July 2010 

NASA/TM-2010-216720

Entry, Descent and Landing Systems Analysis 

Study: Phase 1 Report 

Alicia M. Dwyer Cianciolo, Jody L. Davis, David R. Komar, Michelle M. Munk, Jamshid A. Samareh, 

Julie A. Williams-Byrd, and Thomas A. Zang 

NASA Langley Research Center, Hampton, Virginia 

Richard W. Powell and Jeremy D. Shidner 

Analytical Mechanics Associates, Inc., Hampton, Virginia 

Douglas O. Stanley and Alan W. Wilhite 

Georgia Institute of Technology, Atlanta, Georgia 

David J. Kinney and M. Kathleen McGuire 

NASA Ames Research Center, Moffett Field, California 

James O. Arnold 

University of California, Santa Cruz, California 

Austin R. Howard 

ELORET Corporation, Sunnyvale, California 

Ronald R. Sostaric, Joseph W. Studak, and Carlie H. Zumwalt 

NASA Johnson Space Center, Houston, Texas 

Eduardo G. Llama 

GBTech, Inc., Houston, Texas 

Jordi Casoliva, Mark C. Ivanov, Ian Clark, and Anita Sengupta 

NASA Jet Propulsion Laboratory, Pasadena, California 
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July 2010 

NASA/TM-2010-216720

Entry, Descent and Landing Systems Analysis 

Study: Phase 1 Report 

Alicia M. Dwyer Cianciolo, Jody L. Davis, David R. Komar, Michelle M. Munk, Jamshid A. Samareh, 

Julie A. Williams-Byrd, and Thomas A. Zang 

NASA Langley Research Center, Hampton, Virginia 

Richard W. Powell and Jeremy D. Shidner 

Analytical Mechanics Associates, Inc., Hampton, Virginia 

Douglas O. Stanley and Alan W. Wilhite 

Georgia Institute of Technology, Atlanta, Georgia 

David J. Kinney and M. Kathleen McGuire 

NASA Ames Research Center, Moffett Field, California 

James O. Arnold 

University of California, Santa Cruz, California 

Austin R. Howard 

ELORET Corporation, Sunnyvale, California 

Ronald R. Sostaric, Joseph W. Studak, and Carlie H. Zumwalt 

NASA Johnson Space Center, Houston, Texas 

Eduardo G. Llama 

GBTech, Inc., Houston, Texas 

Jordi Casoliva, Mark C. Ivanov, Ian Clark, and Anita Sengupta 

NASA Jet Propulsion Laboratory, Pasadena, California 
Available from: 

NASA Center for AeroSpace Information 

7115 Standard Drive 

Hanover, MD 21076-1320 

443-757-5802 

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an official endorsement, either expressed or implied, by the National Aeronautics and Space 

Administration. 
i

Table of Contents 

Table of Contents...........................................................................................................................................i

List of Figures ..............................................................................................................................................iii

List of Tables................................................................................................................................................. v

Acronyms..................................................................................................................................................... vi

Acknowledgements...................................................................................................................................... xi

EDL-SA Team.............................................................................................................................................xii

Abstract ......................................................................................................................................................xiii

1 Background ............................................................................................................................................. 1

2 Exploration Class Architectures and Technologies................................................................................ 3

3 Exploration Class Component Model..................................................................................................... 5

3.1 Mass Models .................................................................................................................................... 5

3.1.1 Rigid Mid-L/D Aeroshell .......................................................................................................... 5

3.1.2 Hypersonic Inflatable Aerodynamic Decelerator (HIAD) ........................................................ 7

3.1.3 Supersonic Inflatable Aerodynamic Decelerator (SIAD).......................................................... 9

3.1.4 Supersonic/Subsonic Retro-Propulsion (SRP) .......................................................................... 9

3.2 Aerodynamic and Aerothermal Modeling...................................................................................... 13

3.2.1 Tool Set.................................................................................................................................... 13

3.2.2 Mid-L/D Geometry.................................................................................................................. 14

3.2.3 Flexible Low-L/D Vehicle (HIAD)......................................................................................... 15

3.3 Aerocapture/EDL Guidance Algorithms........................................................................................ 16

3.3.1 HYPAS (Hybrid Predictor-Corrector Aerocapture Scheme) .................................................. 16

3.3.2 Numerical Predictor Corrector Guidance (NPC)..................................................................... 18

3.3.3 Theoretical EDL Guidance ...................................................................................................... 19

3.3.4 Analytic Predictor-Corrector – Apollo-Derived Guidance ..................................................... 19

3.3.5 Guidance Comparison ............................................................................................................. 22

3.4 Retro-propulsion............................................................................................................................. 24

3.5 Thermal Protection Systems........................................................................................................... 27

3.5.1 HIAD – 23 m diameter............................................................................................................ 27

3.5.2 Mid L/D Aeroshell................................................................................................................... 28

4 Simulation Results................................................................................................................................ 32

4.1 Aerocapture Simulation Results..................................................................................................... 32

4.1.1 Aerocapture Fligth Path Corridor Sizing................................................................................. 33

4.1.2 Nominal Trajectory Results..................................................................................................... 34

4.1.3 Monte Carlo Results................................................................................................................ 36

4.1.4 Sensitivity Study Results......................................................................................................... 37

4.1.5 Trades ...................................................................................................................................... 38

4.2 Entry, Descent and Landing ........................................................................................................... 41

4.2.1 Comparison of EDL-SA to DRA5........................................................................................... 41

4.2.2 EDL Strategy ........................................................................................................................... 42

4.2.3 Nominal and Monte Carlo Results .......................................................................................... 44

4.2.4 Sensitivities Study Results....................................................................................................... 52

5 Packaging.............................................................................................................................................. 55

5.1 Preliminary Packaging ................................................................................................................... 55

5.2 HIAD Packaging ............................................................................................................................ 55

6 Figures of Merit .................................................................................................................................... 58

6.1 Assessment Process........................................................................................................................ 58

6.2 Figure of Merit Weighting ............................................................................................................. 61

6.3 Assessment Results........................................................................................................................ 61
ii

6.3.1 Architecture Assessment Results............................................................................................. 61

6.3.2 Technology Assessment Results ............................................................................................. 63

6.4 Summary ........................................................................................................................................ 64

7 Viability Assessments........................................................................................................................... 65

7.1 Dual-Layer TPS.............................................................................................................................. 65

7.2 IAD Controllability ........................................................................................................................ 68

8 Technology Investment Recomendations............................................................................................. 69

9 Exploration-Class Study Open Issues................................................................................................... 70

10 Large Robotic-Class Study ................................................................................................................. 71

10.1 Packaging Feasibility..................................................................................................................... 71

10.2 Parachutes...................................................................................................................................... 71

10.2.1 Modeling Overview............................................................................................................... 72

10.2.2 Deployment Modeling........................................................................................................... 72

10.2.3 Parachute Drag Performance ................................................................................................. 74

10.3 Monte Carlo Simulation Results.................................................................................................... 77

10.4 Robotic Class Summary................................................................................................................. 79

11 Bibliography.......................................................................................................................................... 80

A Appendix A. DRMs, GR&As and FOMs............................................................................................... 83

A.1 Design Reference Missions .............................................................................................................. 83

A.1.1 Exploration-Class Design Reference Mission ........................................................................... 83

A.1.2 Large-Robotic Class Reference Mission................................................................................... 85

A.2 EDL-SA Study Ground Rules and Assumptions.............................................................................. 88

A.2.1 General Ground Rules and Assumptions................................................................................... 88

A.2.2 Exploration Mission Ground Rules and Assumptions............................................................... 88

A.2.3 Robotic Mission Ground Rules and Assumptions..................................................................... 88

A.2.4 Margin Policy............................................................................................................................. 89

A.2.5 Exploration Mission Monte Carlo Parameters........................................................................... 91

A.2.6 Robotic Mission Monte Carlo Parameters................................................................................. 94

A.3 EDL-SA Study Figures of Merit....................................................................................................... 95

A.3.1 Description of Figures of Merit ................................................................................................. 95

A.3.2 Categories for Figures of Merit.................................................................................................. 95

A.3.3 Definitions of Figures of Merit .................................................................................................. 96

A.3.4 Scoring for Figures of Merit ...................................................................................................... 96

A.3.5 FOM Proxy Parameters............................................................................................................. 97

B Appendix B: EDL Transition Events.................................................................................................... 100

B.1 Background ..................................................................................................................................... 100

B.2 Methodology ................................................................................................................................... 100

B.3 Summary of Results........................................................................................................................ 100

B.3.1 Rigid Aeroshell to SRP ........................................................................................................... 101

B.3.2 Inflatable Aeroshell to SRP...................................................................................................... 102
iii

List of Figures 

Figure 1. Study Flow.................................................................................................................................... 1

Figure 2. Exploration Class Architectures .................................................................................................... 4

Figure 3. EDL-SA Major Mass Components................................................................................................ 5

Figure 4. Finite-Element Model of the Rigid Aeroshell .............................................................................. 6

Figure 5. Structural (left) and TPS (right) Mass for the Rigid Aeroshell of Architecture 1 ......................... 7

Figure 6. Reinforcing Fibers Invalid source specified. ................................................................................. 8 

Figure 7. Structural (left) and TPS (right) Mass for the HIAD of Architecture 2 ........................................ 8

Figure 8. SRP Specific Impulse vs. Engine Thrust-to-Weight ................................................................... 11

Figure 9. Mid L/D (left) and HIAD (right) Vehicle Configurations........................................................... 13

Figure 10. Representative DPLR Solution.................................................................................................. 14

Figure 11. Representative CBAERO Solution............................................................................................ 14

Figure 12. Representative CART3D Solution............................................................................................. 14

Figure 13. Anchored CBAERO Solution (left) and Subsonic CART3D Solution (right).......................... 16

Figure 14. Guidance comparison for Architecture 1................................................................................... 23

Figure 15. Guidance comparison for Architecture 8................................................................................... 23

Figure 16. SRP Sample Throttle Profile (left) and SRP ISP vs Expantion Ratio (right)............................ 25

Figure 17. Notional Tank Arrangement ...................................................................................................... 25

Figure 18. HIAD Material Stack ................................................................................................................. 27

Figure 19. HIAD Dual Pulse Surface Heating History and Integrated Heatload Contour.......................... 28

Figure 20. Mid L/D Aeroshell TPS Stack ................................................................................................... 29

Figure 21. Mid L/D Aeroshell Material Splitlines...................................................................................... 29

Figure 22. Mid L/D Aeroshell Dual Pulse Surface Heating History and Integrated Heatload Contour..... 30

Figure 23. Aerocapture Phases.................................................................................................................... 33

Figure 24. Available Skip-out Margin for Variations of L/D and Ballistic Number.................................. 34

Figure 25. Altitude Profiles for All Configurations.................................................................................... 35

Figure 26. Nominal Bank Profiles for the 23 m HIAD............................................................................... 35

Figure 27. Nominal Drag Acceleration and Attitude Rate Profile for 23 m HIAD .................................... 35

Figure 28. Monte Carlo Results for the Rigid Mid L/D Aeroshell ............................................................. 36

Figure 29. Monte Carlo Results for the 23 m HIAD, L/D = 0.3................................................................ 37

Figure 30. Monte Carlo Results for the 55 m HIAD, L/D = 0.2................................................................ 37

Figure 31. Delta V Sensitivity to Post-Aerocapture Orbit ......................................................................... 38

Figure 32. Available Skip-out Margin for Low Combinations of L/D and Ballistic Number.................... 39

Figure 33. Delta V Monte Carlo Results.................................................................................................... 39

Figure 34. Plot of 3 sigma peak heat rate versus diameter for lines of constant L/D ................................. 40

Figure 35. Effect of Flight Path Angle Dispersions for L/D=0.2, Bn=33.33 kg/m2 ................................... 41

Figure 36. EDL-SA Exploration Class Altitude Versus Velocity............................................................... 44

Figure 37. Architecture 1 Altitude vs. Time .............................................................................................. 45

Figure 38. Architecture 2 Altitude vs. Time .............................................................................................. 45

Figure 39. Rocket Plume Hypersonic Flowfield Interactions.................................................................... 46

Figure 40. Architecture 3 Arrival Mass and Delta V .................................................................................. 47

Figure 41. Architecture 4 Altitude vs. Time ............................................................................................... 48

Figure 42. Architecture 6 Altidue vs. Time ................................................................................................ 49

Figure 43. Architecture 7 Altitude vs. Time ............................................................................................... 49

Figure 44. Architecture 8 Altitude vs. Time ............................................................................................... 50

Figure 45. Architecture 2 Packaging........................................................................................................... 55

Figure 46. First Attempt at HIAD TPS Packaging...................................................................................... 56

