predicted and observed performance in this technology area, including IRVE and sup- of the AAES system, and is crucial for improving porting materials and analysis research (ARMD the design of current systems, and for ensuring Hypersonics & ESMD ETDD) and a “fast-start” sufficient system reliability prior to deployment or hypersonic inflatable aerodynamic decelerator use. In addition, entry data can also enhance or program sponsored by OCT. Similar investments enable scientific return from missions, as with the have been sponsored by other government Agen- recession sensors on the Galileo probe, which were cies and industry for applications such as payload used to improve knowledge of vehicle drag as a recovery and stage separation, although primarily function of time as part of the atmospheric recon- in the supersonic regime. However, the only de- struction experiment. Advanced health monitor- ployable entry concept that has seen significant ing instrumentation can have strongly enhancing development to date is inflatables; investment in benefits for missions that require high reliabili- mechanically deployed systems lags significant- ty by ensuring that the entry system is functional

DRAFT TA09-11 prior to use. The current SOA consists of thermo- used to extrapolate ground test results to predict couples, recession sensors, and pressure ports em- flight performance. The current NASA SOA in bedded into the TPS material (such as will fly on entry system modeling ranges from good (flight MSL in 2011), and acoustic impact detection sys- mechanics and 6-DOF trajectory) to fair (aero- tems used on the Shuttle wing leading edge. There thermodynamics and fluid-structure interactions) is little current NASA investment in this area to poor (dynamic aerodynamics). Most analyses other than through a few research grants with- are conducted in an uncoupled fashion – multi- in ARMD/FAP. In addition, there is some devel- disciplinary tools are still at the cutting edge in opment of wireless systems within Constellation this field. In many cases, particularly aerothermo- (CUIP), but that is slated to end in 2011. Recent dynamics, the sophistication of the computation- investments in the SMD ISPT led to the devel- al software outpaces the level of validation; key opment of advanced recession sensors that will fly gaps remain in validation of these codes at flight- on MSL in 2011. Major technical challenges in like conditions (e.g., high enthalpy, high Reynolds entry instrumentation include high temperature number, correct gas composition). Well designed systems capable of direct heat flux measurements, ground tests are a critical component of simula- in situ measurements (temperature and strain) in tion validation. NASA currently supports some flexible TPS, advanced optical and other non-in- core investment in modeling and simulation (pri- trusive measurement techniques, and shock lay- marily in the Aerosciences) in ARMD (Hyperson- er radiation measurements in ablative TPS. Chal- ics) and ESMD (Orion and ETDD), as well as lenges in health monitoring include development some validation activities in the NASA Engineer- of low data low power networks, elimination of ing and Safety Center (NESC) (RCS-aero inter- false positives, and the ability to initiate and mon- actions). Flight missions, such as MSL, typically itor repair of detected damage. Primary areas of invest in the development of mission-specific da- recommended NASA investment include: tabases, but do not support investment in advanc- • Sustained investment in improved intrusive es to the state of the art. Modeling and simulation, entry instrumentation (e.g., faster response and its validation, is an area where a sustained in- times, higher thermal capability, compact vestment in core technical disciplines is essential low mass pressure measurements, direct to maintaining cutting-edge design practices for measurements of heat flux, catalycity), and future missions. In the Aerosciences, specific ad- development of systems for the measurement vances are required in the area of higher order tur- of shock layer radiation, bulence modeling (such as DNS methods), fully • In-situ instrumentation for flexible/deployable unstructured or gridless CFD approaches for hy- TPS systems, personic flow, improved methods for low-density • Semi or non-intrusive instrumentation flows (such as Boltzmann solvers), and higher fi- concepts, including wireless (data and power), delity models for non-equilibrium high tempera- electromagnetic, and acoustic based systems, ture physical phenomena. Next generation NASA missions will rely on larger, heavier entry systems, • Advanced distributed systems for MMOD which will place increased emphasis on improved detection, understanding of turbulent heating, transition to • Advanced ISHM systems, including turbulence, shock layer radiation, and complex multifunctional response and the ability for surface-chemistry interactions. These phenome- closed-loop initiation and monitoring of in situ na must be modeled in the context of a hyperson- repairs, and ic chemically reacting environment, which places • Remote observation platforms for Earth additional constraints on the methods employed. entries. Many improvements to the state of the art for low 1.1.6. Entry Modeling and Simulation speed flows are not applicable to entry systems be- Often overlooked in technology roadmaps, cause they cannot accurately capture the embed- modeling and simulation capabilities, including ded strong shocks that are so prominent in this experimental validation, are a lynchpin of modern flight regime. Note that atmosphere modeling is AAES design. Ground test limitations preclude a included in Section 1.4.5. Primary areas of recom- “test as you fly” approach to AAES, and flight tests mended NASA investment include: are prohibitively expensive for most missions. As • Development of validated multi-disciplinary a consequence, validated high-fidelity models are coupled analysis tools (bridging aerosciences, flight mechanics, structural, and thermal TA09-12 DRAFT analysis), particularly for high reliability and parachutes has demonstrated difficulties in extrap- extreme environment applications, olating deployment and steady state behaviors be- • Sustained investment in improvements in yond qualified scales. Addressing the uncertainties aerodynamics and aerothermodynamics associated with the use of large-scale decelerators modeling, including shock layer radiation introduces the need to test at near full scale, or and high enthalpy ionized turbulent and the need to develop test methodologies that re- separated flows, across the continuum and duce the dependence on testing at scale. Qualifi- non-continuum flight regimes. Includes a cation testing at the needed scales and conditions strong focus on flight relevant experimental is generally beyond the affordability of a flight validation, program, inhibiting the use of anything but “her- • Sustained investment in the underlying itage” systems. Thus, it is important that technol- numerical methodologies and techniques, ogy development programs not only test at appre- taking advantage of expected computer ciable scales but also develop strategies for flight architecture and hardware improvements programs to qualify the technology at larger sizes (overlap with Modeling & Simulation). and more stringent test conditions. 