Co-Optimization of Mid Lift to Drag Vehicle Concepts for Mars Atmospheric Entry
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MARS SCIENCE LABORATORY ORBIT DETERMINATION Gerhard L. Kruizinga(1)
MARS SCIENCE LABORATORY ORBIT DETERMINATION Gerhard L. Kruizinga(1), Eric D. Gustafson(2), Paul F. Thompson(3), David C. Jefferson(4), Tomas J. Martin-Mur(5), Neil A. Mottinger(6), Frederic J. Pelletier(7), and Mark S. Ryne(8) (1-8)Jet Propulsion Laboratory, California Institute of Technology, 4800 Oak Grove Drive, Pasadena, California 91109, +1-818-354-7060, fGerhard.L.Kruizinga, edg, Paul.F.Thompson, David.C.Jefferson, tmur, Neil.A.Mottinger, Fred.Pelletier, [email protected] Abstract: This paper describes the orbit determination process, results and filter strategies used by the Mars Science Laboratory Navigation Team during cruise from Earth to Mars. The new atmospheric entry guidance system resulted in an orbit determination paradigm shift during final approach when compared to previous Mars lander missions. The evolving orbit determination filter strategies during cruise are presented. Furthermore, results of calibration activities of dynamical models are presented. The atmospheric entry interface trajectory knowledge was significantly better than the original requirements, which enabled the very precise landing in Gale Crater. Keywords: Mars Science Laboratory, Curiosity, Mars, Navigation, Orbit Determination 1. Introduction On August 6, 2012, the Mars Science Laboratory (MSL) and the Curiosity rover successful performed a precision landing in Gale Crater on Mars. A crucial part of the success was orbit determination (OD) during the cruise from Earth to Mars. The OD process involves determining the spacecraft state, predicting the future trajectory and characterizing the uncertainty associated with the predicted trajectory. This paper describes the orbit determination process, results, and unique challenges during the final approach phase. -
IMU/MCAV Integrated Navigation for the Powered Descent Phase Of
Overview of Chinese Current Efforts for Mars Probe Design Xiuqiang Jiang1,2, Ting Tao1,2 and Shuang Li1,2 1. College of Astronautics, Nanjing University of Aeronautics and Astronautics, Nanjing 210016, China 2. Space New Technology Laboratory, Nanjing University of Aeronautics and Astronautics, Nanjing 210016, China Email: [email protected], [email protected] Abstract Mars exploration activities have gathered scientific data and deepened the current understanding about the Martian evolution process and living environment. As a terrestrial planet, Mars is also an excellent proving ground of some innovative special-purpose aerospace technologies. In the U.S., Soviet Union (Russia), Europe, and India, missions sending spacecraft to explore Mars currently exist, and the first three have landed their spacecrafts on the surface of Mars. China has programmed five Mars landing missions in next 20 years, and some critical technologies for Mars landing mission have been studied beforehand. In this paper, we will report Chinese Mars exploration mission scenario and the latest progress in dynamics study and guidance navigation and control (GNC) system design for Chinese first Mars probe. Some elementary design rules and index will be described according to simulation experiments, analysis and survey. Several new coping strategies will be proposed to be competent for the challenges of the Mars entry descent and landing (EDL) dynamics process. Details of the work, the supporting data and preliminary conclusions of the investigation will be presented. Based on our current efforts, one candidate system configuration architecture of Chinese first Mars probe will be shown and elaborated in the end. United States launched the “Mars Pathfinder (MPF)” 1. -
Deep Space 2: the Mars Microprobe Project and Beyond