Figure 47. Second Attempt at Packaging the 23 m “Apollo” HIAD .......................................................... 57

Figure 48. FOMs, Sub-FOMs and Sources Used to Evaluate..................................................................... 58
iv

Figure 49. FOM Assessment Process.......................................................................................................... 59

Figure 50. Programmatic Risk FOM Process Flow ................................................................................... 60

Figure 51. Average FOM Category Weights.............................................................................................. 61

Figure 52. Architecture Score from FOM Weigth Surveys and Sensitivity Studies................................... 62

Figure 53. Technology Score Ranges from FOM Weight Surveys and Sensitivity Study ......................... 63

Figure 54. (a) curved panel covered with ~5cm honeycomh filled with PICA blocks; (b) post tested 

stagnation iso-q model after exposure to 620 W/cm2 for 30 seconds; and (c) post tested swept 

cylinder model after exposure at 140 W/cm2 for 25 seconds.............................................................. 65

Figure 55. Iso-Q Model Design of PICA atop LI-900. Dimensions in inches........................................... 66

Figure 56. Test Thermal Cycle to Simulate Dual Heat Pulse ..................................................................... 66

Figure 57. Post tested model at the IHF run condition................................................................................ 67

Figure 58. Post tested model at the IHF run condition................................................................................ 67

Figure 59. : Clips from a movie taken during the second heat pulse simulating entry from out-of-Mars 

orbit after aerocapture .......................................................................................................................... 68

Figure 60. Parachute systems considered. .................................................................................................. 72

Figure 61. High and Low Drag Performance Curves Used for DGB Modeling......................................... 74

Figure 62. Summary of supersonic Ringsail flight test drag data ............................................................... 76

Figure 63. Summary of supersonic Ringsail flight test data with high and low drag performance model 76

Figure 64. Monte Carlo simulation results using 8000 dispersed cases for the reefed parachute 

configuration. ....................................................................................................................................... 78

Figure A-1. Exploration-Class Design Reference Missions ....................................................................... 83

Figure A-2. Mars Transfer Vehicles ........................................................................................................... 85

Figure A-3. Outher Mold Line of a Mars Lander Aeroshell....................................................................... 85

Figure A-4. Large Robotic-Class Design Reference Mission..................................................................... 86

Figure A-5. MSL and Large Robotic-Class Mission Landing Strategies................................................... 87

Figure B-1. Rigid Aeroshell to SRP Transition Option ............................................................................ 101

Figure B-2. Baseline Rigid Aeroshell to SRP Transition Options............................................................ 102

Figure B-3. Inflatable Aeroshell to SRP Transition Option...................................................................... 102

Figure B-4. Baseline Inflatable Decelerator to SRP Option ..................................................................... 103

Figure B-5. Inflatable Hypersonic Aeroshell to Supersonic Inflatable Transition Options...................... 104

Figure B-6. Baseline Hypersonic Inflatable Decelerator to Supersonic Inflatable Decelerator Option ... 105

Figure B-7. Rigid Aeroshell to Inflatable Aerodynamic Decelerator Options ......................................... 105

Figure B-8. Baseline Rigid Aeroshell to Inflatable Aerodynamic Decelerator ........................................ 106
v

List of Tables 

Table 1. Simplified Set of Exploration Class Technologies Considered by EDL-SA.................................. 4

Table 2. Nominal Parameters and Mass Breakdown for Architecture 1....................................................... 7

Table 3. Nominal Parameters and Mass Breakdown for Architecture 2....................................................... 9

Table 4. SRP Independent Variables and Limits for Response Surface Equations.................................... 10

Table 5. SRP Dependent Variables............................................................................................................. 10

Table 6. SRP Mass Breakdown for Architectures 1 and 3.......................................................................... 12

Table 7. SRP Engine Characteristics.......................................................................................................... 26

Table 8. TPS Sizing Margins ...................................................................................................................... 27

Table 9. Mid L/D Aeroshell TPS Mass Estimates (Not including attachment or WGA)........................... 30

Table 10. Mid L/D Aeroshell, Windward Side Estimated Masses with Attachment.................................. 31

Table 11. Aerocapture Monte Carlo Dispersion Values............................................................................. 36

Table 12. Architecture Component Separation Times................................................................................ 43

Table 13. EDL-SA Monte Carlo Dispersions ............................................................................................ 44

Table 14. Architecture 1 Results................................................................................................................. 45

Table 15. Architecture 2 Results ................................................................................................................. 45

Table 16. Architecture 4 Results................................................................................................................. 48

Table 17. Architecture 6 Results................................................................................................................. 49

Table 18. Architecture 7 Results................................................................................................................. 49

Table 19. Architecture 8 Results................................................................................................................. 50

Table 20. Component Dimensions and Masses using Standard Margins................................................... 51

Table 21. Nominal Environment Conditions .............................................................................................. 52

Table 22. Environmental Sensitivity........................................................................................................... 53

Table 23. Performance Sensitivity .............................................................................................................. 53

Table 24. Design Sensitivity ....................................................................................................................... 54

Table 25. Contributors to the FOM Assessment......................................................................................... 59

Table 26. TDP Recommendations .............................................................................................................. 69

Table 27. Summary of Supersonic Ringsail Parachute Inflation Times..................................................... 73

Table 28. Summary of Supersonic Ringsail Test Conditions ..................................................................... 75

Table 29. Supersonic Parachute Performance Results. ............................................................................... 77

Table 30. Supersonic Reefed Parachute Performance Results ................................................................... 78

Table 31. Comparison of MSL and the Scaled Reefed Ringsail Configuration ......................................... 79

Table A-1. Key Parameters for the Exploration-Class DRM...................................................................... 84

Table A-2. Key Parameters fo the Large Robotic-Class DRM................................................................... 87

Table A-3. EDL-SA Subsystem Margins.................................................................................................... 90

Table A-4. TRL Margin Increments ........................................................................................................... 91

Table A-5. Mid-L/D Rigid Monte Carlo Parameters.................................................................................. 91

Table A-6. Low-L/D Inflatable Monte Carlo Parameters........................................................................... 93

Table A-7. Robotic Monte Carlo Parameters.............................................................................................. 94

Table A-8. Definitions of Figures of Merit................................................................................................. 96

Table A-9. Safety and Mission Success Proxy Parameters......................................................................... 97

Table A-10. Performance and Effectiveness Proxy Parameters ................................................................. 98

Table A-11. Programmatic Risk Proxy Parameters.................................................................................... 98

Table A-12. Affordability and Life Cycle Cost Proxy Parameters............................................................. 99

Table A-13. Applicability to Other Mission Proxy Parameters.................................................................. 99

Table B-1. EDL Failure Modes................................................................................................................. 106
vi

Acronyms 

AIAA American Institute for Aeronautics and Astronautics 

ADP Advanced Development Program 

AFE Aeroassist Flight Experiment 

AHF Aerodynamic Heating Facility 

ALHAT Autonomous Landing and Hazard Avoidance Technology 

Al-Li Aluminum Lithium 

Alpha (α) Angle of Attack 

APAS Aerodynamic Preliminary Analysis System 

APC Analytic Predictor-Correction

ARA Analytical Research Associates 

ARC Ames Research Center 

ARMD Aeronautics Research Mission Directorate 

AS Aeroshell 

Beta (β) Side Slip Angle 

BC Ballistic Coefficient 

BLDT Balloon Launched Decelerator Test 

BN Ballistic Number 

CCRB Customer Content Review Board 

CBAERO Engineering code to compute aero and aerthermodynamics 

CECE Common Extensible Cryogenic Engine 

CEV Crew Exploration Vehicle 

CFD Computational Fluid Dynamics 

CG Center of Gravity 

CNES Centre National d'Études Spatiales 

CxP Constellation Program 

DGB Disk Gap Band 

deg degrees 

deg/sec2

 degrees per second squared 

DGB Disk Gap Band 

DOF Degree of Freedom 

DP Dual Pulse 

DPLR Data Parallel Line Relaxation 
vii

DRA5 Design Reference Architecture 5.0 

DRM Design Reference Mission 

DSIAD Drag Supersonic Inflatable Aerodynamic Decelerator 

EDL Entry, Descent and Landing 

EDL-SA Entry, Descent and Landing Systems Analysis 

EFPA Entry Flight Path Angle 

EPR External Peer Review 

ESAS Exploration Systems Architecture Study 

ESMD Exploration Systems Mission Directorate 

ETDP Exploration Technology Development Program 

EXAMINE Exploration Architecture Model for the IN-space and Earth-to-orbit 

FE Finite Element 

FIAT Code to model TPS thermal response 

FOM Figure of Merit 

FY Fiscal Year 

GRAM Global Reference Atmosphere Model 

GRC Glenn Research Center 

GR&A Ground Rule and Assumption 

H/C Honeycomb 

HIAD Hypersonic Inflatable Aerodynamic Decelerator 

HSRI Human-Systems Integration Requirements 

HT424 TPS adhesive used on Orion, Apollo, ARA H/C 

HYPAS Hybrid Predictor-Corrector Aerocapture Scheme 

Hz Hertz 

IAD Inflatable Aerodynamic Decelerator 

IHF Interactive Heating Facility 

IMLEO Initial Mass in Low Earth Orbit 

in inch 

IRDT Inflatable Reentry and Descent Technology 

Iso-q Constant Dynamic Pressure 

Isp Specific Impulse 

ISRU In-situ Resource Utilization 

JANNAF Joint Army, Navy, NASA Airforce 

JPL Jet Propulsion Laboratory 
viii

JSC Johnson Space Center 

kg kilograms

kg/m2 kilograms per square meter 

klbf kilopounds of force

km kilometers 

kN kilo Newtons 

LaRC Langley Research Center 

lbf pounds of force 

LADAR Laser Radar 

LAURA Langley Aerothermodynamic Upwind Relaxation Algorithm 

LCH4 Liquid Methane 

LEO Low Earth Orbit 

LHIAD Lifting Hypersonic Inflatable Aerodynamic Decelerator 

LH2 Liquid Hydrogen 

LM Lunar Module 

LOX Liquid Oxygen 

Ls Solar Longitude 

LSIAD Lifting Supersonic Inflatable Aerodynamic Decelerator 

LoD Length-to-Diameter 

L/D Lift-to-Drag Ratio 

m meters 

m/s meters/second 

MER Mass Estimating Relationship 

MIAS Mars Inflatable Aeroshell entry System 

MJ Mega Joules 

MMOD MicroMeteroids and Orbital Debris 

MOLA Mars Orbiter Laser Altimeter 

MPS Main Propulsion System 

MSL Mars Science Laboratory 

mt metric ton 

MTV Mars Transfer Vehicle 

NA Not Applicable 

NASA National Aeronautics and Space Administration 

NPC Numerical Predictor-Corrector 
ix

NTR Nuclear Thermal Rocket

OCE Office of the Chief Engineer 

OMS Orbital Maneuvering System 

PDF Probability Density Function 

PICA Phenolic Impregnated Carbon Ablator 

POST2 Program to Simulate Optimized Trajectories (version 2) 

PPL Pin Point Landing 

Psia Pounds Per Square Inch in Air 

q Dynamic Pressure 

RCS Reaction Control System 

RD3 Research & Development Degree of Difficulty 

RFA Request for Action 

RP Retro-propulsion 

RPS Radioisotope Systems 

RSE Response Surface Equation 

RTG Radioisotope Thermoelectric Generators 

RTV Room Temperature Vulcanized 

SBIR Small Business Innovative Research 

sec seconds 

SIAD Supersonic Inflatable Aerodynamic Decelerator 

SIP Strain Insulation Pad (Shuttle TPS) 

SIRCA Silicone Impregnated Reusable Ceramic Ablator 

SLA Super Lightweight Ablator 

SMC Strategic Management Council

SMD Science Mission Directorate 

SOA State of the Art 

SRP Supersonic Retro-propulsion 

TDP Technology Development Program 

TIM Technical Interchange Meeting 

TM Technical Memorandum 

TPS Thermal Protection System 

TRL Technology Readiness Level 

T/W Thrust-to-Weight Ratio 

UHF Ultra High Frequency 
x

XCG Center of Gravity Location Along the X Body Axis 

YCG Center of Gravity Location Along the Y Body Axis 

ZCG Center of Gravity Location Along the Z Body Axis 

ΔV Delta velocity 

σ standard deviation 
xi

Acknowledgements 

Many individuals, both inside and outside NASA, have contributed to the results of the EDL Systems 

Analysis Study. We first acknowledge the following individuals who participated directly in EDL-SA 

activities at some time during the course of this study: Karl T. Edquist, Richard T. French, Aga M. 