1.2. Descent 1.2.1. Attached Deployable Decelerators Descent subsystems and technologies are de- Large increases in the drag area of an entry ve- fined as those that bridge the hypersonic portion hicle can be achieved through the use of deploy- of the entry sequence with the terminal phase of able decelerators. These devices differ from the landing. The presence of an atmosphere is inher- entry variant (Section 1.1.4) in that they are de- ently assumed. Descent is generally considered ployed endo-atmospherically after the peak heat- to include flight through supersonic and high- ing and peak deceleration phases of flight. As a re- subsonic conditions. Initiation is predicated on sult, the thermal and structural environments are a staging event such as a parachute deployment considerably less severe, although with the added that may not exist in every mission sequence. De- complexity of a dynamic deployment event. At- scent ends with the initiation of a terminal de- tached decelerators can provide order of magni- scent propulsion system or landing. Historically, tude increases in drag area at Mach numbers and the descent phase of flight has focused on simply dynamic pressures considerably higher than cur- providing sufficient deceleration to stage to a par- rent supersonic decelerators, which in turn en- ticular landing system, and thus the primary tech- ables increased timeline margin and/or increased nology for this phase of EDL has been the para- mass delivery to higher elevation landing sites. chute. As planetary missions move towards larger The primary application of attached deployable payloads with greater emphasis on targeted land- decelerators may be at Mars due to its tenuous at- ings, the SOA in descent technology will require mosphere, but other applications, including ISS major advances. These advances primarily focus downmass and the landing of large payloads on on providing greater deceleration in the superson- other atmosphere bearing bodies, are possible. At- ic and subsonic regimes in a manner that does not tached decelerators can be further categorized as reduce landing accuracy or result in transient un- flexible (e.g. SIADs) or rigid. Attached inflatable steadiness or loss of performance in the transon- decelerators were originally conceived during de- ic regime. Although the thin atmosphere of Mars velopment of the Mars Viking missions and saw provides a challenging condition for descent tech- extensive ground based aerodynamic and struc- nologies, advances made in this area will provide tural testing of small-scale articles (< 1.5 m) at benefits to a variety of mission concepts at oth- Mach numbers approaching 5. Larger articles (11 er planets as well, particularly as larger and larg- m) were drop tested at low velocity conditions al- er landed masses are desired. As with AAES, the though no large-scale flight tests ever took place at team has identified five focused technology invest- supersonic conditions. Even though prior devel- ment areas and each is discussed below. opment on inflatable decelerators largely ceased at Although each of the descent technology areas the conclusion of the Viking program, small-scale identified has a unique set of challenges associat- development, primarily in the form of wind tun- ed with it, the issue of scalability impacts nearly nel testing of alternative configurations, has con- all of them. That is, heavier payloads require in- tinued over the past five years within ARMD/ creasingly larger aerodynamic or propulsive decel- FAP. More recently, SIADs have been identified erators during descent. Historical experience with by OCT and SMD as the target of a fast paced

DRAFT TA09-13 development and flight test program aimed at for larger human class missions. In addition, the achieving TRL 6 in 2013. Development of me- ability to deploy such chutes at higher Mach num- chanically deployed or rigid attached decelerators ber, which requires technology advances in textile is largely non-existent except in conceptual stud- strength and thermal performance, may provide ies. Most envisioned attached deployable decelera- significant timeline benefit for some missions. The tors are purely drag devices, but a push investment amount of heritage in subsonic parachutes for oth- in lifting deployables (such as guidable or steer- er planetary applications is much larger than with able systems) could have game-changing impact supersonic chutes, though not always in a relevant in terms of landing precision. Major technical environment (e.g., temperature and density). En- challenges include scalability, deployment meth- tries to Venus and the Giant planets can use tra- odology (for non-inflatable designs), dynamic sta- ditional subsonic parachutes with good efficacy, bility, and controllability. Primary areas of recom- although investment in higher temperature capa- mended NASA investment include: bility textiles is warranted for those applications. • Progression of SIAD development into large Additional investment in evolutionary concepts, scale atmospheric flight testing, like parachute clusters and multi-stage reefing, • High strength, high temperature textile fabrics is warranted, but is not considered a significant and coatings that can extend the thermal advance to the state of the art over, for example, environment in which deployable decelerators Apollo. Investment in lifting trailing decelerators can operate, (such as paragliders) as descent systems does not • Staging mechanisms or systems for phased drag seem warranted given proposed mission require- deployment, ments. NASA investment in parachute technolo- gy has been nearly non-existent since the Viking • Mechanically deployed decelerators and BLDT program with the exception of three high- methods of active (e.g. drag modulation) altitude subsonic drop tests of a 33.5 m ringsail. control, IADs in a trailing configuration (often termed bal- • Steerable and guided deployable decelerators, lutes) have been previously flight tested at Mach and numbers near 10 and could provide improved sta- • Dual mode attached decelerator systems that bility and drag at Mach numbers above 3. Primary are optimized for supersonic and subsonic areas of recommended NASA investment include: flight. • Large (> 30 m) subsonic and supersonic 1.2.2. Trailing Deployable Decelerators parachutes for low density use, including Trailing deployable decelerators are necessary multi-stage reefing, for providing stabilization and deceleration of the • More capable supersonic parachutes (larger entry vehicle through low supersonic and subson- diameter, > Mach 2.2, 800 Pa dynamic pressure ic flight and into terminal descent, and often have deployments), secondary applications for events like stage-sepa- • High strength, high temperature textile fabrics ration. The SOA in subsonic trailing decelerator and coatings that can extend the thermal technology are ribbon and ringsail parachutes, environment in which trailing decelerators can used individually or in clusters, such as employed operate, on Pioneer Venus Large Probe, Galileo, Apollo, • Trailing inflatable aerodynamic decelerators and planned for Orion. The SOA in trailing decel- (ballutes), and erator technology for supersonic use is the DGB • Embedded wireless sensors and algorithms for parachute developed through the PEPP, SPED, smart feedback. SHAPE, and BLDT flight test programs of the 1960’s and 1970’s [9]. These parachutes have been 1.2.3. Supersonic Retropropulsion (SRP) the primary deployable decelerator for planetary As Mars missions approach human class entry robotic missions for the past 40 years. The quali- masses, the required size of supersonic deploy- fication limits in size and deployment conditions able aerodynamic decelerators renders them im- for these parachutes hinder the ability to land mis- practical due to the tenuous atmosphere [6]. As a sions beyond the size of MSL. Supersonic para- result, initiation of propulsive deceleration must chutes larger and more efficient than the 21.5 m occur earlier in the descent phase while the vehi- DGB used by MSL will enable larger Mars robotic cle is traveling at supersonic velocities. Thus, SRP missions and will likely be used as staging devices becomes an enabling technology for human class Mars missions [6]. Smaller-scale SRP systems may TA09-14 DRAFT have enhancing benefits to other mission classes as • Development of advanced event trigger sensors well. The SOA in SRP is limited to primarily Vi- and algorithms to reduce dispersions. king era wind tunnel testing of a limited number 1.2.5. Descent Device Modeling & of notional configurations using perfect gas jets Simulation [7]. No engine development specifically for su- personic decelerator applications has been fund- The development of new and larger deployable ed. NASA investment in the fundamental phys- decelerators and the limited ability to test them at ics of SRP has been reinvigorated recently through full scale and in relevant environments places in- funding by ARMD FAP and ESMD ETDP [8]. creased emphasis on the maturation of modeling Current systems analysis is focused on cryogenic and simulation codes and methods. For a majority LOX/Methane systems (to take advantage of in-si- of the descent technologies identified above, defi- tu propellant generation), but exploration of alter- ciencies in modeling and simulation fall into the nate fuels should be undertaken as well. Primary category of aerodynamic, structural, or combined areas of recommended NASA investment include: fluid-structure interaction. Reliance on empiri- cal models for flexible decelerators, as is current- • Advanced algorithms and sensors to ly done for parachutes, typically produces uncer- dynamically control and stabilize the entry tainty spreads that