First International Conference on Mars Polar Science 3039.pdf DEEP SPACE 2: THE MARS MICROPROBE PROJECT AND BEYOND. S. E. Smrekar and S. A. Gavit, Mail Stop 183-501, Jet Propulsion Laboratory, California Institute of Technology, 4800 Oak Grove Drive, Pasa- dena CA 91109, USA ([email protected]). Mission Overview: The Mars Microprobe Proj- System Design, Technologies, and Instruments: ect, or Deep Space 2 (DS2), is the second of the New Telecommunications. The DS2 telecom system, Millennium Program planetary missions and is de- which is mounted on the aftbody electronics plate, signed to enable future space science network mis- relays data back to earth via the Mars Global Surveyor sions through flight validation of new technologies. A spacecraft which passes overhead approximately once secondary goal is the collection of meaningful science every 2 hours. The receiver and transmitter operate in data. Two micropenetrators will be deployed to carry the Ultraviolet Frequency Range (UHF) and data is out surface and subsurface science. returned at a rate of 7 Kbits/second. The penetrators are being carried as a piggyback Ultra-low-temperature lithium primary battery. payload on the Mars Polar Lander cruise ring and will One challenging aspect of the microprobe design is be launched in January of 1999. The Microprobe has the thermal environment. The batteries are likely to no active control, attitude determination, or propulsive stay no warmer than -78° C. A lithium-thionyl pri- systems. It is a single stage from separation until mary battery was developed to survive the extreme landing and will passively orient itself due to its aero- temperature, with a 6 to 14 V range and a 3-year shelf dynamic design (Fig. -
Aerothermodynamic Design of the Mars Science Laboratory Heatshield
Aerothermodynamic Design of the Mars Science Laboratory Heatshield Karl T. Edquist∗ and Artem A. Dyakonovy NASA Langley Research Center, Hampton, Virginia, 23681 Michael J. Wrightz and Chun Y. Tangx NASA Ames Research Center, Moffett Field, California, 94035 Aerothermodynamic design environments are presented for the Mars Science Labora- tory entry capsule heatshield. The design conditions are based on Navier-Stokes flowfield simulations on shallow (maximum total heat load) and steep (maximum heat flux, shear stress, and pressure) entry trajectories from a 2009 launch. Boundary layer transition is expected prior to peak heat flux, a first for Mars entry, and the heatshield environments were defined for a fully-turbulent heat pulse. The effects of distributed surface roughness on turbulent heat flux and shear stress peaks are included using empirical correlations. Additional biases and uncertainties are based on computational model comparisons with experimental data and sensitivity studies. The peak design conditions are 197 W=cm2 for heat flux, 471 P a for shear stress, 0.371 Earth atm for pressure, and 5477 J=cm2 for total heat load. Time-varying conditions at fixed heatshield locations were generated for thermal protection system analysis and flight instrumentation development. Finally, the aerother- modynamic effects of delaying launch until 2011 are previewed. Nomenclature 1 2 2 A reference area, 4 πD (m ) CD drag coefficient, D=q1A D aeroshell diameter (m) 2 Dim multi-component diffusion coefficient (m =s) ci species mass fraction H total enthalpy -
Aerocapture FS-Pdf.Indd
NASA Facts National Aeronautics and Space Administration Marshall Space Flight Center Huntsville, Alabama 35812 FS-2004-09-127-MSFC July 2004 Aerocapture Technology NASA technologists are working to develop ways to place robotic space vehicles into long-duration, scientific orbits around distant Solar System destinations without the need for the heavy, on-board fuel loads that have historically inhibited vehicle performance, mission dura- tion and available mass for science payloads. Aerocapture -- a flight maneuver that inserts a space- craft into its proper orbit once it arrives at a planet -- is part of a unique family of “aeroassist” technologies under consideration to achieve these goals and enable robust science missions to any planetary body with an appreciable atmosphere. Aerocapture technology is just one of many propulsion technologies being developed by NASA technologists and their partners in industry and academia, led by NASA’s In-Space Propulsion Technology Office at the Marshall Space Flight Center in Huntsville, Ala. The Center implements the In-Space Propulsion Technolo- gies Program on behalf of NASA’s Science Mission Directorate in Washington. Aerocapture uses a planet’s or moon’s atmosphere to accomplish a quick, near-propellantless orbit capture to place a space vehicle in its proper orbit. The atmo- sphere is used as a brake to slow down a spacecraft, transferring the energy associated with the vehicle’s high speed into thermal energy. Aerocapture entering Mars Orbit. The aerocapture maneuver starts with an approach trajectory into the atmosphere of the target body. The dense atmosphere creates friction, slowing the craft and placing it into an elliptical orbit -- an oval shaped orbit. -
