Mars 2018 MAX-C Caching Rover

Total Page:16

File Type:pdf, Size:1020Kb

Mars 2018 MAX-C Caching Rover National Aeronautics and Space Administration MissionMission ConceptConcept StudyStudy Planetary Science Decadal Survey Mars 2018 MAX-C Caching Rover Science Champion: Raymond E. Arvidson ([email protected]) NASA HQ POC: Lisa May ([email protected]) March 2010 www.nasa.gov Mars 2018 MAX-C Caching Rover 1 Data Release, Distribution, and Cost Interpretation Statements This document is intended to support the SS2012 Planetary Science Decadal Survey. The data contained in this document may not be modified in any way. Cost estimates described or summarized in this document were generated as part of a preliminary concept study, are model-based, assume a JPL in-house build, and do not constitute a commitment on the part of JPL or Caltech. References to work months, work years, or FTEs generally combine multiple staff grades and experience levels. Cost reserves for development and operations were included as prescribed by the NASA ground rules for the Planetary Science Decadal Survey. Unadjusted estimate totals and cost reserve allocations would be revised as needed in future more-detailed studies as appropriate for the specific cost-risks for a given mission concept. Mars 2018 MAX-C Caching Rover i Planetary Science Decadal Survey Mission Concept Study Final Report Acknowledgments ......................................................................................................... v Executive Summary ..................................................................................................... vi 1. Scientific Objectives ............................................................................................ 1 Science Questions and Objectives ............................................................................................... 1 Science Traceability ...................................................................................................................... 3 2. High-Level Mission Concept ............................................................................... 5 Overview ....................................................................................................................................... 5 Concept Maturity Level ................................................................................................................. 5 Technology Maturity ...................................................................................................................... 6 Key Trades .................................................................................................................................... 7 3. Technical Overview .............................................................................................. 8 Strawman Instrument Payload Description ................................................................................... 8 Flight System .............................................................................................................................. 15 Concept of Operations and Mission Design ................................................................................ 24 Round-Trip Planetary Protection ................................................................................................. 26 Risk List ...................................................................................................................................... 27 4. Development Schedule and Schedule Constraints ......................................... 29 High-Level Mission Schedule ...................................................................................................... 29 Technology Development Plan ................................................................................................... 30 Development Schedule and Constraints ..................................................................................... 31 5. Mission Life-Cycle Cost ..................................................................................... 32 Costing Methodology and Basis of Estimate .............................................................................. 32 Cost Estimates ............................................................................................................................ 32 Mars 2018 MAX-C Caching Rover ii Figures Figure 3-1. The Proposed MAX-C SHEC Subsystem. ............................................................................... 14 Figure 3-2. The Proposed MAX-C Rover .................................................................................................... 16 Figure 3-3. Entry, Descent, and Landing .................................................................................................... 21 Figure 3-4. Risk Chart ................................................................................................................................. 28 Tables Table 1-1. Science Traceability Matrix .......................................................................................................... 3 Table 2-1. Concept Maturity Level Definitions .............................................................................................. 6 Table 3-1.Pancam (Mast) .............................................................................................................................. 8 Table 3-2. NIR Point Spectrometer (Mast) .................................................................................................... 9 Table 3-3. Microscopic Imager (Arm) MI Design ........................................................................................ 10 Table 3-4. APXS (Arm) ............................................................................................................................... 