Goodsell, David B. Hash, Tuyen Hua, Ian C. Dupzyk, Eric M. Queen, Adam D. Steltzner, John J. 

Wagner, Carlos H. Westhelle, Douglas S. Whitehead, Henry S. Wright, and Michael J. Wright. 

Next, we appreciate the contributions of the following technology program managers who provided 

guidance on our Design Reference Missions, Ground Rules and Assumptions and Figures of Merit: David 

J. Anderson, Samad A. Hayati, Frank Peri and James L. Pittman. 

The identification of the transition mechanisms was assisted by Douglas S. Adams, Charles H. 

Campbell, Allen Chen, Ian Clark, Chirold D. Epp, Karl T. Edquist, Richard T. French, John C. Gallon, 

Carl S. Guernsey, Steve J. Hoffman, Christian de Jong, Devin M. Kipp, Ashley Korzun, Chris M. 

Madsen, Donald R. Pettit, Tommaso P. Rivellini, A. Miguel San Martin, Steven W. Sell, Adam D. 

Steltzner, Mike Tigges, and Aron A. Wolf, 

The data used in the assessment of the Programmatic Cost & Risk Figures of Merit were provided by 

Robin A. Beck, Anthony M. Calomino, Charles H. Campbell, F. McNeil Cheatwood, Karl T. Edquist, 

Stephen J. Hughes and Charles J. Player. 

The data used in the assessment of the Applicability to Other Missions Figures of Merit were 

provided by David J. Anderson, John W. Dankanich, Samad A. Hayati, Tommaso P. Rivellini, Adam D. 

Steltzner and Michael J. Wright. 

The weighting of the Figures of Merit was performed by F. McNeil Cheatwood, Douglas Craig, 

Robert Dillman, Bret G. Drake, Dana C. Gould, Samad A. Hayati, James J. Reuther, L. Dale Thomas, 

Ethiraj Venkatapathy, Henry S. Wright (and Thomas A. Zang). 

The preliminary EDL-SA results were critiqued in an Internal Peer Review by F. McNeil Cheatwood, 

Allen Chen, Juan R. Cruz, Dean A. Kontinos, Charles J. Player, Alan L. Strahan, Ethiraj Venkatapathy, 

David W. Way and Paul F. Wercinski.. 

The External Peer Review was performed by Mark Adler, Robert Braun, Anthony M. Calomino, John 

Connolly, James M. Corliss, Steven G. Labbe and Joseph Olejniczak. 

The team is appreciative of the efforts of Farhad Tahmasabi and Harold M. Bell of the Office of the 

Chief Engineer for their guidance and support, and especially for the efforts of our sponsor, Mike 

Ryschkewitsch, the NASA Chief Engineer. 

Finally, The EDL Systems Analysis Study was only possible because of the persistent and effective 

advocacy of Walter C. Engelund and Stephen P. Sandford. 
xii

EDL-SA Team 

Gary A. Allen, Jr. 

James O Arnold 

Jeffery V. Bowles 

Jordi Casoliva 

George T. Chen 

Ian G. Clark 

Melmoth A. Covington 

Jody L .Davis 

Ian C. Dupzyk 

Alicia M. Dwyer Cianciolo 

Karl T. Edquist 

Walter C. Engelund 

Wenhong Fan 

Richard T. French 

Joseph A. Garcia 

Howard E. Goldstein 

Aga M. Goodsell 

Sergey Gorbunov 

David B. Hash 

Austin R. Howard 

Tuyen Hua 

Mark C. Ivanov 

David J. Kinney 

David R. Komar 

Mary E. Livingston 

Eduardo G. Llama 

M. Kathleen McGuire 

Michelle M. Munk 

Richard W. Powell 

Dinesh K. Prabhu 

Eric M. Queen 

Jamshid A. Samareh 

David A. Saunders 

Anita Sengupta 

Stephen A. Sepka 

Jeremy D. Shidner 

Ronald R. Sostaric 

Margaret M Stackpoole 

Douglas O. Stanley 

Adam D. Steltzner 

Joseph W. Studak 

Farhad Tahmasebi 

Ethiraj Venkatapathy 

John J. Wagner 

David W. Way 

Carlos H. Westhelle 

Susan M. White 

Douglas S. Whitehead 

Alan W. Wilhite 

Julie A. Williams-Byrd 

Henry S. Wright 

Michael J. Wright 

Thomas A. Zang 

Carlie H. Zumwalt 
xiii

Abstract 

NASA senior management commissioned the Entry, Descent and Landing Systems Analysis (EDLSA) Study in 2008 to identify and roadmap the Entry, Descent and Landing (EDL) technology 

investments that the agency needed to make in order to successfully land large payloads at Mars for both 

robotic and human-scale missions. This paper summarizes the motivation, approach and top-level results 

from Year 1 of the study, which focused on landing 10–50 mt on Mars, but also included a trade study of 

the best advanced parachute design for increasing the landed payloads within the EDL architecture of the 

Mars Science Laboratory (MSL) mission. 

The EDL-SA Study focused on Exploration-class missions in FY 09, i.e., cargo or crewed missions 

requiring between 10 and 50 mt of landed payload. Candidate technology areas were assessed against a 

set of eight EDL-SA Architectures, i.e., representative architectures (high-level designs) against which the 

benefits of the technology areas were evaluated. The Study used Design Reference Missions (DRMs), 

Ground Rules & Assumptions (GR&As) and Figures of Merit (FOMs) that were approved by the 

managers of the relevant NASA technology programs in May 2009, prior to the execution of the 

simulations and the evaluations of the FOMs. In evaluating the FOMs, the Study used simulation-based 

results whenever possible and subjective assessments otherwise. The major simulation-based result was 

the Mars Arrival Mass, i.e., the total mass of the payload plus the systems needed for Mars Orbit Insertion 

and Mars EDL. 

The key enabling technology areas identified for investment were rigid decelerators, flexible 

(inflatable) decelerators, supersonic retro-propulsion, precision landing with hazard avoidance, 

aerocapture, and all-propulsive EDL systems. 
1

1 Background 

The steps in the study for the Exploration-class missions are shown in Figure 1. The NASA Strategic 

Management Council (SMC) commissioned the Entry, Descent and Landing Systems Analysis (EDL-SA) 

Study in May 2008 to identify and roadmap the Entry, Descent and Landing (EDL) technology 

investments that the agency needed to make in order to successfully land large payloads at Mars for both 

robotic and human-scale missions. This decision and the governance structure for its implementation are 

described in internal Agency documents—the Program Decision Memorandum [1] and the Terms of 

Reference [2] respectively, from August 2008. The EDL-SA Team was composed of members from four 

NASA centers—Ames Research Center (ARC), Jet Propulsion Laboratory (JPL), Johnson Space Center 

(JSC) and Langley Research Center (LaRC). The Aeronautics Research Mission Directorate (ARMD), 

the Exploration Systems Mission Directorate (ESMD), the Science Mission Directorate (SMD) and the 

Office of the Chief Engineer (OCE) provided funding and oversight. LaRC led the EDL-SA Team, and 

the oversight was coordinated by the OCE. The project plan was approved in October 2008. Technical 

work on the project commenced in November 2008. 

Figure 1. Study Flow 

The EDL-SA Team held a brainstorming session in November 2008 to identify candidate 

technologies for consideration. The context for the technology brainstorming was the Mars Design 

Reference Architecture 5.0 [3] (DRA5). In keeping with standard practice in systems analysis for 

technology evaluation, the technologies were assessed against the suite of EDL-SA Architectures. The 

set of EDL-SA Architectures only needs to include options that encompass all candidate technology 

areas. The architectures and technologies are described in Section 3. 

The Report of the Design Reference Missions, Ground Rules and Assumptions, and Figures of Merit 
2

present the basic guidelines for the study and are provided in Appendix A. The Design Reference 

Mission was taken to be the Aerocapture and EDL Phases of DRA5. The Figures of Merit are described 

in Section 6. 

To the maximum extent possible existing processes and tools were employed in the EDL-SA Study. 

Nevertheless, numerous refinements to processes and tools were needed in order to address the specifics 

of this study. Details on the models and simulations utilized are given in Section 3. The initial 

simulations and FOM data gathering were performed, and then an External Peer Review (EPR) in 

September 2009 provided recommendations for improvement. Most of the Requests for Action from this 

EPR were implemented; however, resource constraints and the priority given by the oversight committee 

on concentrating on the Large-robotic-class Study in FY 2010 left some recommended work incomplete. 

Recommendations for follow-on work are provided in Section 9. The Exploration-class Study was 

wrapped up in February 2010; the principal results from it are summarized in Section 4 to Section 7. 

Some preliminary work on Large-robotic-class missions was performed the first year. See Section 10 for 

the formulation, methodology and results. 
3

2 Exploration Class Architectures and Technologies 

The EDL-SA Team held a brainstorming session at its November 2008 Workshop to identify 

candidate technologies for consideration. The context for the technology brainstorming was the 

Exploration-class DRM, namely, the Aerocapture and EDL phases of the Mars Design Reference 

Architecture 5.0 [3] (DRA5). In fact, one of the Ground Rules for the Study permits only changes to the 

Aerocapture and EDL phases of DRA5. The baseline EDL architecture—Architecture 1—was that of 

DRA5: a rigid, mid-L/D aeroshell used for aerocapture and hypersonic entry deceleration, followed by 

supersonic retro-propulsion. Alternatives to this concept were then identified using a common “tree” 

approach, and refined by engineering judgment and the availability of reasonable models. 

The technology areas that emerged from this brainstorming and subsequent refinement were 

* Rigid Mid Range Lift to Drag Ratio Aeroshell (Rigid Mid-L/D AS) 

* Lifting Hypersonic Inflatable Aerodynamic Decelerator (LHIAD) 

* Supersonic Retro-Propulsion (SRP) 

* Drag Supersonic Inflatable Aerodynamic Decelerator (DSIAD) 

* Deployable Lifting Supersonic IAD With Skirt (LSIAD–Skirt) 

* Dual-Pulse Thermal Protection System (DP TPS–Flexible and DP TPS–Rigid) 

The working example of a Rigid Mid-L/D Aero shell is the “ellipsled” used in DRA5, which has an 

elliptical (current analysis assumes a hemispherical) nose and a cylindrical aft section. No attempt was 

made to further refine this aerodynamic shape. For Inflatable Aerodynamic Decelerators (IADs), the 

main distinctions are whether the IAD primarily provides lift or drag and whether it is used at hypersonic 

or supersonic (or both) speeds. For the thermal protection systems, there is an important distinction 

between TPS for flexible “aeroshells” (for IADs) or for rigid aeroshells. This set of technologies as 

reviewed with technology program managers from ARMD, ESMD, and SMD in May 2009. The group 

did not identify additional technologies to be considered. 

During the initial screening process, methods of hard landing such as airbags were eliminated, so all 

of the architectures assumed powered subsonic phases as the only practical way to attain pinpoint 

landings next to pre-deployed assets, as required by DRA5. Engineering details such as leg design, 

crushable segments, and stroke length were not studied. A hazard detection and avoidance system (such 

as ALHAT) is considered necessary, but sensor requirements were not evaluated in the Year 1 activities. 

At the External Peer Review in September 2009, the EDL-SA team received feedback that a rigid, 

deployable system should have been considered. This option was not carried explicitly because it is in the 

class of large, blunt-body vehicles to which inflatable decelerators belongs, but the areal mass of 

deployables is judged to be greater than inflatables, so inflatables were studied as the bounding case. In 

the future, rigid deployables will appear in the EDL taxonomy and will likely be considered on the 

roadmaps developed jointly by the EDL-SA and the technology programs. Other systems studies and 

contracts within the Small Business Innovation Research (SBIR) program have investigated deployables. 

Towed ballutes were briefly considered but ultimately judged unwieldy for decelerating masses of 

this size. 

For the purposes of 3-DOF performance simulations, and to stay within the constraints of the study 

schedule and budget, simplifications were made for modeling the various technologies. The important 

parameters of aerodynamic decelerators at this level are frontal area, aerodynamic characteristics (L/D), 

and mass. For instance, an LHIAD was assumed to have enough internal pressure to act as a rigid blunt 

body, and the mass model reflected the gas and gas generation system necessary to prevent buckling. 

(The validity of this assumption was investigated in a separate, parallel effort.) All propulsion stages, 
4

whether in the hypersonic, supersonic or subsonic regimes, used a common mass and sizing model. 