Magnetoshell Aerocapture: Advances Toward Concept Feasibility
Magnetoshell Aerocapture: Advances Toward Concept Feasibility Charles L. Kelly A thesis submitted in partial fulfillment of the requirements for the degree of Master of Science in Aeronautics & Astronautics University of Washington 2018 Committee: Uri Shumlak, Chair Justin Little Program Authorized to Offer Degree: Aeronautics & Astronautics c Copyright 2018 Charles L. Kelly University of Washington Abstract Magnetoshell Aerocapture: Advances Toward Concept Feasibility Charles L. Kelly Chair of the Supervisory Committee: Professor Uri Shumlak Aeronautics & Astronautics Magnetoshell Aerocapture (MAC) is a novel technology that proposes to use drag on a dipole plasma in planetary atmospheres as an orbit insertion technique. It aims to augment the benefits of traditional aerocapture by trapping particles over a much larger area than physical structures can reach. This enables aerocapture at higher altitudes, greatly reducing the heat load and dynamic pressure on spacecraft surfaces. The technology is in its early stages of development, and has yet to demonstrate feasibility in an orbit-representative envi- ronment. The lack of a proof-of-concept stems mainly from the unavailability of large-scale, high-velocity test facilities that can accurately simulate the aerocapture environment. In this thesis, several avenues are identified that can bring MAC closer to a successful demonstration of concept feasibility. A custom orbit code that dynamically couples magnetoshell physics with trajectory prop- agation is developed and benchmarked. The code is used to simulate MAC maneuvers for a 60 ton payload at Mars and a 1 ton payload at Neptune, both proposed NASA mis- sions that are not possible with modern flight-ready technology. In both simulations, MAC successfully completes the maneuver and is shown to produce low dynamic pressures and continuously-variable drag characteristics. -
Space Vehicle Conceptual Design Blue Team March 20, 2021
Hephaestus: Space Vehicle Conceptual Design Blue team March 20, 2021 Connor Cruickshank, Joel Lundahl, Birgir Steinn Hermannsson, Greta Tartaglia M.Sc. students, KTH Royal Institute of Technology, Stockholm, Sweden Abstract—A hypothetical contest has been issued to summit ms Structural mass Olympus Mons. This report details a conceptual design of a space P Power vehicle intended for that purpose. Functional requirements of the p Combustion chamber pressure vehicle were defined, and the primary subsystems identified. The 02 SpaceX Starship was chosen to serve as a design baseline and pe Exhaust gas exit pressure a comparison of suitable launch vehicles was made. Subsystems S Drag surface such as propulsion, reaction control, power generation, thermal T Thrust management and radiation shielding were considered, as well as ue Exhaust gas exit velocity interior design of the space vehicle and its protection measures W Weight for Mars entry and activities. The consequences of an engine-out scenario were studied. t Metric tonne The Starship launch system was deemed the most suitable launch vehicle candidate. Six Raptor engines were selected for the spacecraft main propulsion system, with 8 smaller thrusters I. INTRODUCTION for reaction control. A radiator mass of 890 kg was estimated In the year 2038 a challenge was issued for a team of for the thermal management system, and lithium metal hydride was determined an adequate radiation shielding material for the pioneers to be the first to reach the peak of Olympus Mons, the crew quarters. A radiation shelter was integrated into the pantry, highest mountain in the solar system. The reward for the first to and layouts of the crew quarters and cargo bay were prepared. -
Development of Supersonic Retropropulsion for Future Mars Entry, Descent, and Landing Systems