are prohibitively large for high vehicle in the presence of complex fluid precision landings. Rigid body static and dynamic dynamic interactions, aerodynamics in the supersonic and subsonic re- • SRP configurations and packaging studies for gimes are heavily influenced by the aftbody and high-mass entry vehicles, wake interactions flow field, which in turn dom- • Flowfield modulation through the use of SRP, inate uncertainties in aerodynamic coefficient es- • Deep-throttling (10:1), high thrust (100s of timates. Progressing beyond empirical estimation kN) engines for Mars descent, for flexible decelerators inherently requires ad- • Sustained investment in computational tool vanced FSI modeling capabilities that are still in V&V efforts (particularly fluid mechanics), their infancy. Modeling of SRP flowfield interac- and tion is similarly at a very low level of maturity. The primary technology gap is the application and val- • Expansion of the SRP initiation envelope into idation of current SOA CFD and structures codes hypersonic regimes. to the dynamic simulation of these descent devic- 1.2.4. Guidance, Navigation, & Control es. At the current level of fidelity, the communi- Sensors ty does not yet know what specific advances in GN&C during the descent phase is typically the SOA are required on either the CFD or the non-unique and utilizes common instrumenta- structural analysis sides, however investment in tion with the entry phase. A major exception is in low dissipation flux methods and high spatial and the area of the triggers used to initiate events such temporal accuracy CFD solvers with high-order as parachute deployment and heatshield jettison. turbulence closure is certainly required. NASA Timers or inertial based velocity/altitude triggers support in maturing FSI and SRP modeling ca- represent the SOA in descent-phase event timing. pabilities has been through the ARMD FAP and More advanced event trigger capabilities, based on recent ESMD EDTP technology development new sensor technologies and advanced on board tasks. Primary areas of recommended NASA in- computing power will enable greater landing and vestment include: targeting precision, and more robust EDL sys- • Sustained development of FSI tools for static tems in general. New sensor capabilities for event and dynamic assessment of flexible decelerators, triggers may include real time atmospheric densi- Including acquisition of data sets useful for FSI ty and wind measurements via flush air data sys- validation efforts at relevant aerodynamic and tems, optical terrain relative navigation, “smart aerothermodynamic environments, parachute” deployment, and real time load mea- • Sustained development of SRP modeling tools, surements and heath monitoring. Investments in and alternate event trigger mechanisms have largely been in the area of conceptual studies, although • Sustained investments in improved industries outside of NASA have developed sen- aerodynamic modeling of aftbody and wake sors with potential application to future plane- interaction flows. tary missions. The primary area of recommended NASA investment is: DRAFT TA09-15 1.3. Landing and-go” sample acquisition capability such as at- The landing phase begins with the sensing of tempted by the Japanese Hayabusa probe. These the surface and the operation of a terminal de- technologies are discussed in more detail in Sec- scent propulsion system (if present) and ends with tion 1.3.5. Technology investments are needed in the landing event itself, which is complete when the development of improved landing system dy- the kinetic energy of impact has been dissipated namic analysis and test techniques. Application of and the vehicle is at zero velocity relative to the these techniques enables rapid and thorough ex- surface. In the case of aerial vehicles, landing re- ploration of landing system architectures needed fers to the transition from deceleration to aerial to meet mission specific payload and terrain re- free flight. The landing event may also include an quirements. Systems for safe landing on unconsol- egress/deployment phase to bring the system to idated steep to vertical surfaces and weak surfaces operational state. The landing phase surface sens- including liquids, saturated granular media, and ing may begin before the descent phase ends, re- snow are needed for landing on the various moons sulting in an overlap between the two phases. The of the outer gas giant planets. key areas of technology development are the sys- Large Robotic Landers (100-1500 kg): This tems to sense the surface, descent propulsion mo- category covers robotic exploration of the Moon tors and plume-surface interaction mitigation, and Mars, as well as landing systems for Ve- touchdown systems, high-G survivable systems, nus and Titan. Touchdown system landing per- and small-body guidance. Environmental charac- formance is a major factor in landing site selec- terization of the atmosphere and surface is covered tion. Local topography and mechanical properties in Section 1.4.5. of desired landing sites will dictate the complexi- 1.3.1. Touchdown Systems ty of the touchdown systems employed, converse- Small Robotic Landers (<100 kg): ly, the limitations of the touchdown systems im- This cate- poses restrictions on the sites that can be targeted. gory covers a wide range of potential architectures Large robotic missions will likely entail surface from legged systems to airbags and high G surviv- mobility and /or sample return systems. Technol- able systems (>5000 Gs). Generally, small land- ogy advancements should be focused on: systems ers have and will continue to exploit their inher- designed for sample return class payloads, multi- ent ability to accept high impact loads in order functionality such as landing on a mobility sys- to maximize performance. This is particularly true tem, deployable structures (inflatable or rigid), for impactors such as DS-2 at Mars, and potential and active landing gear for greater performance on ice penetrators to Europa or Enceladus. It is also rocks and slopes. As with small landers, systems expected that continuation of this trend can make for safe landing on unconsolidated steep to verti- these systems more affordable and less sensitive to cal surfaces and weak surfaces including liquids, surface topography. The lack of Agency/industry saturated granular media, and snow are needed for experience in designing and qualifying ruggedized landing on the various moons of the outer gas gi- instruments and electronic and power systems is ant planets. generally their key limitation to broader adoption. Human Class (1500-45000 kg): This category Extensive ruggedization experience exists within covers not only human exploration of the Moon, the defense sector, which in many cases could be NEO’s, and Mars, but also large-scale robotic pre- adapted for space flight. Venus and the icy moons cursor missions. Understanding the technology of Jupiter represent high value scientific destina- needs for human class landing systems will first re- tions with uniquely challenging landing require- quire a system level understanding of the configu- ments. The potential inability to obtain high- ration of the entry system envelope, the payloads, resolution landing site selection prior to arrival and the requirements on the surface mission that requires these systems to be designed for more ex- are not available today. The challenges for these treme surface conditions. Venus landers will also large-scale landing systems will be those of config- require terrain sensing and staging systems capa- uration and mass fraction. Landing performance ble of sustaining elevated temperatures and pres- on large rocks and slopes is not anticipated to be sures. Icy moon landers will require means of ad- the driving challenge due to their size and the pre- dressing icy structures and topographies, either sumption that by the time these missions are real- via soft controlled landing or penetration. Small ized, the ability to actively avoid dangerous surface body landers may require advanced grappling ca- topography will have been achieved. Develop- pabilities for extremely rough surfaces, or “touch- ment of touchdown architectures that are com-