In keeping with standard practice in systems analysis for technology evaluation, the technologies 

were assessed against a suite of EDL-SA Architectures, i.e., a collection of representative architectures 

(high-level designs) against which the benefits of specific technology areas can be evaluated. The set of 

EDL-SA Architectures only needs to include options that encompass all candidate technology areas. The 

architecture suite is illustrated in Figure 2 and the resulting simplified set of technologies is listed in Table 

1. Evaluation of the technologies is accomplished by evaluating metrics at the architecture level, and then 

extracting the benefits (or penalties) of the technologies pairwise by comparison of architectures that 

differ only in the specific technologies. 

Figure 2. Exploration Class Architectures 

Table 1. Simplified Set of Exploration Class Technologies Considered by EDL-SA

Aerocapture Hypersonic Supersonic Subsonic

Architecture 1 Rigid Mid-L/D AS Rigid Mid-L/D AS Propulsion Propulsion

Architecture 2 Lifting HIAD Lifting HIAD Propulsion Propulsion

Architecture 3 N/A Propulsion Propulsion Propulsion

Architecture 4 Rigid Mid-L/D AS Lifting HIAD Propulsion Propulsion

Architecture 5 Rigid Mid-L/D AS Lifting HIAD Same LHIAD Propulsion

Architecture 6 Lifting HIAD Lifting HIAD Same LHIAD Propulsion

Architecture 7 Rigid Mid-L/D AS Rigid Mid-L/D AS Drag SIAD Propulsion

Architecture 8 Lifting HIAD Lifting HIAD LSIAD–Skirt Propulsion
5

3 Exploration Class Component Model 

3.1 Mass Models 

There are two key requirements for the EDL-SA mass models: they shall be parametric and consistent 

across all architectures. Parametric models are mathematical representations that relate the component 

mass to vehicle dimensions and mission key environmental parameters such as maximum deceleration 

and total heat load. The model consistency is achieved by sharing similar components across all eight 

architectures. 

The EDL-SA architecture suite contains five unique components (see Figure 3): the rigid mid-L/D 

aeroshell, a lifting hypersonic inflatable decelerator (LHIAD), a drag supersonic inflatable decelerator 

(DSIAD), a lifting supersonic inflatable decelerator implemented with a skirt on an LHIAD (LSAID–

Skirt), and subsonic/supersonic retro-propulsion (SRP). The details of the mass modeling approach are 

being documented in a separate NASA report and an AIAA paper to be available in late 2010. The next 

four subsections provide an overview of parametric mass models for rigid mid-L/D aeroshell, HIAD, 

SIAD, and SRP. 

Figure 3. EDL-SA Major Mass Components 

3.1.1 Rigid Mid-L/D Aeroshell 

The rigid mid-L/D aeroshell is a modified version of the dual-use Ares V shroud used by the DRA5 

study. The aeroshell has a straight barrel section with a hemispherical nose cap. The nominal total length 

is 30 m and the nominal outside diameter is 10 m. (Recent packaging results indicate that a rigid 

aeroshell with either SRP or SIAD for supersonic deceleration can comfortably fit within the Ares V 
6

shroud; however, simulation results are not yet available for this option.) The mass model for rigid midL/D consists of six subcomponents: structure, acoustic blanket, separation mechanism, body flaps, 

avionics, and TPS. 

The Ares-V finite-element (FE) analysis process was 

used to generate the structural mass estimates. The work 

was performed by Daniel Pinero and Lloyd Eldred. Loft 

[4], an in-house computer program, was used to automate 

the FE model generation with appropriate launch, 

aerocapture, and entry load cases. NASTRAN and 

Hypersizer were used to analyze and determine optimal 

structural mass subject to material and buckling 

constraints that were developed for the Ares V project. 

The barrel section consists of eight longerons and six 

frames (divided into five design groups). Hemisphere 

nose section consists of 8 longerons formed into one 

design group. Payload is attached to the second and the 

fifth frames: see Figure 4. A 25% mass growth 

allowance was added to the optimal mass to account for 

minimum gage design and required fasteners and other 

structural components not included in the FE model to obtain a current best estimate of the mass. A 

response surface equation (RSE) for the structural mass estimate was developed based on FE mass 

estimates. The RSE includes the following independent variables: diameter, total length, arrival mass, 

maximum dynamic pressure, and maximum lateral and axial decelerations. Figure 5 shows structural 

mass variation for nominal cases, excluding system-level mass growth allowance and system-level 

margin. Acoustic constraints for Mars EDL-SA payload are presently unknown. Mars surface power 

system may include radioisotope systems (RPSs), which may have a considerable impact on the acoustic 

blanket design. Standard acoustic blankets are most effective at 400 Hz and above (e.g., Titan IV has a 

three inch blanket with a one kg/m2 areal density). The Cassini blanket design was driven by radioisotope 

thermoelectric generators (RTGs) environment, which was qualified for the Galileo and Ulysses missions. 

The blanket was designed for 200-250 Hz (six inch blanket with a 3.9 kg/m2 areal density). Ares V is 

currently (October 2009) using a heavier, one-inch thinner, blanket (five inch blanket with 6.28 kg/m2 

areal density), and this blanket is used for the rigid aeroshell model. It is recognized that, depending on 

the packaging schemes selected for the architectures utilizing IADs, the IAD material may serve a dual 

use as acoustic blanket. Possible packaging arrangements are provided in Section 5. However additional 

detailed analysis and testing are needed and so for this analysis, the acoustic blanket mass is book kept 

separately. The mass estimate will be adjusted when there is additional information and a better 

understanding of Mars EDL payload acoustic requirements and packaging arrangements. 

The mass for the body flaps is a point design mass that is added to the aeroshell mass. There are two 

flaps that are 2 m long by 13.1 m wide; assuming 150 degrees warp angle. The areal density is 16.7 

kg/m2

 for flaps and 9.8 kg/m2 for the TPS. The mass estimate for flaps includes additional mass for 

actuation, hydraulics, and AU consumables. However the body flaps were not required in the final 

analysis. The TPS is a dual-layer PICA on top of LI-900, and the mass model is function of reference 

area and total heat load for aerocapture and entry. The TPS mass includes an attachment mass, which is 

44% of the TPS mass and is shown in Table 10. Figure 5 shows TPS mass contours, excluding attachment 

mass, the system level mass growth allowance, and system margin. Table 2 shows nominal simulation 

parameters and mass breakdown for rigid mid-L/D aeroshell for Architecture 1. 

Figure 4. Finite-Element Model of the Rigid 

Aeroshell
7

Figure 5. Structural (left) and TPS (right) Mass for the Rigid Aeroshell of Architecture 1 

Table 2. Nominal Parameters and Mass Breakdown for Architecture 1 

Variable Value Mass Components kg

Diameter, m 10 Structure 5482

Length, m 30 Acoustic Blanket 6415

Aerocapture Heat Load, MJ/m2 345 Separation System 2065

Entry Heat Load, MJ/m2 130 Avionics 222

Max Dynamic Pressure, kPa 11 Flap 1729

Max Lateral Deceleration, m/s2 29 TPS 9199

Max Axial Deceleration, m/s2 4 Total 25112

Arrival Mass, mT 110

3.1.2 Hypersonic Inflatable Aerodynamic Decelerator (HIAD) 

The HIAD design is based on Mars Inflatable Aeroshell Entry System (MIAS) model [5] which is a 

60o

 sphere-cone aeroshell. The model consists of an inflatable structure, flexible TPS, avionics, 

separation system, payload adapter and a rigid payload containment structure known as a heatshield. The 

inflatable mass model is based on the models developed by NASA in the 1960’s and 1970’s, and the 

model incorporates a double stacked-toroid consisting of radial straps to tie toroids together and carry 

radial loads, gores to carry circumference pressure loads, axial straps to carry the buckling loads, torus 

reinforce fabric to counter the hoop stress, a gas barrier, inflation gas, and gas generators. The straps and 

reinforcing fabrics are made of Kevlar-49, and the gores and gas barrier are made of Upilex. The 

mechanical properties of fabrics are reduced for operations in an elevated thermal environment. The 

design factors of safety for the HIAD follow the NASA standard for soft goods [6]. 

The toroidal structural concept is based on Brown’s design [7] that uses a minimum-weight fiberreinforced film. The design uses widely spaced reinforcing fibers bonded to the surface of the film, as 

shown in Figure 6. Brown concludes that the 12X advantage in specific strength of fiber compared to 
8

film results in a 7X lower mass, compared to the same size torus fabricated with unreinforced film for the 

same burst pressure. It is assumed that the fabric bondline temperature is 200o

C with material knockdown 

factor of 0.5. The material knockdown factor needs further testing for better understanding. The load 

factor of safety is set to 4 per NASA requirements for soft goods. Figure 7 shows the HIAD inflatable 

mass contours for various diameter and maximum 

dynamic pressures based on a 3.6 m heatshield 

diameter. The inflatable mass includes radial straps, 

gores, tori, inflation gas, and inflation system with 

appropriate knockdown factors due to an elevated 

temperature environment and NASA factors of 

safety. The solid gas generator is used to produce 

the inflation gas. 

Brown [8] recommends using 7.55% of launch 

mass for the adapter. The payload adapter for this 

study is set to 2% of arrival/entry mass because the 

adapter is assumed to carry small mechanical load 

during launch and it is primarily used during the 

aerocapture and entry phases. 

The flexible TPS is silica felt/silicone, and the 

parametric model is a function of reference area, the 

aerocapture heat load, and entry heat load. The current flexible TPS mass model is for an ablator that is 

limited to diameters less than 50 m. The TPS areal density for aeroshell diameters greater than 50 m is 

held fixed at the areal density of a 50 m aeroshell. The TPS model is suitable for high to moderate heat 

rates and loads. The use of this TPS model for architecture 6 could produce less accurate results. The next 

generation of mass model will include a mass model for a flexible insulator that will be suitable for lower 

heat rates and heat loads. Figure 7 also shows TPS mass contours for Architecture 2 as a function of heat 

loads. Table 3 shows the nominal parameters and mass breakdown for Architecture 2. 

Figure 7. Structural (left) and TPS (right) Mass for the HIAD of Architecture 2 

Figure 6. Reinforcing Fibers 
9

Table 3. Nominal Parameters and Mass Breakdown for Architecture 2 

Variable Value Components Mass, 

mT

% Areal Density, 

kg/km3

Diameter, m 23.0 Adapter* 2.2 21.2 5.3

Heatshield diameter, m 9.0 Heatshield 1.1 10.2 2.6

Aerocapture Heat Load, MJ/m2 87.3 Inflatable 1.8 16.8 4.2

Entry Heat Load, MJ/m2 26.1 Avionics 0.1 0.9 0.2

Max Dynamic Pressure, Pa 4240.1 Separation 1.3 12.4 3.1

Payload Mass, mT 40.0 TPS 4.0 38.5 9.7

Arrival Mass, mT 83.6 Total 10.5 100.0 25.1

HIAD Mass, mT 10.5 * [8] 0.0755*support mass +50

3.1.3 Supersonic Inflatable Aerodynamic Decelerator (SIAD) 

The SIAD, deployed after peak heating, is a hypercone model with no TPS and no knockdown factors 

for fabric due to high temperature. The SIAD model does include NASA recommended factors of safety 

for loads. A modeling approach similar to HIAD was used to design the SIAD components. 

3.1.4 Supersonic/Subsonic Retro-Propulsion (SRP) 

Architectures 1, 2, and 4 use supersonic RP modules and Architectures 5 through 8 use subsonic RP. 

Architecture 3 uses RP for the entire EDL segment. 

The Exploration Architecture Model for the IN-space and Earth-to-orbit (EXAMINE) [9] modeling 

tool, developed in-house at LaRC, was used to develop the parametric mass estimates of the SRP stage 

for all architectures. To eliminate the need for manual performance and sizing iterations, the SRP stage 

sizing and mass estimation was provided to the trajectory in the form of RSEs that enabled the SRP stage 

sizing and closure process to be completed simultaneously with trajectory optimization. 

Three RSE mass models were generated: one for aero-entry architectures that do not jettison discrete 

dry mass prior to entry (Architectures 1, 2, 6, 7, 8); one for aero-entry architectures that jettison a portion 

of the entry system dry mass prior to entry (Architectures 4, 5); and one for the all-propulsive architecture 

(Architecture 3). Table 4 shows independent variables as well as the upper and lower limits for the 

response surface equations. Table 5 shows the dependent variables. 