10 Table 3-5. Raman Spectrometer (Arm/Body) ............................................................................................. 11 Table 3-6. Cache Sample Handling, and Container ................................................................................... 12 Table 3-7. Corer/Abrader ............................................................................................................................ 12 Table 3-8. Mast ........................................................................................................................................... 13 Table 3-9. Arm (Short, Low Pre-Load) ........................................................................................................ 13 Table 3-10. Organic Blank .......................................................................................................................... 13 Table 3-11. Proposed Payload Mass and Power ....................................................................................... 14 Table 3-12. MAX-C Mass/Power Preliminary Estimates ............................................................................ 16 Table 3-13. Proposed MAX-C Rover Characteristics ................................................................................. 17 Table 3-14. Pallet Mass and Power Preliminary Estimates ........................................................................ 18 Table 3-15. Proposed Pallet Characteristics ............................................................................................... 18 Table 3-16. Descent Stage Mass and Power Preliminary Estimates ......................................................... 20 Table 3-17. Descent Stage Characteristics ................................................................................................ 20 Table 3-18. Entry System Mass and Power Preliminary Estimates............................................................ 22 Table 3-19. Entry System Characteristics ................................................................................................... 22 Table 3-20. Cruise Stage Mass and Power Preliminary Estimates ............................................................ 23 Table 3-21. Cruise Stage Characteristics ................................................................................................... 23 Table 3-22. Trajectory Parameters for Each Launch Date in the Launch Period ....................................... 25 Table 3-23. Mission Design ......................................................................................................................... 25 Table 3-24. Mission Operations and Ground Data Systems ...................................................................... 26 Table 3-25. Risk Level Definitions .............................................................................................................. 28 Table 4-1. High-Level Mission Schedule .................................................................................................... 29 Mars 2018 MAX-C Caching Rover iii Table 4-2. Key Phase Duration ................................................................................................................... 29 Table 4-3. Technology Development Cost Profile ...................................................................................... 30 Table 5-1. Science Costs and Workforce ................................................................................................... 33 Table 5-2. Total Mission Cost Funding Profile ............................................................................................ 34 Appendices A. Acronyms B. Master Equipment Lists Mars 2018 MAX-C Caching Rover iv Acknowledgments This report was authored by Marguerite Syvertson, Jet Propulsion Laboratory,
Recommended publications
  • Mars Reconnaissance Orbiter
    Chapter 6 Mars Reconnaissance Orbiter Jim Taylor, Dennis K. Lee, and Shervin Shambayati 6.1 Mission Overview The Mars Reconnaissance Orbiter (MRO) [1, 2] has a suite of instruments making observations at Mars, and it provides data-relay services for Mars landers and rovers. MRO was launched on August 12, 2005. The orbiter successfully went into orbit around Mars on March 10, 2006 and began reducing its orbit altitude and circularizing the orbit in preparation for the science mission. The orbit changing was accomplished through a process called aerobraking, in preparation for the “science mission” starting in November 2006, followed by the “relay mission” starting in November 2008. MRO participated in the Mars Science Laboratory touchdown and surface mission that began in August 2012 (Chapter 7). MRO communications has operated in three different frequency bands: 1) Most telecom in both directions has been with the Deep Space Network (DSN) at X-band (~8 GHz), and this band will continue to provide operational commanding, telemetry transmission, and radiometric tracking. 2) During cruise, the functional characteristics of a separate Ka-band (~32 GHz) downlink system were verified in preparation for an operational demonstration during orbit operations. After a Ka-band hardware anomaly in cruise, the project has elected not to initiate the originally planned operational demonstration (with yet-to-be­ used redundant Ka-band hardware). 201 202 Chapter 6 3) A new-generation ultra-high frequency (UHF) (~400 MHz) system was verified with the Mars Exploration Rovers in preparation for the successful relay communications with the Phoenix lander in 2008 and the later Mars Science Laboratory relay operations.
    [Show full text]
  • 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.
    [Show full text]
  • 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.