TA09-16 DRAFT patible with launch vehicle and entry system form but are rather captured within TA-02. Regard- factors, provide ample margin on overturning sta- less of the propulsion system employed, rocket bility due to residual horizontal velocity at touch- plumes pose four major hazards during the land- down and minimize the need for complex egress ing event; soil erosion also referred to as trench- and deployment systems is needed. The quantifi- ing, dust cloud generation which can obfuscate cation of the architectural needs of the touchdown sensors and leave deposits on surfaces, generation system will in turn place constraints on the rest of of high velocity debris which can damage near- the flight system. System level design studies are by surface assets, and ground effects interactions needed to co-evolve human class touchdown sys- with the flight system which generate destabiliz- tems in conjunction with the rest of the EDL sys- ing forces on the lander. Mitigation, prediction, tem and to more clearly identify technology gaps protection are the three main ways of addressing which will in turn create more options for design- plume effects. ers. • Mitigation techniques such as rocket motor 1.3.2. Egress and Deployment design (Viking shower-head nozzle) and Robotic Class Egress and Deployments: Egress vehicle architecture (MSL SkyCrane) are the and deployment systems must be tailored to specif- current SOA. Alternate mitigation techniques ic lander and payload needs. As such it is not pos- could include in situ landing site preparation, sible to anticipate specific systems for technology hard landers, or other concepts. Apollo did not development. Instead, there are a few general cat- need to do anything specific to mitigate plume egories of component technology that have broad- effects primarily because the lunar regolith has er application to this area. They are: high power very high inherent strength and did not erode density short life actuation systems including but significantly under plume forces. Mitigation not limited to electromechanical, pneumatic and techniques will be required for Human class pyrotechnic, and rigid and inflatable load bearing landers targeting Mars where the regolith deployable structures that can be used as ramps, mechanical properties are not as robust as those cranes, leveling devices. The MSL SkyCrane sys- on the moon, and plume forces and size are tem is an example in which the spacecraft con- significantly larger than robotic class vehicles. figuration was designed specifically to eliminate • Prediction techniques are largely CFD-based egress. As mobile systems become large, this archi- and in their early phase of development. Plume tectural feature will become more important. Al- effects modeling is needed to ensure that ternative landing system architectures that avoid vehicle configuration, landing site topography the need for egress are needed as well. and flight system control system and control Human Class Egress and Deployments: As authority are compatible with the predicted landers grow in size the criticality of addressing levels of interaction. Continued investment egress and deployment of primary payload be- in CFD-based plume interaction (including comes more pronounced. System design stud- multi-phase continuum flow modeling) is ies are needed initially to identify the architectur- needed. Methods of predicting soil erosion/ al needs of the surface payloads, the touchdown trenching are also needed. systems and subsequently the egress and deploy- • Plume accelerated debris protection techniques ment systems. These studies must be done as part are required for mission architectures which of an overall system study since the influence of have surface elements landing within less than the egress and deployment systems will have first ~1 km from each other. order influence on the touchdown systems and 1.3.4. Landing Guidance, Navigation, and entry system configurations. It is likely that any Control for Large Bodies given landed system must be self sufficient in en- suring that it can place the desired payload on the Previous landers with terminal descent propul- planetary surface. sion have utilized various levels of knowledge of position and velocity relative to the surface. Ra- 1.3.3. Landing Propulsion dars have been used to determine altitude and ver- Technology advances in rocket propulsion tech- tical velocity (Mars Pathfinder had only that). In nology (such as nano-energetic fuels and high- addition, horizontal velocity has been determined thrust LOX/Methane engines) could provide sig- using Doppler radar (Surveyor, Apollo, Viking, nificant benefits to EDL. However, such advances Phoenix), as well as passive optical imaging with are not directly considered within this roadmap, onboard correlation (Mars Exploration Rover). To

DRAFT TA09-17 enable landing at more challenging and hazard- capabilities needed for small-body in situ explo- ous locations, to enable surface rendezvous, and ration, sample acquisition, or crewed operations. to land with greater safety for crewed missions, fu- Japan’s Hayabusa spacecraft attempted proximity ture landers will require knowledge of location rel- operations at Itokawa, but was damaged by un- ative to the target, the detection of landing haz- controlled contact with the surface [10]. Proximi- ards (both provided by crew on Apollo), and the ty maneuvering around and on small bodies such efficient use of propellant in the redirection of as near-Earth asteroids and comets is a terminal the terminal descent to the desired targets and in descent guidance and landing problem. It differs avoiding hazards. The challenges in meeting these from large-body terminal descent in that several needs are: new modes are added such as ascent, hover, and • Robust algorithms for terrain tracking using touch, all in concert with fault protection respons- passive and/or active imaging and on-board es, and it differs in that the time scales for knowl- maps derived from orbital data, compensating edge and control are longer. Future robotic and for variations in distance, orientation, lighting, crewed missions to small bodies will require ter- and visibility (also enabling for small bodies rain-relative guided flight around and in contact — see Section 1.3.5, and potentially enabling with the small body, precision targeting to avoid for lower cost altitude and velocity sensors to hazards and apply tools between them, guidance replace existing radars when terrain tracking is laws and fault protection integrated together to not needed), assure only controlled contact with the body, and • Robust algorithms for the identification and the ability to remain contact with the body either location determination of landing hazards actively or passively and apply normal force to using just-acquired passive and/or active the body with surface sampling and measurement imaging, with a low probability of false positive tools. The challenges in meeting these needs are: or negative indications, • Terrain tracking as in Section 1.3.4, • Passive and/or active visible or near-IR imaging • Guidance that targets tool contact with the sensors to provide the data for terrain tracking body using terrain tracking, and that robustly and hazard detection, and that are compact reacts to unforeseeable forces from contact, and low-power, controlling both attitude and position to avoid • Algorithms that predict and correct landing inadvertent contact by other portions of the location errors relative to a changing target from spacecraft, the position and hazard detection components, • State machines to control descent, touch, land, and do so with optimal fuel usage, hover, and ascent modes in the presence of • Adaptive or reconfigurable control algorithms faults, to react to failed systems or components, and • Touchdown and anchoring/grappling systems • Highly capable and low power on-board to enable the application of order 100N normal dedicated compute elements to support terrain forces to surface material by end-effector tools tracking, hazard detection, and possibly for periods of time from seconds to hours trajectory optimization at rates suitable for use depending on the application, and in terminal descent on large bodies. • A physics-based end-to-end simulation A worthy goal would be development of an inte- capability with hardware in the loop for the grated Terrain Tracker assembly, that would incor- qualification of terrain tracking and proximity porate the sensors, compute elements, algorithms, operations during project development. and maps all into a single, small box with a simple 1.4. Vehicle Systems Technology interface, much like modern Star Trackers. Haz- A comprehensive understanding of component, ard detection algorithms could be hosted on the subsystem, and system level performance is inher- device as well. ent to all successful entry vehicle systems. Prior 