The primary SRP structure is an 8.8 m diameter aluminum-lithium (Al-Li) cylinder that supports the 

tank system and payload. This primary structure mass is estimated from a historically based empirical 

curve fit [10]. Thrust structure mass is based on a historical fit accounting for stage diameter, the number 

of engines and the thrust load. Secondary structure mass is 25% of the primary plus thrust structure 

masses. Landing gear mass is 2.5% of the landed mass on Mars. Multilayer insulation (MLI) is 5 cm 

thick (39.4 kg/m3

) covering the exterior structure, providing thermal control of the spacecraft. It is 

assumed that the payload provides power. The design includes a fluid-cooling loop that collects heat from 

the avionics cold-plates and cryogenic tankage (up to 10 kW), and heat is returned to payload thermal 

cooling system for heat rejection. The avionics model includes UHF, X-band, Ka-band communication 

systems, quad-fault tolerant flight computer, ranging and Doppler used for interplanetary position 

determination, and dual-fault tolerant laser radar (LADAR) altimeter for precision landing and hazard 

avoidance. 
10

Table 4. SRP Independent Variables and Limits for Response Surface Equations 

Architectures 1,2,4,5,6,7,8 3

Independent Variable Lower

Bound

Upper

Bound

Lower

Bound

Upper

Bound

Payload, mT 10 60 10 60

Terminal Descent ΔV, km/s 0.2 1.5 4 5.5

Initial T/W (Mars g’s) 3 11 1 4

Area Ratio 10 200 10 200

Aeroshell (Struc+TPS+misc), mT 5 55 NA NA

Aerocapture Apo-Correct. ΔV, m/s 0 150 NA NA

Descent Orbit insertion ΔV, m/s 0 500 NA NA

Percent pre-entry aeroshell Jettison, % 20* 90* NA NA

*Used for Architectures 4 and 5 

Table 5. SRP Dependent Variables 

 Dependent Variables

SRP Initial Mass, mT

Aeroshell initial Mass*, mT

Stack Mass at Arrival, mT

Stack Mass at Entry, mT

Stack Mass at Terminal Descent Initiation

Stack Mass at Landing, mT

SRP Propellant Mass, mT

SRP RCS Propellant Mass, mT

Aeroshell RCS propellant mass*, mT

RP Thrust per Engine

Engine T/W (Mars g’s)

*Not Applicable for Architecture 3

Liquid oxygen (LOX) and liquid methane (LCH4) propellants are used for both the main propulsion 

system (MPS) and the reaction control system (RCS). The MPS has four pump-fed expander engines 

each operating at 650 psia chamber pressure and a mixture ratio of 3.5. Because stage thrust-to-weight 

(T/W) and engine area ratio were selected as independent variables, the required thrust varies from case to 

case and in the overall closure/optimization. Thus, a set of RSEs for the MPS were developed to quickly 

predict the engine characteristics (vacuum specific impulse, engine thrust-to-weight, engine length and 

exit diameter) as a function of required thrust and area ratio. Figure 8 shows the specific impulse (Ispv) 

and engine T/W data used in the performance and sizing analysis. 
11

Figure 8. SRP Specific Impulse vs. Engine Thrust-to-Weight 

For all architectures except Architecture 3, two Al-Li cylindrical LOX tanks and two Al-Li 

cylindrical LCH4 tanks are packaged within the primary structure with the maximum diameter of each 

MPS tank limited to 3 m. For the all-propulsive Architecture 3, the 2x2 tank packaging arrangement 

yielded tanks with extremely high length-to-diameters (L/D) due to the need to package the additional 

propellant required by the all-propulsive case. At high tank L/D the tank and stage structure mass grows 

fast and does not allow model convergence. Thus, to limit the maximum tank length-to-diameter, an 

inline tank arrangement was used for Architecture 3 with one forward LOX tank, one aft LCH4 tank, each 

limited to 8.8 meters in diameter. For all architectures, the MPS tanks stored propellant at 50 psia and 

utilized advanced cryogenic propellant management technology to minimize boiloff (50 layers of MLI 

plus single-stage cryocooling system) and provide autogeneous pressurization and control. 

The RCS has sixteen pressure-fed thrusters each producing 100 lbf. Each thruster operates at a 

chamber pressure of 125 psia, a mixture ratio of 3.0, and an area ratio of 40, delivering a vacuum specific 

impulse of 334.5 sec. The RCS propellants are stored at 250 psia in two spherical graphite-wrapped 

aluminum tanks, one for LOX and one for LCH4. To minimize boiloff, 30 layers of MLI plus a singlestage cryocooling system are employed while a 6000 psia gaseous helium system provides consumables 

for RCS tank pressurization. 

 For all architectures, 100 m/s ΔV is allocated for RCS operation during landing. For Architecture 3, 

an additional 100 m/s ΔV is allocated for RCS operation during entry. 

Ground rules of the study required the dry mass growth allowance be 15% of the basic dry mass and 

an additional 30% (of the basic mass) is carried as system level margin. Thus, a total of 45% dry mass 

reserve is included in the mass estimates. Table 6 shows the mass breakdown for Architectures 1 and 3. 
12

Table 6. SRP Mass Breakdown for Architectures 1 and 3 

Mass Item

Arch 1 -

Rigid MidL/D Aeroshell

Arch 1 - Retro 

Propulsion 

Stage

Arch 3 - Retro 

Propulsion 

Stage

Primary Body+Thrust Structure 0.0 2076.3 4353.2

Secondary Body Structure 0.0 519.1 1088.3

Aeroshell Structure, TPS, Misc Mass 25111.8 0.0 0.0

Multilayer Insulation 0.0 107.2 83.2

Space Engines & Installation 0.0 1845.9 2623.4

RCS Engines & Installation 202.5 153.7 153.7

MPS Fuel Tanks & Feed/Fill/Drain Sys. 0.0 471.6 1877.4

MPS Oxidizer Tanks & Feed/Fill/Drain 

Sys. 0.0 512.1 3150.3

RCS Fuel Tanks & Feed/Fill/Drain Sys. 129.9 74.0 267.1

RCS Oxidizer Tanks & Feed/Fill/Drain 

Sys. 134.4 81.8 310.9

Pressurization System 0.0 90.9 1244.9

Power Management & Distribution 0.0 366.1 366.1

Command, Control, and Data Handling 0.0 12.7 12.7

Guidance & Navigation 0.0 10.3 10.3

Communications 0.0 61.0 61.0

Vehicle Health Management 0.0 0.0 0.0

Cabling and Instrumentation 0.0 35.4 35.4

TCS Heat Acquisition 0.0 120.1 120.1

TCS Heat Transport 0.0 322.9 322.9

TCS Heat Rejection 0.0 325.0 325.0

Landing Legs 0.0 1317.7 1801.6

System Level Margin 7673.6 2551.2 5462.3

Mass Growth Allowance 3836.8 1275.6 2731.1

Dry Mass w/ Growth 37089.1 12330.6 26400.9

Pressurant 2.9 53.2 684.8

Unused Fuel 41.8 73.5 1113.6

Unused Oxidizer 68.9 252.0 3863.4

Inert Mass 37202.6 12709.2 32062.8

Usable OMS Fuel 0.0 3155.8 52258.4

Usable OMS Oxidizer 0.0 11045.3 182904.3

Usable RCS Fuel 2088.9 517.7 3422.1

Usable RCS Oxidizer 3446.7 1553.2 10266.3

Gross Mass 42738.2 28981.2 280913.9

Masses do not include 15% Project Mass Growth Allowance
13

3.2 Aerodynamic and Aerothermal Modeling 

Two classes of vehicle configurations were analyzed during the first year of the study: 1) a rigid, midL/D vehicle and 2) a low-L/D flexible vehicle often referred to as the Hypersonic Inflatable Aerodynamic 

Decelerator or HIAD concept. Figure 9 presents the vehicle configurations as analyzed for the first year 

effort. 

 

Figure 9. Mid L/D (left) and HIAD (right) Vehicle Configurations 

3.2.1 Tool Set 

A suite of engineering and high-fidelity aerodynamic and aerothermal analysis tools was used to 

develop the aerodynamic and aerothermal databases for the two vehicle configurations depicted in Figure 

9. The analysis tools including CBAERO [11], DPLR [12] [13], LAURA [14] [15], NEQAIR [16], and 

CART3D [17]. 

The CBAERO software package is an engineering level aero-thermodynamics tool for predicting the 

aerodynamic and aero-thermodynamic environments of general vehicle configurations. CBAERO makes 

use of an unstructured surface grid of triangles to define the outer mold line of the vehicle configuration. 

No volume mesh is required. Solution times are on the order of seconds on a typical desktop computer 

for providing integrated forces and moments as well as the full aerothermal environments over the entire 

vehicle surface. For the present work CBAERO is used as the basis of the engineering-level aerodynamic 

and aerothermal analyses upon which corrections, based on high-fidelity CFD, are applied. 

The two high-fidelity CFD codes used in the development of the aerodynamic and aerothermal 

databases included: The Langley Aerothermodynamic Upwind Relaxation Algorithm (LAURA) code 

and the NASA Ames' Data Parallel Line Relaxation (DPLR) code. These codes solve finite-volume 

formulations of the Reynolds-averaged Navier-Stokes equations, including chemistry affects. 

The prediction of shock radiation was performed with NEQAIR. NEQAIR is a line-by-line 

spectroscopy and one-dimensional radiation transport tool. The code computes radiation transport (from 

the free stream to the vehicle surface) along straight lines of sight using a tangent-slab approximation. 

The data at points on a line of sight are obtained through interpolation from a volume solution computed 

using DPLR or LAURA, i.e., radiative heating is obtained assuming the flow and radiative processes are 

uncoupled. NEQAIR constructs accurate spectra at every point on a line of sight, computes the integrated 

radiative intensity from the spectral distribution, and transports the energy to the next point on the line of 

sight through one-dimensional transport. 

Finally, CART3D is a high-fidelity inviscid analysis package for conceptual and preliminary 

aerodynamic design, allowing users to quickly perform automated CFD analysis on complex geometries. 

The package includes utilities for geometry import, surface modeling and intersection, mesh generation, 

flow simulation and post-processing of results. The CART3D packaged was used to provide rapid 
14

turnaround of aerodynamic solutions in the subsonic and low supersonic regimes. 

3.2.2 Mid-L/D Geometry 

The mid-L/D rigid vehicle concept was based on a nominal 10x30 meter reference geometry used 

during the Mars DRA5 study. The original DRA5 configuration, based on a simple cylinder and 

spherical nose cap geometry had no provisions for body flaps or speed brakes. The present concept 

slightly modified the aft body section to allow for the 

incorporation of both a body flap and speed brakes. 

The nominal hypersonic L/D is 0.5 at angle of attack of 

55o

. The configuration was constrained to fit within 

the Ares V shroud, as specified defined by the EDL-SA 

ground rules. 

The aerodynamic model covers Mach 1.3 through 

50, angles of attack of 0 through 90o

, and dynamic 

pressures of 1.E-7 through 0.75 bars. The aerodynamic 

models covers body flap deflections in the range of -10 

to 50o

, and the speed brake for the range of 0 to 60o. 

The aerodynamic model was developed using three 

separate aerodynamic models generated with the DPLR, 

CART3D, and CBAERO software packages. Over 600 

high-fidelity CART3D solutions were run on the 

baseline, as well as a control-surface-deflected 

configuration. DPLR was run at a single (Mach 33) 

flight condition to anchor the high Mach number 

aerodynamics. Representative solutions from DPLR, 

CART3D, and CBAERO are shown in Figure 10, 

Figure 11 and Figure 12. Note that while the CART3D 

and CBAERO configurations included the full details 

of the body flap and speed brake, the DPLR geometry 

was simplified to remove the gaps between the control 

surfaces. 

The aerodynamic database is a merger of the 

aerodynamic solutions from DPLR, CART3D, and 

CBAERO. The DPLR solutions were used to adjust 

the CBAERO solutions at high Mach number using a 

simple delta correction form. The CART3D solutions 

were used exclusively below Mach 5, with the 

exception that the CBAERO viscous forces and 

moments were used to adjust the inviscid CARTD 

solutions. For the Mach numbers from 5 to 10 the 

CART3D and CBAERO solutions were linearly 

weighted such that the solution ramps from purely 

CART3D solution at Mach 5 to the purely CBAERO 

solution (anchored against DPLR) at Mach 10. 

The detailed aerothermal model is based on the 

previous DRA5 analyses and covers Mach 1.3 through 

50, dynamic pressures of 1.E-7 through 0.75 bars, at a 

Figure 10. Representative DPLR Solution

Figure 12. Representative CART3D Solution

Figure 11. Representative CBAERO Solution
15

single angle of attack of 55o. The 55o angle of attack corresponds to the nominal L/D of 0.5. The 

database was created using CBAERO and is anchored against a single DPLR/NEQAIR solution set at 

Mach 33, corresponding to the DRA5 peak total heating point. The aerothermal database provides the 

aerothermal environments over the entire vehicle surface. The flow was assumed to be fully turbulent, 

fully catalytic, with a radiative equilibrium wall boundary condition. The surface emissivity was assumed 

to be 0.85, although this value is subsequently adjusted in the TPS analysis based on the material 

selections. 