JOURNAL OF SPACECRAFT AND ROCKETS Vol. 51, No. 3, May–June 2014 Development of Supersonic Retropropulsion for Future Mars Entry, Descent, and Landing Systems Karl T. Edquist∗ and Ashley M. Korzun† NASA Langley Research Center, Hampton, Virginia 23681 Artem A. Dyakonov‡ Blue Origin, LLC, Kent, Washington 98032 Joseph W. Studak§ NASA Johnson Space Center, Houston, Texas 77058 Devin M. Kipp¶ Jet Propulsion Laboratory, California Institute of Technology, Pasadena, California, 91109 and Ian C. Dupzyk** Lunexa, LLC, San Francisco, California 94105 DOI: 10.2514/1.A32715 Recent studies have concluded that Viking-era entry system deceleration technologies are extremely difficult to scale for progressively larger payloads (tens of metric tons) required for human Mars exploration. Supersonic retropropulsion is one of a few developing technologies that may enable future human-scale Mars entry systems. However, in order to be considered as a viable technology for future missions, supersonic retropropulsion will require significant maturation beyond its current state. This paper proposes major milestones for advancing the component technologies of supersonic retropropulsion such that it can be reliably used on Mars technology demonstration missions to land larger payloads than are currently possible using Viking-based systems. The development roadmap includes technology gates that are achieved through ground-based testing and high-fidelity analysis, culminating with subscale flight testing in Earth’s atmosphere that demonstrates stable and controlled flight. The component technologies requiring advancement include large engines (100s of kilonewtons of thrust) capable of throttling and gimbaling, entry vehicle aerodynamics and aerothermodynamics modeling, entry vehicle stability and control methods, reference vehicle systems engineering and analyses, and high-fidelity models for entry trajectory simulations. -
Mars Exploration Entry, Descent and Landing Challenges1,2
Mars Exploration Entry, Descent and Landing Challenges1,2 Robert D. Braun Robert M. Manning Georgia Institute of Technology Jet Propulsion Laboratory Atlanta, GA 30332-0150 California Institute of Technology (404) 385-6171 Pasadena, CA 91109 [email protected] (818) 393-7815 [email protected] Abstract—The United States has successfully landed five interesting locations within close proximity (10’s of m) of robotic systems on the surface of Mars. These systems all pre-positioned robotic assets. These plans require a had landed mass below 0.6 metric tons (t), had landed simultaneous two order of magnitude increase in landed footprints on the order of hundreds of km and landed at sites mass capability, four order of magnitude increase in landed below -1 km MOLA elevation due the need to perform accuracy, and an entry, descent and landing operations entry, descent and landing operations in an environment sequence that may need to be completed in a lower density with sufficient atmospheric density. Current plans for (higher surface elevation) environment. This is a tall order human exploration of Mars call for the landing of 40-80 t that will require the space qualification of new EDL surface elements at scientifically interesting locations within approaches and technologies. close proximity (10’s of m) of pre-positioned robotic assets. This paper summarizes past successful entry, descent and Today, robotic exploration systems engineers are struggling landing systems and approaches being developed by the with the challenges of increasing landed mass capability to 1 robotic Mars exploration program to increased landed t while improving landed accuracy to 10’s of km and performance (mass, accuracy and surface elevation). -
Mars Microprobe Entry Analysis
Mars Microprob e Entry Analysis y Rob ert D. Braun Rob ert A. Mitcheltree F. McNeil Cheatwo o d NASA Langley Research Center NASA Langley Research Center Vigyan Inc. Hampton, VA 23681-0001 Hampton, VA 23681-0001 Hampton, VA 23666-1325 (757) 864-4507 (757) 864-4382 (757) 864-2984 [email protected] [email protected] f.m.cheatwo o [email protected] Abstract{The Mars Microprob e mission will 1 Introduction provide the rst opp ortunity for subsurface The ob jective of NASA's New Millennium pro- measurements, including water detection, near gram is to demonstrate and ight qualify tech- the south p ole of Mars. In this pap er, p er- nology elements required for the science mis- formance of the Microprob e aeroshell design is sions of the next century [1 ]. The program's evaluated through development of a six-degree- second ight pro