1.3.5. Small Body Systems to having sufficient flight test and operational ex- The SOA for NASA in small-body proximity perience, which validates a design and identifies maneuvering was achieved by NEAR, which ap- anomalies, design verification is provided via large proached and landed on Eros in an end-of-mis- scale ground tests, flight tests and evaluation of sion experiment. This demonstrated the ability to design margins compared to the environmental navigate a descent, but did not demonstrate the factors that stress a system’s capability. In addition, systems technology capabilities perform a key role TA09-18 DRAFT of identifying, characterizing and maturing system gether. In addition technologies oriented around level integration and design. EDL systems are by payload mass and configuration staging, EDL for their nature an integrated framework of technolo- atmospheric bodies will require atmosphere char- gies that necessitate system level efforts for robust acterization to a higher degree than currently maturation. At the entry vehicle level, systems are achieved. Given the current state of atmospheric defined by their payload mass class as well as the models for potential planetary entries to Mars, Ve- means by which the vehicle transitions from one nus or some of the moons within our solar system, stage or configuration to another. In addition, ar- additional technologies are needed to characterize chitectural studies provide a strategic and criti- the altitude and seasonal variations. Application cal function to define a path forward that aligns of these technologies would provide a framework with national goals and objectives and enable in- of scientific data to support development of vari- formed technology investment decisions. Within ational models for day-of-entry assessments and this context, there are several core technology fun- weather pattern prediction and modeling. damentals that must be addressed to enable this 1.4.1. Separation Systems decision making process. This roadmap atempts EDL traverses flight segments usually involving to identify those technologies that, will have the essentially static vehicle configurations, punctu- most significance in shaping such capabilities. As ated by transitions between these configurations. a distinction from individual component or sub- The time scales of vehicle reconfiguration vary sig- system level technologies, certain EDL technolo- nificantly between exo-atmospheric conditions, gy frameworks encompass many aspects of a vehi- where there are no aerodynamic forces to contend cle and will be captured in the context of a system with, down and through aerodynamic loading as- technology. Vehicle Systems Technology will thus sociated with hypersonic, supersonic, and subson- be segmented into three areas that have implica- ic conditions for atmospheric entries. Many of tions across the entire EDL architecture: Separa- the component level technologies associated with tion Systems, Vehicle Technologies, and Atmo- Separation Systems are mentioned in previous sec- spheric Modeling & Surface Characterization. tions. However, it is essential that particular at- Major challenges exist to mature advanced tech- tention be given to the technology involved with nologies for micro payloads (<100 kg), science ro- integrating component level technologies for stag- botic payloads (< ~1500 kg), large payloads (<~ ing and separation into flight demonstrated sys- 3000 kg) and human class payloads (>~3000 kg) tems. Within the context of this roadmap, this is which would eventually lead to human class capa- considered to be of sufficient criticality that specif- bilities. Identification of system technologies that ic characterization of the SOA and push technol- enable a wide range of payload mass is essential. ogies will be included. A brief characterization of The likely need to accomplish technology devel- each EDL flight phase, in addition to several other opment with reduced ground infrastructure will key areas, is included below. Due to the difficulty also drive modeling and simulation to provide in- in successfully and reliably accomplishing vehicle tegrated physical modeling environments. These staging and reconfiguration, sustained investment integrated modeling capabilities must be ground- in establishing integrated technology solutions for ed in physics, and coupled in a manner that ad- each of these flight phases will be necessary and dresses the gap between what can be tested on the will ultimately define what combinations of entry ground and the actual flight environment. As a systems are possible. Component and subsystem separate consideration from payload mass class, technologies must be integrated and flight dem- the technology to successfully stage, transition, onstrated due to the high risk and single point and avoid recontact of vehicle components dur- failure modes that are critical to successful vehicle ing exo-atmospheric, hypersonic, supersonic, ter- staging and separation. Flight testing wrings out minal descent or touchdown phases will always the inadequacies in such integrated systems, and encompass some of the most challenging vehicle is a necessary component of technology matura- technology issues covered by EDL. As a necessary tion. A vehicle transition example for the system subset of vehicle reconfiguration, the technologies level framework can be appreciated by consider- that enable a transition from one configuration or ing possible re-contact after an initially successful stage to another are critical. Vehicle transition and separation. All the component technologies lead- staging technologies ultimately determine which ing up to separation can be successful, but the in- technologies can be successfully integrated to- tegrated approach must reliably and successfully

DRAFT TA09-19 lead to complete separation of the vehicle compo- discussion above that are crosscutting and nents without recontact. enabling for vehicle transition and staging. • Exoatmospheric. The current SOA includes Other component technologies that may shroud separation, vehicle rendezvous and have significant roles to play in enabling new docking, and in-space construction (e.g., ISS). vehicle transition contexts include advanced Push technologies include on-orbit component pyrotechnics, springs, airbags, and drag robotic construction, mechanical or inflatable augmentation devices (see Sections 1.2 & 1.3). deployment of staged systems, or rigidizable 1.4.2. Vehicle Technology aeroshell sub-systems. EDL systems naturally separate themselves into • Hypersonic-Supersonic. The current SOA categories driven by external factors. At the en- includes aerodynamic control surfaces try vehicle level, payload mass is one of the pri- and trailing vehicle separation/disposal by mary characteristics that drives the option space propulsive, pyrotechnic or mechanically for technology utilization. Systems that are viable assisted components. Push technologies for one class of mass are not appropriate for other include mechanical or inflatable deployment classes. With this factor in mind, EDL vehicles are of staged hypersonic aeroshell separation, or separated here into four payload mass categories propulsive based hypersonic stage separation. that require technology development and tailor- • Supersonic-Subsonic. The current SOA ing to suit the needs of each vehicle scale. Focus- includes aerodynamic control surfaces, payload ing effort into categories will enable the technolo- bay door opening/closing, Space Shuttle gy solutions that are appropriate for different mass Solid Rocket Booster staging and separation, ranges to be distinguished as such, and pursued subsonic heat shield separation for Mars robotic separately. In addition to this, one of the principle vehicles, parachute (drogue) deployment, techniques to facilitate investigations of integrat- mortar deployment or mass ejection assisted ed design solutions and flight test performance is by propulsive, pyrotechnic or mechanically the use of integrated mission simulations. For this assisted components. Push technologies reason there is a strong dependency on the tech- include mechanical or inflatable deployment of nologies identified earlier, in addition to probabi- staged systems, supersonic aeroshell and entry listic design technologies identified in the TA12 shroud separation, or propulsive-based stage (Materials and Structures) and integrated mod- separation. eling approaches such as the digital vehicle con- • Terminal Descent – Touchdown. The current cept identified in TA11 (Modeling and