3.2.3 Flexible Low-L/D Vehicle (HIAD) 

Initial studies for the low-L/D flexible vehicle concept pointed led to a design L/D of 0.3 and a 

baseline diameter of 23 meters. An initial configuration based on a 60o

 and 70o sphere-cone 

configurations were chosen based on similar studies, specifically the Russian MIAS concepts, and the 

heritage of the Viking 70o

 sphere-cone geometry. However, the poor aerothermal behavior of these 

sphere-cone shapes, specifically the augmented turbulent heating due to entropy swallowing affects, led 

the design towards a constant radius heat shield configuration based on the Apollo geometry. The final 

configuration is shown in Figure 9 has a nominal L/D of 0.3 at approximately 20o

 angle of attack. Note, 

for this study the payload and 'backshell' configurations were undefined, and were assumed to have small 

impacts on the supersonic and hypersonic aerodynamics. For subsonic Mach numbers a simplified 

payload configuration was assumed. 

In the supersonic and hypersonic ranges, the similarity of the HIAD configuration to the Crew 

Exploration Vehicle (CEV) heat shield permitted the leveraging of the existing CEV aerothermal 

database. The CEV aerothermal database is a combination of high-fidelity CFD and engineering methods. 

The high-fidelity CFD codes, DPLR and LAURA, are used for the prediction of convective heating, and 

NEQAIR for the prediction of shock radiation heating. The high-fidelity CFD codes are run sparingly at 

critical conditions that bound the expected flight envelope. The engineering-level analysis code, 

CBAERO, is then anchored against the high-fidelity CFD and used to create a dense aerothermal database 

for use in TPS selection and sizing. 

For the present flexible configuration, the existing CEV aerothermal database was leveraged by using 

CBAERO to adjust for both scale (5 m to 23 m) and planet (Earth to Mars). The resultant aerothermal 

database provided the aerothermal environment over the entire vehicle surface. A representative, 

anchored, CBAERO solution for the flexible configuration is shown in Figure 13. Integration over the 

surface of the predicted pressures and shears provided the integrated forces and moments used in the 

aerodynamic database for all supersonic and hypersonic Mach numbers. For subsonic and transonic Mach 

numbers the CART3D package was used to calculate the forces and moments on the 23-meter HIAD with 

a representative payload container attached. Estimates for the viscous drag at these lower Mach numbers 

were obtained using CBAERO. A representative subsonic CART3D solution for the flexible 

configuration is also shown in Figure 13. The final aerodynamic and aerothermal model covers Mach 0.3 

through 50, angles of attack of 0 through 33o

, and dynamic pressures of 1.E-7 through 0.75 bars. 
16

Figure 13. Anchored CBAERO Solution (left) and Subsonic CART3D Solution (right) 

3.3 Aerocapture/EDL Guidance Algorithms 

Several guidance algorithms were developed and used to design both the aerocapture and EDL 

portions of flight. The guidance algorithms considered, assessed the feasibility of each architecture to 

aerocapture with acceptable margin and to maneuver during entry such that the vehicle could reach a 

specified target. Algorithms considered included the Hybrid Predictor-Corrector Aerocapture Scheme, 

(HYPAS), a Numerical Predictor Corrector Guidance (NPC), a theoretical “all knowing” guidance and an 

Apollo derived Analytical Predictor Corrector Guidance. For entry analysis, the theoretical guidance was 

used because of its ease of implementation and relatively quick run time. However to validate the EDL 

results, entry performance was compared to select simulations using both the NPC and APC guidance. 

3.3.1 HYPAS (Hybrid Predictor-Corrector Aerocapture Scheme) 

The HYPAS aerocapture guidance algorithm [18] [19], developed at JSC, guides a lifting vehicle 

through the atmosphere to a desired exit orbit apoapsis and inclination using bank angle control. The 

guidance uses an analytically derived control algorithm based on deceleration due to drag and altitude rate 

error feedback. Inputs to the guidance algorithm are the current vehicle position, velocity, sensed 

acceleration, and body attitude. The algorithm outputs a commanded bank angle and the direction to bank 

from the current attitude. The guidance algorithm is adaptable to a wide range of initial state vectors, 

vehicle lift-to-drag ratios and ballistic coefficients, planetary atmospheres, and target apoapsis and 

inclinations by only changing a set of initialization constants. Furthermore, by tuning these constants, 

other trajectory constraints can be controlled such as maximum dynamic pressure, deceleration, heat rate, 

and the amount of the theoretical corridor captured. 

A significant feature of the HYPAS algorithm is that no reference trajectories are computed prior to 

flight; all references are computed and updated during flight. This analytic, “on-the-fly” approach leads 

to efficient code, minimal data storage requirements, and minimal preflight effort. The non-numerical, 

non-iterative scheme ensures fast and consistent execution times. 

The original version of the HYPAS algorithm was developed for the Aeroassist Flight Experiment 

(AFE) program. During the AFE program, the algorithm was tested, compared, and evaluated against 
17

other guidance algorithms in three and six degree-of-freedom computer-based simulations. The HYPAS 

guidance algorithm was selected for the space flight test, and development of the flight code was on 

schedule until the AFE program was cancelled. Since the AFE program, the HYPAS algorithm has been 

used in numerous human and robotic exploration mission studies. These studies involved developing 

nominal and dispersed trajectory simulation results for aerocapture at Earth, Mars, Titan and Neptune for 

a wide range of vehicle L/D, ballistic coefficients, entry conditions, and target orbits. HYPAS was also in 

the evaluation phase for the NASA/CNES 2005 Mars Sample Return Mission [20] [21], which included 

an aerocapture orbiter at Mars. This work has provided the opportunity to gain a deep understanding of 

how the guidance algorithm performs in a variety of situations. Modifications have been made as 

necessary to improve performance and robustness. 

The HYPAS guidance algorithm consists of two phases. In the first phase, or capture phase, bank 

angle commands are generated to stabilize the trajectory and drive the vehicle toward equilibrium glide 

conditions, where lift, gravity, and centripetal forces are balanced. When the vehicle decelerates to a 

specified velocity, the second phase, or exit phase, begins. In the exit phase, the velocity vector at 

atmospheric exit altitude is analytically predicted each guidance computation cycle. Bank angle 

commands are then generated so that the velocity achieved at exit altitude will produce an orbit with the 

target apoapsis. This two-phase approach allows separate tuning of initialization constants to maximize 

robustness during capture and maximize performance during exit. 

Target apoapsis is achieved by controlling the vertical component of the lift vector through bank 

angle commands. The guidance algorithm follows a reference altitude rate and drag profile, generating 

bank angle commands using the control equation 

where φcmd is the commanded bank angle, is the altitude rate, is the dynamic pressure, D is the 

deceleration due to drag, and φeq.gl.is the bank angle required for equilibrium glide. and Gd are gains 

selected to provide the desired natural frequency and damping ratio response to a second order differential 

equation in altitude. Their values can be adjusted to maximize performance or robustness. 

During the capture phase, when equilibrium glide is targeted, the reference altitude rate, is zero, 

and the reference drag is computed each cycle by 

where VI is current inertial velocity, R is current radius vector magnitude, g is acceleration due to gravity, 

CL and CD are lift and drag coefficients, and K is a factor to determine how much of the lift vector should 

be used to maintain equilibrium glide. The guidance will attempt to insure the vehicle will not skip out 

by balancing the vertical forces. Selection of the gains controls the point in the entry in which equilibrium 

glide is established, thereby controlling heat rate and g-loads. 

During the exit phase of the guidance, a constant reference altitude rate is computed so that the 

velocity at atmospheric exit will provide the desired target apoapsis altitude. This reference altitude rate 

is updated each guidance cycle based on the difference between the predicted and desired exit velocity. 

The equation for updating the reference altitude rate is 
18

where Vmiss is the difference between the predicted velocity at exit and the velocity required at exit to 

achieve the apoapsis target with given reference altitude rate. The predicted exit velocity is computed 

from an analytic equation that assumes an exponential atmosphere and constant radial velocity to 

atmosphere exit for estimating velocity loss due to drag. The desired exit velocity is simply computed 

from orbital mechanics equations. 

The guidance algorithm requires an estimate of the current atmospheric density. An estimate of the 

density is derived from the measured deceleration due to drag assuming a nominal ballistic coefficient 

using 

where Vr is the current relative velocity, and Dmes is the sensed deceleration due to drag. The guidance 

algorithm models the atmosphere density as a simple exponential law of altitude 

where r0, and h0, are the reference density and altitude, and hs is the scale height. A density scale 

multiplier Kr is defined as the ratio of the density expected from the exponential model and the measured 

density. A first order low-pass filter of the form 

is used to smooth high-frequency atmosphere disturbances and control the guidance response rate to 

density changes. Then, the estimate of the atmosphere density is computed by 

. 

Note that with this approach, the guidance algorithm automatically compensates for dispersions in 

both atmospheric density and vehicle drag coefficient. 

Bank reversals are performed periodically to achieve the target orbit inclination. The lateral logic in 

the guidance algorithm uses an inclination or wedge angle dead band that is a function of inertial velocity. 

Whenever the inclination or wedge angle error exceeds this dead band, a roll reversal is commanded. The 

direction to bank is selected through a series of tests that examine current velocity, angular distance to 

roll, and difference between reference and navigated altitude rate. 

3.3.2 Numerical Predictor Corrector Guidance (NPC) 

The Numerical Predictor Corrector algorithm [22] was originally developed as candidate guidance for 

both the aerocapture and entry of the Mars Surveyor Program 2001 mission and the aerocapture of CNES 

Mars 2005 Sample Return Orbiter. Orion is developing a NPC for its lunar return guidance. The NPC 

algorithm operates using an outer loop called every simulation time step and an inner loop, called at some 

specified interval (e.g. every 10 seconds). For the EDL-SA implementation, the NPC receives perfect 

navigated state conditions (position and velocity) as well as current roll angle and sensed body-axis 

accelerations. The algorithm integrates 3DOF translational equations of motion using a fourth-order 

Runge-Kutta integration scheme. 

The inner loop includes models of the planet, gravity field, vehicle aerodynamics, body attitude, mass 

properties and atmosphere. The inner loop of the algorithm produces a command vector that includes roll 
19

angle and reversal time. A state history, including range to target, energy and time rate of change in 

energy is also stored. The state information is used to modify the roll angle magnitude between updates. 

The alteration is done in the outer loop by comparing the actual state conditions with the predicted 

state conditions and then modifying the roll angle appropriately. A half-interval search method is used to 

determine the appropriate roll angle and reversal time. Sensed acceleration data is used to update the 

internally stored atmosphere density profiles and aerodynamics. This is done by calculating the ratios of 

normal and axial sensed acceleration and comparing them to the values predicted by the guidance 

algorithm. In addition, the guidance routines are used to calculate a local density scale height. By 

knowing both the aerodynamic accelerations ratios and the local density scale height, scalar multipliers to 

the guidance predicted aerodynamics and an altitude bias to the guidance atmosphere model can be 

calculated. These scalar multipliers and altitude bias quantities are averaged over a 10 second intervals 

and the averages are used within the inner loop. 

The predictor corrector inner loop is activated on a deceleration trigger that is tailored for each 

architecture. Once the predictor-corrector is invoked, the vehicle guidance strategy is a combination of 

roll angle magnitude and roll reversal. The general strategy is to iterate on the roll angle magnitudes until 

the predicted miss distance is within a preselected tolerance. Roll reversals are commanded if the heading 

error exceeds a preset limit. Roll angle magnitude and reversal times are then passed to the outer loop for 

execution. 

3.3.3 Theoretical EDL Guidance 

The “Theoretical EDL Guidance” was developed as highly adaptable to support all the architectures. 

It also had the ability to be detuned such the results would be comparable to those expected from flight 

guidance algorithms. It is POST2 based and has full knowledge of all environmental parameters such as 

aerodynamics, atmospheric density and wind profiles. The flight phases are representative of the 

numerical predictor-corrector and the Apollo-derived algorithms that were used in this study. The 

theoretical guidance used bank angle magnitude to control the range to the target and bank angle direction 

to control crossrange. After the range control phase the guidance switched to a heading alignment phase 

where the guidance used bank angle to remove any residual cross range error, and placed the vehicle at 

the proper velocity, altitude and range from the target to initiate the staging and propulsive descent phase. 