ject, Deep Space Two (DS{ of-freedom (6-DOF) aero dynamic database and 2) is fo cused on the design of two small Mars ight dynamics simulation. Numerous mission entry prob es. As a result, DS{2 is often re- uncertainties are quanti ed and a Monte-Carlo ferred to as the Mars Microprob e mission. This analysis is p erformed to statistically assess mis- pro of-of-concept system is intended to demon- sion p erformance. Results from this 6-DOF strate key elements of future network science Monte-Carlo simulation demonstrate that, in a missions [2 , 3]. Attached to the cruise stage ma jority of the cases (approximately 2{ ), the of the Mars 98 Surveyor Lander, these two p enetrator impact conditions are within current Microprob e vehicles will b e launched to Mars design tolerances. -
Mars Polar Lander JPL Report Part 5 (Pdf)
8 SUMMARY OF POTENTIAL DS2 FAILURE MODES This section provides a summary description of the DS2 potential failure modes considered by the Board. It also provides a high-level assessment of the plausibility of each of the failure modes based on the Board’s review of the relevant design, implementation, and test history, and what may be inferred from data obtained throughout the mission operations. References are made to Sections 9, which contains detailed descriptions of the failure modes summarized in this section, along with findings and Lessons Learned. Failure modes are divided into two groups: ! Failure Modes Affecting Both Probes ! Failure Modes Affecting a Single Probe The four DS2 potential failure modes identified by the Board as plausible are highlighted as FLAG 1, FLAG 2, FLAG 3, and FLAG 4. Figure 8-1 illustrates the DS2 entry, descent, and impact (EDI) sequence and shows potential failure modes. Mars Polar Lander/Deep Space 2 Loss — JPL Special Review Board Report JPL D-18709 — page 124 • Probes experience battery drain prior to entry, descent, and impact sequence • Probes separate prematurely • Probes fail to separate from cruise stage (flaw in separation system) • Aeroshells fail (design flaw) • MPL fails to separate from cruise stage • MPL overheating, skip-out, excessive down-track entry points • Probes skip out or land too far down track; can’t • Random part failure communicate with MGS • Undetected aeroshell handling damage • Probes bounce on impact • Probes encounter surface conditions that exceed design capabilities • Structural failure on impact • Electronic or battery failure on impact • Ionization breakdown in Mars atmosphere • DS2 UHF link fails • Probes land on their sides, interfering with antenna performance Figure 8-1. -
Mars Lander Vehicle/Parachute Dynamics
The Space Congress® Proceedings 1968 (5th) The Challenge of the 1970's Apr 1st, 8:00 AM Mars Lander Vehicle/Parachute Dynamics R. D. Moog Denver, Colorado Martin Marietta Denver, Colorado Follow this and additional works at: https://commons.erau.edu/space-congress-proceedings Scholarly Commons Citation Moog, R. D. and Marietta, Martin, "Mars Lander Vehicle/Parachute Dynamics" (1968). The Space Congress® Proceedings. 3. https://commons.erau.edu/space-congress-proceedings/proceedings-1968-5th/session-10/3 This Event is brought to you for free and open access by the Conferences at Scholarly Commons. It has been accepted for inclusion in The Space Congress® Proceedings by an authorized administrator of Scholarly Commons. For more information, please contact [email protected]. MARS LANDER VEHICLE/PARACHUTE DYNAMICS R. D. Moog, Martin Marietta Denver, Colorado SUMMARY DISCUSSION Parachute decelerators used exclusively or in The Voyager mission involves entry into the combination with retro rockets have been consid Martian atmosphere from an orbiting spacecraft ered prime candidates for the terminal descent and of a scientifically instrumented payload for the landing system of a scientifically instrumented purpose of atmospheric determination and surface Mars lander. The objective of this study is to experimentation. A high drag entry vehicle com understand basic relationships between parameters bined with a terminal deceleration system involv affecting dynamic response of the parachute and ing either parachutes or retro rockets or both capsule and to define those aspects of the system will be used to achieve a soft landing on the which have a sensitive effect on the design of the planet surface.