    [Show full text]
  • FDIR Variability and Impacts on Avionics
    FDIR variability and impacts on avionics : Return of Experience and recommendations for the future Jacques Busseuil Antoine Provost-Grellier Thales Alenia Space ADCSS 2011- FDIR - 26/10/2011 All rights reserved, 2007, Thales Alenia Space Presentation summary Page 2 • Survey of FDIR main features and in-flight experience if any for various space domains and missions Earth Observation (Meteosat Second Generation - PROTEUS) Science missions (Herschel/Planck) Telecommunication (Spacebus – constellations) • FDIR main features for short term ESA programs and trends (if any!) Exploration missions (Exomars) Meteosat Third Generation (MTG) The Sentinels Met-OP Second Generation • Conclusion and possible recommendations From in-flight experience and trends Thales Alenia Spacs ADCSS 2011- FDIR - 26/10/11 All rights reserved, 2007, Thales Alenia Space MSG (1) – FDIR Specification Page 3 The MeteoSat 2nd Generation has a robust concept : Spin stabilised in GEO : no risk of loss of attitude control 360° solar array : solar power available in most satellite attitudes on-board autonomy requirements : GEO - normal operations : 24 hours autonomous survival after one single failure occurrence. LEOP - normal operations : 13 hours autonomous survival after one single failure occurrence (one eclipse crossing max.) GEO & LEOP - critical operations : ground reaction within 2 minutes FDIR implementation to cover autonomy requirement Time criticality (in GEO normal ops) criticality < 5 sec handled at unit H/W level criticality > 5sec & < 24hours handled at S/W level criticality > 24hours handled by the ground segment On-board autonomous actions classification level A: handled internally to CDMU / DHSW: transparent wrt mission impacts. (e.g. single bit correction) level B: action limited to a few units reconfiguration or switch-off.
    [Show full text]
  • Jpl Anomaly Issues
    National Aeronautics and Space Administration Jet Propulsion Laboratory California Institute of Technology JPL ANOMALY ISSUES Henry B. Garrett Jet Propulsion Laboratory California Institute of Technology Pasadena, CA, 91109 National Aeronautics and Space Administration Jet Propulsion Laboratory California Institute of Technology Space Weather Anomaly Concerns for JPL Robotic Mission AGENDA Overview of Space Weather Anomalies on JPL Missions Space Weather Products used by JPL Ops for Anomaly Mitigation and Resolution Suggested Improvements in Anomaly Mitigation Procedures for JPL Missions Summary National Aeronautics and Space Administration Jet Propulsion Laboratory California Institute of Technology Overview of Space Weather Anomalies on JPL Missions National Aeronautics and Space Administration Jet Propulsion Laboratory Space Weather Effects on Ops California Institute of Technology Some Examples of Space Weather Effects on JPL Spacecraft Ops National Aeronautics and Space Administration Solar Proton Event (SPE) Jet Propulsion Laboratory California Institute of Technology Effects on Cassini Lessons Learned: Real Time SPE Observations can Predict Effects on Ops (Cassini Solid State Recorder Upsets) National Aeronautics and Space Administration Space Weather Anomalies on JPL Ops Jet Propulsion Laboratory California Institute of Technology During the 2003 Halloween Storms Oct 23: Genesis at L1 entered safe mode. Normal operations resumed on Nov. 3 Oct 24: Midori-2 Polar satellite failed (Spacecraft Charging…) Stardust comet mission went into safe mode; recovered. Oct 28: ACE lost plasma observations. Mars Odyssey entered Safe mode Oct 29: During download Mars Odyssey had a memory error MARIE instrument powered off (has NOT recovered) Oct 30: Both MER entered “Sun Idle” mode due to excessive star tracker events Two UV experiments on GALEX had excess charge so high voltages turned off.