3.3.4 Analytic Predictor-Corrector – Apollo-Derived Guidance 

The Analytic Predictor-Corrector (APC) used in the EDL-SA is a combination of a modified Apollo 

entry guidance [23] and the Apollo powered descent guidance [24]. The modified Apollo entry guidance 

was a candidate guidance algorithm for both the aerocapture and entry of the Mars Surveyor Program 

2001 mission, and the aerocapture of CNES Mars 2005 Sample Return Orbiter. It is the guidance used for 

the MSL Entry, and will be the guidance for Orion return from the ISS. The modified Apollo powered 

descent guidance is currently under evaluation for Altair’s lunar propulsive descent guidance. 

Atmospheric Entry Phase 

Guidance Logic – In this section, an overview and the main characteristics of the Apollo-derived 

entry terminal point controller are presented. The original Apollo atmospheric entry guidance was 

designed for lunar return; however, sufficient mission flexibility was required to accommodate the large 

variations in actual entry conditions, including those of Earth orbit test flights and all types of lunar 

mission aborts. To satisfy re-designation requirements for a weather alternate landing area, an upper 

altitude controlled skip entry capability was included. However, the Mars entry application considered in 

this paper does not require a skip phase. Therefore, only the final entry phase will be used. 

The final entry phase algorithm controls to a terminal range and velocity target using pre-derived 

influence coefficients with respect to errors about a nominal L/D reference trajectory. This reference 
20

trajectory is defined by range-to-go, drag acceleration, and altitude rate as a function of relative velocity. 

The predicted range-to-go (Rp) is calculated as a function of drag (D) and altitude rate ( ) errors with 

respect to the nominal reference trajectory profile and the associated influence coefficients; see equation 

1. 

€

Rp = Rref +

∂R

∂D ( ) D − Dref − ∂R∂r  ̇ 

r  ̇ − r  ̇ ( ) ref 1

The commanded vertical component of the lift-to-drag ratio is calculated as the addition of the 

reference L/D plus a function of the difference between the actual and predicted range to go 

 2 

The commanded bank angle is then calculated as 

€

φc = Kbank cos−1 ( ) L /D c

L /D











 3 

where Kbank is the bank directional control (±1). The sign of Kbank is reversed each time the target 

crossrange out-of-plane central angle exceeds the reversal criterion, which is a quadratic function of the 

relative velocity. 

The partial derivatives of predicted range in equations 1 and 2 are the controller gains, which are 

derived using linear perturbation theory with the nominal reference trajectory by reverse integration of the 

differential equations adjoint to the linearized equations of motion. Optimized control gains for 

converging the dispersed trajectory are implemented in the guidance as tabular functions of relative 

velocity. Because of slow system and trajectory responses to guidance commands, performance is 

empirically enhanced by the use of the over-control gain K in equation 2 to improve range convergence 

behavior. 

To minimize disturbance of downrange control near the target, bank reversals are inhibited below a 

certain relative velocity magnitude. If the bank command is not in the same direction as the target at this 

point, then a final reversal is commanded prior to setting the inhibit flag. After that, a heading control 

phase is initiated in which the bank command is defined by 

where Khc is the overcontrol gain for the heading control phase and where Cr and Dr are the crossrange 

and downrange to the target, respectively. 

Methodology – The reference trajectories basically consist of two segments. The first one, covering 

velocities approximately higher than 2 km/s, usually requires large bank angles such that the constraints 

on heat load, heat rate, and deceleration loads are met. The second segment, covering velocities lower 

than 2 km/s, is designed such that the vertical lift of the entry vehicle be increased in order to further 

reduce the terminal relative velocity at a given terminal altitude. Vertical lift margin is provided in this 

segment in order to allocate dispersions. However, the reference trajectory for the final part of this last 

segment calls for a full lift up configuration below the velocity at which the heading control phase is set 
21

to start. 

Powered Descent Phase 

Guidance Logic – The guidance logic for this phase is a quadratic guidance that uses a pre-stored 

gravity turn steering thrust reference profile that reaches the landing site at the touchdown conditions. It 

commands thrust and attitude to reduce position and velocity errors at touchdown. 

In this logic, the acceleration profile (a) is restricted to a quadratic equation 

 4

The acceleration profile is integrated to develop the analytic equations for velocity and position 

 5

 6 

Equations 4, 5 and 6 can be rewritten as 

Where the target tgo, rt, vt, and at come from the reference profile. Solving for the coefficients c0, c1, 

and c2 in matrix form results in 

Hence, the generic acceleration command for each axis is given by 
22

Thus, the thrust command is given by 

At each guidance call, the coefficients are recalculated in each axis and t is set to zero. The initial 

state for the two-point value boundary value problem is the current vehicle state. The burn length to go, 

tgo, will decrease as the vehicle approaches the target state. As tgo becomes very small, the coefficient 

terms will tend toward infinity. There are typically two different methods employed to prevent this from 

happening. One method is to use a target state that forces the vehicle to follow the desired trajectory path 

without the intent of actually reaching that target. The guidance phase is then terminated prior to tgo 

approaching zero. The second method is to utilize a fine count, where the coefficients are not 

recalculated after tgo goes below a pre-selected value tfc (typically 1-2 sec). During the fine count, the 

coefficients are not recalculated, and t is not reset at each guidance call. 

The pointing direction in terms of pitch and yaw are calculated as follows 

Methodology – The reference profile is a gravity turn trajectory that consists of three segments: a 

constant thrust level, followed by a ramp down, and a smaller constant thrust level. Two target points are 

determined for the whole trajectory, one is set at the end of the ramp down segment for a few seconds 

before the end of the trajectory, and the other is set at the end of the trajectory, which ends at the 

predetermined final conditions. The fine count method is the one that has been used in the study, with a 

tfc = 2 sec. 

3.3.5 Guidance Comparison 

 Figure 14 shows the comparison of a Monte Carlo analysis using APC, NPC and theoretical 

guidance’s for Architecture 1. Both high fidelity guidance algorithms (APC & NPC) can provide the 

required targeting accuracy (10 m with perfect navigation) and have minimal impact on the design 

parameters (max g’s, max heat rate and load, staging conditions, etc). Results are similar for Architecture 

2 results not shown here. 

Figure 15 shows a similar comparison for Architecture 8. Unlike the Architecture 1 and 2 results, the 

current detuning level of theoretical guidance may not be sufficient as the theoretical guidance analysis 

uses significantly less propellant. Also note that the APC terminal descent guidance produces tighter 

propellant statistics than current NPC strategy. 
23

Figure 14. Guidance comparison for Architecture 1. 

Figure 15. Guidance comparison for Architecture 8. 
24

3.4 Retro-propulsion 

The primary objectives of the Supersonic RetroPropulsion (SRP) study element for EDL were to 

provide: (a) Isp and thrust/weight ratio guidance for use in higher-level vehicle parametric modeling and 

(b) high level system design insight to the team. A literature review, commercial modeling software, and 

system design experience were used to provide these products. Using the provided system requirements, 

an assessment of the state-of-the-art for liquid oxygen/liquid methane (LOX/LCH4) propulsion, engine 

throttle range considerations, a notional design including main engine thrust, weight, Isp performance, 

cycle options, engine configuration, tank sizing and forward work were provided. 

For the purposes of the study, DRA5 was the starting point, with the ranges identified in that report 

limited to a representative point design, as specified by the EDL study lead. Key requirements included a 

LOX/LCH4 methane propulsion system. This propellant combination is generally considered spacestorable and ISRU (in-situ resource utilization) compatible. Given an initial vehicle combined mass of 62 

mT and a 3 g (earth) initial deceleration, 1.8 MN (400,000 lbf) initial thrust is required. Throttle-able for 

landing, the final thrust is ~10% at shutdown. For sizing purposes, a delta V of 600 m/s was used. 

Similarly, for the ellipsled vehicle, with 5 deg/s^2 control rates, a smaller RCS thruster of 9kN (2000 lbf) 

for roll/pitch and a larger RCS thruster of 62kN (14,000 lbf) for yaw are required. 

The SOA for LOX/LCH4 is driven primarily by recent development efforts funded through EDTP. 

These efforts are primarily geared toward upgrade concepts for the Altair Lunar Ascent Vehicle, although 

the same technologies would also be useful for a non-toxic, high performance propulsion system for the 

ORION Service Module, and potentially a Lunar Lander Descent Stage. These efforts are all geared 

toward relatively low thrust and low delta-V missions, as appropriate for a Lunar Ascent vehicle 

application. Many of these efforts begin to ramp down in the 2011 timeframe. The total technology effort 

is ongoing at 6 NASA centers, a dozen contractors in 30 or 40 separate technology development 

programs. Overall, the TRL could be considered medium for RCS and small Main Engines, and low for 

large, throttle-able main engines. (Significant engine development efforts include a prototype 22kN/5klbf 

fixed thrust ascent engine, a 176kN/40klbf breadboard injector with heat sink chamber, a 44kN/10klbf 

workhorse engine and several 444N/100lbf RCS thrusters delivered). In addition to the EDTP work, 

numerous studies have been performed on LOX/LCH4 engines, from 44kN to 900kN. A Korean engine 

of 90kN is advertised as space qualified, although it has not been flown to date. 

 Engine throttle range considerations are primarily related to the lower ~50% limit typical of most 

large pump fed engines. Injector dynamics, pump and turbine design, engine cooling and performance 

drop-off at low throttle levels limit the practical throttle range for most engines. A recent demo of the 

CECE LOX/LH2 achieved 13%, which suggests potential for at least one well-characterized engine. A 

recent LOX/kerosene Russian engine achieved 30% throttle on the test stand. Pressure fed engines are 

more amenable to wide throttle range, but are impractical for the large thrust levels needed for the Mars 

Lander. 

Throttle considerations, along with a desire from the vehicle control team to have engines in multiples 

of three, led to a notional design of 6 engines for further study. Such a cluster could have 1⁄2 the engines 

shutdown at 40-50% throttle, with the balance throttled up to 100% to maintain the descent rate. Then 

these 3 engines would throttle down to ~20-30% for landing, a reasonable, if optimistic goal for throttle 

range with many years left to develop such an engine. With 1⁄2 the engines operating at low throttle, the 

desired 10% total thrust at shutdown is thus achieved. An example of the profile settings is provided on 

the left in Figure 16. Given the DRM sizing constraints, the notional propulsion system design would 

include 6 engines of 300kN thrust each. Using commercially available engine modeling software, a 

parametric analysis was performed and a final engine design was chosen that optimized for Isp. Not 

surprisingly, the high Isp solution of 374s is achieved at the expense of a very high expansion ratio of 
25

180, which drives a very tall nozzle. The figure in the right of Figure 16 illustrates a typical curve of 

expansion ratio for increasing Isp, albeit for a slightly different mixing ratio. The engine T/W ratio is 60, 

although several papers and historical precedent suggest that 70 is achievable with adequate development. 

Further improvements in T/W could be achieved at additional expense, although an optimized engine is 

well beyond the scope of this study. 

Figure 16. SRP Sample Throttle Profile (left) and SRP ISP vs Expantion Ratio (right) 

Similar to the propulsion system, a notional design for the propellant tanks is based on the DRA5 

sizing of a 62 mT vehicle (with 40 mT payload, 10 mT structure and 12 mT propellant) and the proposed 

mixture ratio of 3.5. Relative to most lander designs (e.g. all propulsive descent as in Apollo LM), the 

tanks are rather modest in size, as expected with a modest delta-V requirement and the relatively short 

burn time of less than 3 minutes. Assuming a common tank diameter of 1.5 m, 9330 kg of LOX would be 

stored in a pair of 8.2 m3

 tanks, 2.5 m tall. Similarly, 2670 kg of LCH4 would be stored in a pair of 6.1 

m^3 tanks, 2.2 m tall. A notional tank arrangement is shown in Figure 17. Specifics of the tank design are 

provided in Table 7. 

Figure 17. Notional Tank Arrangement 
26

Table 7. SRP Engine Characteristics 

Characteristic Metric English

Total Thrust at Initiation 1.8 MN 400,000lb

Engine Thrust 300 kN 67,000 lb

Chamber Pressure 6.2 MPa 900 psia

Engine Height 2.8 m 9.2 ft

Nozzle Height 2.3 m 7.6 ft

Nozzle Exit Diameter 2.5 m 8.2

Engine Mass 550 kg 1210 lbm

Forward work identified in this study may be grouped into engine, tank and vehicle-level 

considerations. The engine startup in the reverse flow field during Mars entry has two major components: 

engine startup and nozzle size/stiffness. The startup pressure in the chamber was calculated (by the 

aerothermal team) at 5.2 kPa (0.75 psia). This pressure is well within our experience base, since sea-level 

and vacuum starts are both common operations. A much more significant question is the interaction 

between a large nozzle and the unsteady flow stream around the vehicle at startup, and the dynamic 

response of the flow stream to mass flow introduced behind the bow shock, along with nozzle stiffness. 