Simula- SOA includes parachute, mortar, or landing tion). Ultimately, the basis of these simulation gear deployment, or shroud separation assisted capabilities in concert with ground and flight test- by propulsive, pyrotechnic or mechanically ing will provide a foundation for down select de- assisted components. Push technologies include cisions on technology suites and certification anal- larger scale tethering devices for SkyCrane yses for the robotic, robotic-precursor and human type systems (which separate the landing and spaceflight missions of the future. Such simula- propulsion systems during terminal descent). tion capabilities provide the only means to evalu- • Landing Site Surface Preparation. The ate how sub-scale flight tests performed to validate potential severity and complexity associated system technology approaches will be applicable with having a design robust to the interaction to full scale. It is thus imperative that higher fidel- of soil/debris with large propulsive system ity simulation and testing capabilities be integrat- plumes suggests that technology investments ed into the fabric of early technology development should be considered for preparing a landing efforts, and that conscious and strategic empha- zone prior to touchdown. Risks associated with sis is placed on developing these technologies to plume-soil interaction are discussed in landing support and enable Vehicle Technology. To enable Section 1.3.3. The current SOA is at a very this process, additional perspective on the vehicle low TRL and includes push technologies such payload mass classes and integrated simulation are as microwave conversion of soil to sintered provided below. This perspective is defined in or- surfaces or other soil processing methods to der to capture a framework that will enable com- prepare a ‘landing pad’. ponent and systems level technologies to be pur- • Component Technologies. Several component sued within an approach that directly recognizes level technologies have been mentioned in the that technologies will be best suited to a limited payload mass range. TA09-20 DRAFT • Micro Payloads. This class of payload mass suited to a range of potential large mass entry is targeted to less than approximately 100 systems. Current SOA is undefined – no high kg. The current SOA is characterized by the fidelity vehicle designs exist for greater than unsuccessful Deep Space-2 and Beagle 2 Mars 1000 kg. Push technologies include the use entry vehicles. The high visibility and success of inflatable/deployable aeroshells, higher of the NASA micro-satellite activities presents L/D configurations, precision or pinpoint a compelling case for this model to be extended landing capabilities, high-G capabilities to micro-entry vehicles. The low entry price for for landing that enable different design this weight class of vehicle makes it an ideal approaches, aerocapture, aerobraking, more candidate to promote STEM efforts that will capable supersonic parachutes, SRP, utilization enable and engage a skilled workforce for of terrain relative navigation or navigation the future. Push technologies could include aids/beacons, novel combinations of these high-g self-guided navigation aids/beacons or technologies, etc. sub-surface investigators, collaborative nano- • Human Scale Payloads. This class of payload satellite/micro-payload systems for return of mass is targeted to approximately 30 to 60 t. A biological or material samples from orbit, or significant impediment to the development of low-power/mass MEMS based accelerometers technologies for this scale is the cost of launch or IMUs. vehicles. Technology developments achieved in • Science Robotic Payloads. This class of payload payload mass ranges for previous, lower mass, mass is targeted to less than approximately categories may or may not be applicable to 1500 kg. Current SOA is characterized by such large entry masses for atmospheric entry. the Mars robotic vehicles, which utilize the Technologies not required for atmospheric Viking aeroshell configuration, reaction entry are likely to be much more relevant. It control systems for high-speed control, is assumed that there would be at least one DGB supersonic parachutes, and a variety human precursor mission without a crew and of terminal descent systems. These terminal at a reduced scale before an actual crewed descent systems include the use of airbags, sub- mission. This mission would serve as a final sonic retropropulsion, or the SkyCrane. Push robotic precursor, potentially preposition technologies include the use of non-Viking assets necessary to support a crewed mission, entry configurations (The NASA 2018 Mars and demonstrate all the EDL system level “Max-C” mission is planning an Apollo-derived technologies and their integration prior to a entry configuration), inflatable/deployable crewed mission. aeroshells, higher L/D configurations, precision • Integrated Mission Simulation and Risk/ or pinpoint landing capabilities, high-G Reliability Characterization. The importance capabilities for landing that enable different of integrated mission simulations of the design approaches, aerocapture, aerobraking, entry profile and design performance is often SRP, etc. underestimated. Modeling capabilities • Exploration Precursor Payloads. This class that are inter-disciplinary, incorporate the of payload mass is targeted to less than ~3000 appropriate physical models, and are robust kg. It is expected that initial robotic precursor are necessary to address ground to flight missions will be limited to the current upper scaling, flight test anomaly resolution, risk limit of landed mass afforded by the Delta IV-H mitigation, appropriately evaluate and manage (2-3 t). This class is intended to encompass risk for flight demonstrations and establish more than the current upper limit for payload certification for the human missions. The cost mass to Mars. Smaller mass requirements for of performing many, or even several, flight tests intermediate payload mass technologies will to validate designs is assumed to be prohibitive increase the magnitude of technology jump to and is a motivation for high fidelity modeling reach human scale missions. A larger robotic capabilities which are validated by ground and precursor upper end mass requirement, coupled sub-scale, sub-system level flight testing. High with increasingly larger entry vehicles would fidelity integrated simulation capabilities will allow confidence building before performing be necessary As pointed out earlier, the ground a human scale precursor mission. Exploration testing and simulation capabilities that are Precursors would thus provide a natural developed during initial TRL advancement, in framework for development of technologies addition to DDT&E, will form the foundation DRAFT TA09-21 for TRL advancement to 6 and beyond. In mechanisms to encourage revolutionary and order to declare that a given ground or flight out of the box technology thinking and game test is relevant to flight requires modeling that changing architectural level impacts on mission demonstrates applicability of the TRL 6 test systems must be encouraged and supported. to the ultimate flight scale or environment. 1.4.3. Atmospheric Modeling and Surface Current SOA is much higher for individual Characterization technical disciplines than for such integrated Atmospheric modeling is important to all aero- analyses. To facilitate such analyses and risk dynamic phases of flight including aerocapture, characterization, the application of well aerobraking, entry and descent. Precise landings accepted frameworks involving statistical require guided vehicles to navigate through varia- methods, Monte Carlo assessment, uncertainty tions in atmospheric density and winds. Addition- quantification, Probabalistic Design, etc. are ally, uncertainty in atmosphere density is typically necessary. These technologies have not been handled at the expense of performance via conser- implemented for hypersonic atmospheric vatively sized decelerator systems. Improved un- entry simulations and will potentially require derstanding of the general magnitude of atmo- new techniques to integrate EDL simulation spheric variations would directly improve landed capabilities into effective multi-physics mass performance and facilitate guidance algo- capabilities. The process of TRL advancement rithm development, selection and tuning while provides the progression of activities that real-time measurement would directly enable in- are necessary to motivate and execute the creased landing