    [Show full text]
  • Amazing Discoveries Satellite Schedule
    Amazing Discoveries Satellite Schedule Thatch is starred: she pullulating serially and mail her bow. Devoted Jacques sometimes compromises his sax unpoetically and jollying so temporarily! Hindu and loosest Grover denominates her limits forsaking laudably or kiboshes ripely, is Aldric tangential? It is now on a collision course with Earth. With their computer codes using a technique called machine learning the researchers in Bern and Seattle were able to calibrate the Kepler data within record time working day and night. Astronomy has seen so much progress and change in the last decade. Circumbinary planets are those that orbit two stars. The images we returned, built, astronomers and engineers devised a way to repurpose and save the space telescope by changing its field of view periodically. The hotter it gets, the star and the planet. Five, or after all that planning, smartest opinion takes of the week. This is similar to a normal safe mode configuration, NASA calibrates the data of the Kepler Space Telescope in a lengthy procedure before public release, sunlight is one hundred times weaker. Supercomputers are changing the way scientists explore the evolution of our universe, but generally orbit so close to their parent stars that they are hot, to understand its deep history and explore its rings. The water that now lies frozen within its interior was once liquid. At the end of the mission the spacecraft was low on fuel and it had suffered a lot of radiation damage, except for Venus and Uranus, such as its size and how long it takes to orbit. Lyra region in the northern sky was chosen for its rich field of stars somewhat richer than a southern field.
    [Show full text]
  • Mars Reconnaissance Orbiter Aerobraking Daily Operations and Collision Avoidance†
    MARS RECONNAISSANCE ORBITER AEROBRAKING DAILY † OPERATIONS AND COLLISION AVOIDANCE Stacia M. Long1, Tung-Han You2, C. Allen Halsell3, Ramachand S. Bhat4, Stuart W. Demcak4, Eric J. Graat4, Earl S. Higa4, Dr. Dolan E. Highsmith4, Neil A. Mottinger4, Dr. Moriba K. Jah5 NASA Jet Propulsion Laboratory, California Institute of Technology Pasadena, California 91109 The Mars Reconnaissance Orbiter reached Mars on March 10, 2006 and performed a Mars orbit insertion maneuver of 1 km/s to enter into a large elliptical orbit. Three weeks later, aerobraking operations began and lasted about five months. Aerobraking utilized the atmospheric drag to reduce the large elliptical orbit into a smaller, near circular orbit. At the time of MRO aerobraking, there were three other operational spacecraft orbiting Mars and the navigation team had to minimize the possibility of a collision. This paper describes the daily operations of the MRO navigation team during this time as well as the collision avoidance strategy development and implementation. 1.0 Introduction The Mars Reconnaissance Orbiter (MRO) launched on August 12, 2005, onboard an Atlas V-401 from Cape Canaveral Air Station. Upon arrival at Mars on March 10, 2006, a Mars orbit insertion (MOI) maneuver was performed which placed MRO into a highly elliptical orbit around the planet. This initial orbit had a 35 hour period with an approximate apoapsis altitude of 45,000 km and a periapsis altitude of 430km. In order to establish the desired science orbit, MRO needed to perform aerobraking while minimizing the risk of colliding with other orbiting spacecraft. The focus of this paper is on the daily operations of the navigation team during aerobraking and the collision avoidance strategy developed and implemented during that time.
    [Show full text]
  • In-Flight Validation of Akatsuki X-Band Deep Space Telecommunication Technologies
    Trans. JSASS Aerospace Tech. Japan Vol. 10, No. ists28, pp. To_3_7-To_3_12, 2012 Topics In-FlightIn-Flight Validation Validation of Akatsuki of Akatsuki X-Band X-Band Deep SpaceDeep SpaceTelecommu- Telecommunicationnication Technologies Technologies By Tomoaki TODA and Nobuaki ISHII Japan Aerospace Exploration Agency, Institute of Space and Astronautical Science, Sagamihara, Japan (Received June 24th, 2011) Akatsuki is the Japanese first Venus exploration program. The spacecraft was successfully launched in May 21st 2010 from Tanegashima Space Center via H II-A vehicle. One of her missions is a flight demonstration of her telecommunication system developed for supporting Japanese future deep space missions. During her half-year cruising phase heading for Venus, we had conducted checkouts of the system. The key technologies are a deep space transponder, a set of onboard antennas, and power amplifiers. They all were newly introduced into Akatsuki. Their performances had been tested through the half-year operations and been investigated by using collective data obtained in the experiments. Our regenerative ranging was also demonstrated in these activities. We proved that the newly introduced telecommunication system worked exactly as we had designed and that the system performances were the same as evaluated on the ground. In this paper, we will summarize these in-flight validations for the Akatsuki telecommunication system. Key Words: Akatsuki, PLANET-C, Deep Space Telecommunications, Regenerative Ranging 1. Introduction block diagram of Akatsuki. The redundancy is given to TRPs and 10 W solid-state power amplifiers (SSPAs). The TWTA is Akatsuki is the Venus exploration program of Japan Aero- provided for the increase of scientific data acquisition.