precision. Characterization of the development of calibrated and validated dust environment at Mars and the potential im- high fidelity integrated analyses. A plan for pact for entry systems will be critical for human technology development of integrated EDL Mars missions. Terrain tracking will require on- systems should include a cyclical application board maps of the surface for use that are generat- of increasingly higher fidelity and validated ed from orbital imagery and altimetry. Automated integrated analyses to guide the selection of systems to convert orbital data to onboard maps technologies, technology combinations and will enable small body missions with limited time architectures appropriate for each landed mass between orbital data collection and proximity op- class. These technologies must be developed for erations. The SOA in atmospheric modeling varies EDL if they are to be in a position to support with the planetary body. At Mars, pressure cycles this need. and atmospheric density modeling is anchored by • Game Changing Technologies. Clearly, a paucity of surface pressure measurements from there are new technologies, many of which the Viking landers and subsequent robotic mis- are unknown today, which could change sions while other planets have even less data with the entire way we approach the EDL phases which to anchor models. Post-landing data collec- of any particular mission. For example, if tion of pressure and low altitude winds will pro- nuclear thermal propulsion (NTP) or other vide ground truth for mesoscale wind models used high energy density propulsion systems to validate precision landing. Methods of measur- are realized then many of the significant ing density and wind are common in Earth ap- EDL challenges associated with hyperbolic plications though these have not been applied to entry velocities could be greatly relieved other planets. There may well be opportunities to and solved propulsively with significantly leverage current and ongoing Earth science invest- reduced mass and improved operational ments for atmospheric measurements and charac- flexibilities. Other non-aerodynamic drag terization to help enable the same for other plan- options for decelerating systems might include etary bodies with atmospheres (e.g. advanced gravitational or electromagnetic repulsion orbital platform LIDAR instrument development systems, revolutionary energetics for retro for Earth atmosphere CO measurements could energy, or ingestion and reuse of mass during 2 be extended to Mars orbiters for CO2 atmospher- atmospheric traversal to greatly enhance ic density and wind profiling). NASA’s invest- propulsive efficiencies. In the past, very little ment in planetary atmospheric modeling specific emphasis has been placed on the more far out to EDL capabilities is primarily in the continued technology opportunities and the mission development of Global Reference Atmosphere architecture impacts, beyond just vehicle Models (recently only for Earth – Orion, and mis- components or subsystems. Funding and sion-specific Mars – MSL, Titan - Huygens) and TA09-22 DRAFT in remote measurements made by orbiting space- flight test) that is ultimately required to develop craft. Primary areas of recommended future in- new EDL technologies and architectures for hu- vestments include: man and robotic exploration. • Distributed short and long duration surface Entry - No single ground-based facility exact- weather measurements at Mars, ly replicates high-energy flight conditions associ- • Development of orbiter instruments for wind ated with hypervelocity entry into atmospheres and atmospheric property characterization at surrounding planetary bodies. Instead, individu- multiple altitudes including those relevant for al facilities have been developed that replicate a aerocapture and aerobraking, particular aspect of high speed flight, and when • Development of automated data fusion for combined with analysis and flight test capabilities distinct orbital data types (visible stereo (e.g., suborbital sounding rockets, high altitude imagery, multi-spectral, and altimetry) and its rocket assisted balloon launch, Earth based reen- conversion to onboard topography and albedo try flight tests), these ground-based facilities serve surface maps suitable for use by a terrain to anchor component EDL technology develop- tracker, ment and flight system qualification. Ground-based wind tunnels, with operating • Development of a low-impact standard conditions from subsonic through hypersonic ve- atmospheric data package for all Mars landed locities, achieve fluid dynamic similarity to flight, missions to benefit future Mars landers with generally at subscale conditions. These facilities surface pressure and upward looking wind are required to obtain vehicle aerodynamics across measurements, a large range of relevant Mach numbers, com- • “Scout” Probes designed for measuring ponent aerodynamic loads, control effector and atmospheric properties and entered ahead of a aero/jet interactions, acreage heating patterns on primary entry vehicle, and the vehicle surface, and estimates of configuration • Vehicle based sensors and instrumentation specific turbulent transition and heating effects for making real-time assessments of local for the specific vehicle shapes. and far-field atmospheric properties. Arc-jets are required to understand thermal pro- tection system response at moderate and high 2. Infrastructure heat flux during hypersonic entry. These facilities EDL technology development, qualification, achieve flight-like heating environments, i.e., heat and certification in flight systems will require ac- rate, temperature, heat load, and shear to TPS test cess to numerous ground test facilities and labo- specimen samples. In this manner, the thermal re- ratories around the country. Many of these facili- sponse of flight hardware can be determined. Up- ties require unique and highly specialized resident graded capability over that currently available will engineering talent that has been honed over many likely be required to achieve conditions needed decades of practice. NASA and the U.S. space in- to simulate full radiative and convective heating dustries that are developing new EDL capabilities conditions associated with human scale Earth en- should also strive to take advantage of the capa- try vehicles from beyond LEO (e.g., direct entries bilities of international partners, ITAR issues not- from NEO or Mars return trajectories. The Giant withstanding. Maintenance and stewardship of fa- Planet Facility, a leg on the Ames Research Cen- cilities through sporadic periods of utilization is ter (ARC) arc-jet complex, was used to test ther- essential to NASA’s ability to conduct EDL mis- mal protection material in a radiative/ convective sions. Advancements in computational modeling H/He environment in the 1970’s. This portion of and simulation will most certainly be relied upon the Ames test complex is no longer operational, to a much greater extent than in the past for en- but a similar testing capability is likely required gineering design and ground to flight traceability, for TPS development of future probe missions to but advancements in analytical tools and physics the gas giants. based models still require verification and vali- Shock tubes are used to understand the high dation data, which are typically obtained either temperature atomic, chemical kinetic, and gas dy- through ground based component testing or sub- namic behavior of the atmospheric gases at high scale flight testing. Flight testing, in relevant en- temperature, which is essential for nonequilibri- vironments at sub- or full-scale, is often the final um chemistry and shock layer radiation model- step in full system qualification. It is the combi- ing. This information is used to develop detailed nation of the three (ground test, simulation, and physical models required for aerothermodynamic DRAFT TA09-23 flight prediction. 