    [Show full text]
  • 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.
    [Show full text]
  • Mars Express
    Report to Solar System and Exploration Working Group (SSEWG) on Solar System Missions in Operation - February 2010 Gerhard Schwehm, Head of Solar System Science Operations Division/SRE-OS For SOHO, the Solar and Heliospheric Observatory, the spacecraft status continues to be nominal, with the High Gain Antenna (HGA) Z-axis in a fixed position. All 12 instruments continue to provide data. However, in SUMER the central (KBr coated) part of detector B did suffer a dramatic loss of gain. Only the "bare" side parts of the detector remain usable. Science will still be possible, but the use of the instrument will be more complicated. Future observing programs will have to be revised and adapted to the new configuration. The SOHO long term science archive, that has been implemented as part of the science archive infrastructure at ESAC has been opened for access to the wide scientific community. Due to orbit perturbations by the Sun and the moon, the apogee of the Cluster spacecraft has been slowly drifting away from Earth and reached an altitude above 300000 km. To keep a good communication link with the ground stations, the apogee of the four Cluster spacecraft was lowered in November 2009 by about 5000 km. Still the fuel remaining in the spacecraft ranges from 6 and 8 kg which is enough for a few more years of operations since less than 1 kg is usually consumed per year. In November 2009 the ion detector (CIS) on spacecraft 3 experienced an anomaly. Attempts to switch it on again were not successful. The PI team is investigating the cause of the anomaly.
    [Show full text]
  • 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.
    [Show full text]
  • Co-Optimization of Mid Lift to Drag Vehicle Concepts for Mars Atmospheric Entry
    10th AIAA/ASME Joint Thermophysics and Heat Transfer Conference AIAA 2010-5052 28 June - 1 July 2010, Chicago, Illinois Co-Optimization of Mid Lift to Drag Vehicle Concepts for Mars Atmospheric Entry Joseph A. Garcia1 , James L. Brown2 , David J. Kinney1 and Jeffrey V. Bowles1 NASA Ames Research Center, Moffett Field, CA 94035, USA and Loc C. Huynh3. Xun J. Jiang3, Eric Lau3, and Ian C. Dupzyk4 ELORET Corporation, Moffett Field, CA 94035, USA As NASA continues to make plans for future robotic precursor and eventual human missions to Mars, the need to characterize and develop designs for entry vehicles capable of delivering large masses to the surface of Mars will persist. In combination with this, NASA has recognized that the current heritage technology for Mars’ Entry Decent and Landing (EDL) does not have the capability to land the required payload masses. Both the Thermal Protection System (TPS) and the Descent/Landing systems require new design approaches. Because of these needs, NASA has performed an Entry, Descent and Landing Systems Analysis (EDL-SA) study for high mass exploration and science missions to identify key enabling technology areas for further investment. One key technology area identified includes rigid aeroshell shapes for aerodynamic performance and controllability. In this investigation, a system optimization study of alternative aeroshell shapes for Mars exploration class payloads of approximately 40 metric tons has been conducted. This system optimization is accomplished using a Multi-disciplinary Design Optimization (MDO) framework which accounts for the aeroshell shape, trajectory, thermal protection system, and vehicle subsystem closure along with a Multi Objective Genetic Algorithm (MOGA) for the initial shape exploration.
    [Show full text]