3. Immediate Actions Indoor and free flight ballistic range facilities are Given the current state of NASA investments, often utilized to determine dynamic aerodynam- the authors recommend the following immediate ic force coefficients, which are of significance for actions that are not currently receiving adequate aerostability assessment, particularly in the tran- attention: sonic and supersonic regimes. These facilities can • Conduct advanced architectural level studies also be used to obtain stagnation-point heating for human and large robotic Mars surface and noise-free transition data. missions that are developed to a Phase-A Descent - Test requirements frequently demand level of design detail to drive out critical EDL access to low-density environments at high speed, technology needs and schedules. often near or at full scale. Such access is typically made possible through high-altitude Earth based • Develop a NASA policy for required EDL flight testing. NASA’s Balloon Program supports instrumentation and data acquisition in order high altitude Earth payloads up to 8,000 pounds. to advance and build confidence in models Payloads released from the high altitude balloons that are essential to EDL system qualification. can then be accelerated to relevant test conditions • Implement a plan for instrumenting the Mars either through simple gravity assist (subsonic) or 2018 mission with a system at least as capable via rocket propelled kick stages (supersonic), sim- as the MEDLI system on MSL. ilar to the approach utilized in the Viking super- • Requalify heritage Carbon Phenolic TPS to sonic parachute qualification effort. Full-scale support MSR earth entry reliability and Venus tests that exceed the capacity of existing balloons entries. Requires immediate carbonization will require either the development of new bal- of the NASA stockpile of heritage Rayon loon systems or specialized ground launched rock- before domestic carbonization capability is et systems. Continued access to large-scale wind mothballed. tunnel facilities is also critical to testing and qual- • Investigate deployable entry systems that ifying aerodynamic decelerators. would be an alternative to inflatable approaches Landing – Terminal descent, including maneu- (currently being investigated) in order to not vering, hazard detection and avoidance, integrated be dependent on a single path. in-situ sensor performance and GN&C algorithm • Investigate alternative high mass deceleration testing, and performance assessments of high and approaches to SRP (currently being low gravity body touchdown systems, will require investigated) in order to not be dependent on capabilities to test at the component and integrat- a single path. ed system level. These capabilities may range from • Flight-qualify large subsonic or supersonic simple drop tests (crane or large gantry facilities parachutes and more capable higher-Mach with suspension systems to simulate low-G grav- supersonic drag devices to support MSR itational environments) to helicopter, aircraft or surface payloads. low-altitude rocket systems for sensor testing. It is recommended that NASA form a test fa- • Develop a supersonic Earth atmosphere test cilities team to develop a uniform cost basis and capability to support decelerator and retro- long-term utilization requirements for these facili- propulsion developments for high mass ties. Because of the critical nature of the test facil- systems. ities and the resident expertise, this cost and utili- • Develop terrain tracking and hazard detection zation information is vital for planning EDL and technology to support Mars 2018 and MSR- other technology and system-level capability de- class missions as well as robotic and human velopment. asteroid missions. In addition to facilities, the “human infrastruc- • Advance the state of NASA pyrotechnic ture” component is also a critical part of EDL initiator technology to take advantage of technology. Because of the specialized NASA- significant advances in initiator performance unique nature of many of the necessary technical and reliability in the defense sector. skills within the discipline, it is critical that NASA • Form a NASA EDL Test Facilities Team to accelerate its development of STEM and early ca- periodically assess and prioritize the needs reer opportunities for the next generation of ex- and future utilization of existing and new perts in key EDL technologies. test capabilities for EDL systems, in order to provide a reviewed and accepted source of

TA09-24 DRAFT information for NASA and other government Administration Agency facilities investments planning. NEAR Near-Earth Asteroid Rendezvous NESC NASA Engineering and Safety Center NFAC National Full-Scale Aerodynamic Complex Acronyms OCE Office of the Chief Engineer AAES Aerobraking, Aerocapture, and Entry Sys- OCT Office of the Chief Technologist tems p Pressure (stagnation) AEDC Arnold Engineering Development Center PEPP Planetary Entry Parachute Program AFRSI Advanced Flexible Reusable Surface PICA Phenolic Impregnated Carbon Ablator Insulation q Heat Rate ARC Ames Research Center RCC Reinforced Carbon-Carbon ARMD Aeronautics Research Mission Directorate RCS Reaction Control System BLDT Balloon Launch Decelerator Test RV Reentry Vehicle CFD Computational Fluid Dynamics SHAPE Supersonic High Altitude Parachute C.G. Center of gravity Experiment COTS Commercial Orbital Transportation SIAD Supersonic Inflatable Aerodynamic Services Decelerator CUIP Constellation University Institutes Program SLA Super Lightweight Ablator CY Calendar Year SMD Science Mission Directorate DDT&E Design, Development, Testing, and Evaluation SOA State-of-the-Art DGB Disk Gap Band SPED Supersonic Planetary Experiment DNS Direct Numerical Simulation Development DRA Design Reference Architecture SRP Supersonic Retro Propulsion EDL Entry, Descent, and Landing STEM Science, Technology, Engineering, and ESMD Exploration Systems Mission Directorate Mathematics ETDD Exploration Technology Development t Metric ton (1000 kg) Demonstrations TABS Technology Area Breakdown Structure FSI Fluid-Structure Interaction TDP Technology Development Project G One Earth surface gravitational TPS Thermal Protection System acceleration TRL Technology Readiness Level GN&C Guidance, Navigation & Control U.S. United States HEO High Earth Orbit V&V Verification and Validation HIAD Hypersonic Inflatable Aerodynamic VSTAA Vehicle Systems Technology and Decelerator Architecture Analyses IAD Inflatable Aerodynamic Decelerator WLE Wing Leading Edge IR Infrared IRVE Inflatable Reentry Vehicle Experiment ISHM Integrated System Health Monitoring Acknowledgements ISPT In-Space Propulsion Technology The NASA technology area draft roadmaps were ISS International Space Station developed with the support and guidance from ITAR International Traffic In Arms Regulations the Office of the Chief Technologist. In addition kg Kilogram to the primary authors, major contributors for the kN Kilo Newton TA09 roadmap included the OCT TA09 Road- L/D Lift/Drag mapping POC, James Reuther; the reviewers pro- LCC Life Cycle Cost vided by the NASA Center Chief Technologists LEO Low Earth Orbit and NASA Mission Directorate representatives, m Meter and the following individuals who served as a Red MAV Mars Ascent Vehicle Team: Brett Drake, John Connolly, Ethiraj Ven- MAX-C Mars Astrobiological Explorer - Cacher katapathy, Dean Kontinos, Michele Munk, and MEMS Micro Electro Mechanical Systems Neil Cheatwood. MMOD Micro Meteoroid Orbital Debris MSL Mars Science Laboratory MSR Mars Sample Return NASA National Aeronautics and Space DRAFT TA09-25 References 1. B.G. Drake (ed.), Human Exploration of Mars Design Reference Architecture 5.0, NASA-SP-2009-566, July 2009. 2. NASA Agency Mission Planning Manifest, Microsoft Excel Worksheet, Working Draft Revision 4-15-10 3. Human Exploration Framework Team Overview for A-STAR Technology Roadmapping Teams, Version 9, Microsoft PowerPoint Presentation, August 2010 4. "Science Plan for NASA's Science Mission Directorate 2007 - 2016," National Aeronautics and Space Administration, accessed via web: http://science.nasa.gov/ media/medialibrary/2010/03/31/Science_ Plan_07.pdf, Oct. 2010. 5. Cassell, A.M., “Development of Thermal Protection Materials for Future Mars EDL Systems,” AIAA Paper No. 2010-5049, Jun. 2010. 6. Cianciolo, A.M., “Entry Descent and Landing System Analysis Study: Phase 1 Report,” NASA TM-2010-21620, July 2010. 7. Korzun, A.M., “Survey of Supersonic Retropropulsion Technology for Mars Entry, Descent and Landing,” Journal of Spacecraft and Rockets, Vol. 46, No. 5, pp 929-937, 2009. 8. Edquist, K.T., “Development of SRP for Future Mars EDL Systems,” AIAA Paper No. 2010-5046, Jun. 2010. 9. Moog, R. D., and Michel, F. C., "Balloon Launched Viking Decelerator Test Program Summary Report," NASA Contractor Report CR-112288, Mar. 1973. 10. Yano, H. et al., “Touchdown of the Hayabusa Spacecraft at the Muses Sea on Itokawa,” Science, Vol. 312, Jun. 2006, pp. 1350-1353.

TA09-26 DRAFT This page is intentionally left blank

DRAFT TA09-27 November 2010

National Aeronautics and Space Administration

NASA Headquarters Washington, DC 20546 www.nasa.gov TA09-28 DRAFT