The relative size of the nozzles is also much larger than has ever been tested in wind tunnels, as is the 

significant mass flow into the oncoming flow stream. Multiple nozzles interacting both with each other 

and the environment adds another layer of complexity. 

Nozzle issues aside, the development of a large, throttle-able LOX/LCH4 engine is also required, 

although some progress has been made on some elements of the engine at both NASA and various 

contractor facilities (most relevant is the Pratt Whitney 60klb LOX/LCH4 effort which has performed 

some study and turbo-pump testing of an upgraded variant of the well-known RL10 engine). Numerous 

other studies are available in the AIAA and JANNAF databases, although relatively little hardware has 

been built or tested. One promising concept for reducing engine height is an aerospike solution, although 

throttling would likely be a challenge with a single large engine. The aerospike concept is generally 

chosen for its inherently altitude-compensating nozzle-less external flow field, but here the elimination of 

a large nozzle would be a good fit for the Mars Lander mission. 

Propellant tankage forward work should include liner-less composite overwrapped tanks, which have 

been demonstrated at smaller scale, including a Garvey Aerospace sounding rocket. The elimination of 

the metal liner may reduce the tank mass up to 50%. The relatively small size of the tanks, relative to the 

vehicle size, suggests that an integrated tank/structural element solution is unlikely. Transit times with 

LOX/LCH4 should be compatible with the Mars mission timeline, since a Ball Aerospace study on a 

proposed ORION Service Module concept design was capable of 6 months lunar loiter with minimal boiloff losses. Large area MLI and MMOD may be a challenge. 

In summary, an Isp range of 365 to 375 seconds appears reasonable, with the higher numbers 

requiring more development. For conventional nozzles, the upper end of the Isp range will lead to large 

nozzles, awkward for the Mars Lander, suggesting the practical solution will either be at the lower end of 

performance or depend on an alternate concept such as an aerospike solution. Similarly a T/W range of 

60 to 70 should be realistic, with the higher end requiring more work. No large (300kN) throttleable 

engine technology is available, although very preliminary work has been performed for a fixed thrust 

280kN engine. Most of the current LOX/LCH4 propulsion development work begins to ramp down in 

2011, and the ongoing work is targeted at much smaller systems than the Mars Lander will require. 

Engine startup and interaction with the unsteady flowfield behind the vehicle bow shock will require 

significant CFD analysis, wind tunnel testing and subscale flight tests, e.g. sounding rocket experiments. 

Vehicle drag, flutter and stiffness will interact and should be considered a significant development area. 
27

Propellant tank size is rather modest due to the modest delta-V requirements, but tank, MLI and MMOD 

weight are significant contributors to vehicle mass. 

3.5 Thermal Protection Systems 

Estimated thermal response models were used in the sizing analysis of the thermal protection systems 

(TPS) needed for the HIADs and rigid mid-L/D aeroshells. For both aeroshells the sizing was performed 

on a dual heat pulse mission (aerocapture, cool down in orbit and out-of orbit entry). The sizing analysis 

is based on the tools and practices developed by the Orion TPS ADP. The TPS sizing tool (which makes 

use of the ablative thermal analysis tool FIAT [25]) was extended for EDL-SA to include the capability to 

size dual layer TPS. The margins applied to TPS masses are documented in [26] and summarized in Table 

8. The margins were developed based on the Orion TPS ADP experience. (Note that the rigid mid-L/D 

aeroshell analysis was run before the margin document was finalized so those analyses used the Orion 

values for initial temperature and recession uncertainty.) 

Table 8. TPS Sizing Margins 

Parameter Value

Initial Temperature 10oF (Rigid Aeroshell 70oF )

Radiation Sink Temperature 70oF

Allowable Bondline Temperature 550oF

Blowing Factor 0.4

Ablator Fail Lien 50%

Thermal Margin 108oF

Gap Heating Factor 1.0

Recession Uncertainty, Default FOS 1.3, 1.1 (Rigid Aeroshell 1.2)

3.5.1 HIAD – 23 m diameter 

TPS Material Properties and Stack – Currently no flexible ablative material exists that meets the 

estimated HIAD requirements. For this study, estimates for possible properties of a flexible 

Silica/Silicone ablator were created and used in the TPS sizer. The properties were based on SIRCA-15 

(15 lb/ft3

) with some modifications. The virgin density reduced to 10 lb/ft3 and the char density was 

assumed to be 8 lb/ft3. The thermal conductivity was taken to be 80% of the SIRCA-15 conductivity. 

The Silicone decomposition and virgin and char emissivities were kept the same as SIRCA-15. The 

material stack up used for HIAD is shown in Figure 18. For these analyses a conservative assumption of 

no structure or thermal mass was used. 

Figure 18. HIAD Material Stack 
28

Aerothermal Environment – The HIAD dual pulse environment for aerocapture and entry is shown 

in Figure 19. The plot on the left shows maximum total heating for the nominal aerocapture in red and 

the 80 metric ton entry in blue. (The curves are plotted one after the other for visualization, but in reality 

there is a long on-orbit cooling period between them.) The integrated heatload contour is shown on the 

right. Fully margined analysis estimates that the 23 m HIAD experiences around 100 W/cm2

 during 

aerocapture and 30 W/cm2

 during entry. 

 Figure 19. HIAD Dual Pulse Surface Heating History and Integrated Heatload Contour (Fully Margined) 

TPS Analysis Results – The computed mass for a bond line temperature of 5500F was predicted to be 

3.2 mt, including the RTV (Room Temperature Vulcanized bonding agent) attachment mass and not 

including WGA. These mass calculations are employed in the EDL-SA HIAD mass estimating 

relationships (MERs). Some bondline temperature sensitivity studies were run which showed a 12% 

increase in the TPS mass if the allowable bondline temperature were reduced to 2500

C (4820F). 

3.5.2 Mid L/D Aeroshell 

Background – For DRA 5, TPS masses for a forebody heat shield made of a single material (PICA) 

were very large. The DRA 5 TPS subteam conceived a dual-layer TPS with the PICA ablator atop a low 

thermal conductivity substrate such as the Shuttle tile material or SLA 561. Individually, these TPS 

materials are at a high TRL, but a dual layer TPS system made of these materials has never been 

evaluated or tested. Using existing thermal response models for PICA and LI-900, preliminary hand 

analysis of the dual layer concept at three body points indicated that the concept could reduce TPS masses 

by about 27%, not accounting for attachments. 

This updated analysis used the new sizing code to do entire body dual layer TPS sizing. The concept 

for PICA atop the LI-900 Shuttle tile material is shown to be capable of reducing the windward TPS mass 

by 37%, not accounting for TPS attachments or weight-growth allowance. 

TPS Material Stack and Analysis Methodology – The TPS material stack used for the mid L/D 

aeroshell analysis is shown in Figure 20. The tile was attached to the substructure with a standard 

RTV/SIP configuration. The analysis substructure was a titanium plate and represents the thermal mass 

of typical aeroshell thermal masses. The ablator was modeled directly above the tile. No adhesive or 

attachment between the ablator and tile was modeled. This was considered to be sufficiently conservative 

for this analysis cycle. 
29

Figure 20. Mid L/D Aeroshell TPS Stack 

The material splitlines for the mid L/D vehicle are determined by the maximum temperature seen at 

each location during aerocapture and entry. They are shown in Figure 21. The leeward sections are tile 

only and were sized by a transient run that included both aerocapture and entry with a long on-orbit cool 

off between. For the windward dual layer sections, it is assumed that the on-orbit cool off between the 

aerocapture and entry phases is sufficiently long enough for the TPS system to cool completely to the 

initial temperature. It is also assumed that the ablator is needed to protect the tile during aerocapture, but 

is not needed during entry. So for the windward dual layer TPS analyses, the aerocapture and entry 

phases was separated and run in two steps. For the first step, the ablator was removed from the model and 

the tile layer was sized to survive the entry environment while maintaining the allowable bondline 

temperature at the tile/RTV interface. For the second step, the ablator was reattached to the model. The 

ablator was then sized to maintain the tile allowable surface temperature at the ablator/tile interface (as 

well as the allowable bondline temperature at the tile/RTV interface) during aerocapture. 

Note, that the above described analysis method assumes that ablator is needed to protect the tiles in a 

dual-layer stack during the aerocapture phase only. However, the validation testing of the dual heat pulse, 

dual layer concept presented in Section 7.1 below shows some slumping of the LI-900 at the peak entry 

heating rates once the PICA has ablated away. Therefore, future analysis should consider leaving a thin 

amount of ablator (or char) at the end of aerocapture to protect the tile during entry at the highest heat flux 

areas of the windward dual-layer section. 

Figure 21. Mid L/D Aeroshell Material Splitlines 

Aerothermal Environment – The mid L/D aeroshell dual pulse environment for aerocapture and 

entry is shown in Figure 22. The plot on the left shows maximum total heating for the aerocapture in red 

and the entry in blue. (The curves are plotted one after the other for visualization, but in reality there is a 

long on-orbit cooling period between them.) The integrated heatload contour is shown on the right. Fully 

Ablator/Tile:Ablator sized to maintain 

tile allowable temperature for aerocapture

Tile sized to maintain bondline allowable 

temperature for entry

LI900/LI2200 Tiles (leeward & base)

Sized to maintain allowable bondline 

temperature for both aerocapture and 

entry

Sizing Points

Material Splitlines 10x29
30

margined analysis estimates that the mid L/D aeroshell sees around 450 W/cm2 during aerocapture and 

120 W/cm2

 during entry. 

Figure 22. Mid L/D Aeroshell Dual Pulse Surface Heating History and Integrated Heatload Contour (Fully 

Margined) 

 TPS Analysis Results Not Including Attachments – Results for the mid L/D aeroshell TPS sizing 

analysis are shown in Table 9. Analyses were run for the ablator PICA over two different tiles, LI-2200 

and LI-900. For comparison, a case was also run with PICA only on the windward surface of the vehicle. 

The results show that if the lightweight LI900 tile is used with PICA, there is an estimated TPS mass 

savings over the whole vehicle of 34%. These estimates do not include masses for the attachment of the 

ablator to the tile or for any structure to hold the TPS. 

Table 9. Mid L/D Aeroshell TPS Mass Estimates (Not including attachment or WGA) 

Windward TPS PICA Only Dual Layer PICA/LI-2200 Dual Layer PICA/LI-900

Mass, kg Mass, kg Savings Mass, kg Savings

Nose Windward 1670 1350 19% 1000 40%

Body Windward 8750 7020 20% 5580 36%

TOTAL Vehicle TPS* 11260 9210 18% 7420 34%

*Includes Leeward Tiles

Estimates for Attachment Weights – Estimates for TPS system attachment weights were performed 

and are shown in Table 10. Two attachment methods were considered, a shuttle like tile system with 25.4 

cm square tiles and a honeycomb system with 5 cm blocks bonded into a honeycomb, which was bonded 

to the substructure. The 5 cm honeycomb system estimates were made by scaling measured masses from 

Applied Research Associates, ARA [27] for 5 cm thick PICA blocks in phenolic impregnated silica fabric 

honeycomb with RTV on 5 sides of the PICA blocks. (Note, the “TPS Mass” column differs slightly 

from the total values in Table 9 because mass was removed for the gaps and the volume calculations were 

updated to account for the curvature of the vehicle). The shuttle like tile systems are considered to be 

unrealistic for manufacturing of the dual layer system, so method 2 was discounted. However, method 3 

(5 cm blocks in honeycomb) using PICA over LI900 tiles still shows a 24% mass savings over PICA 

alone (method 1). The mass calculations for method 3 are employed in the EDL-SA mid L/D mass 

estimating relationships (MERs). 
31

Table 10. Mid L/D Aeroshell, Windward Side Estimated Masses with Attachment 

TPS Attachment Method TPS Mass, 

kg

Attachment 

Mass, kg

Total Mass, 

kg Comments

1 PICA Only

Shuttle like Tile System

25.4 cm tiles

0.07 RTV gap filler

RTV/SIP/RTV to structure

10,600 1,100 11,700 Note*

2 PICA/LI900

Shuttle like tiles

25.4 cm tiles

0.07 RTV gap filler

RTV/SIP/RTV to structure

6,700 900 7,600

Un-bonded PICA on LI900

Unrealistic from 

Manufacturing 

Consideration

3 PICA/LI900

5 cm HC System

HT424 HC to structure

RTV blocks to HC

RTV blocks to structure

6,500 2,900 9,400

Feasibility demonstrated 

at coupon level. 

23% heavier than method 

2, but

24 % lighter than PICA 

only

* Orion considered the PICA (#1) option for use with a titanium honeycomb carrier structure system but 

chose an Avcoat option instead based on overall system level risk and reliability. MSL used a system of 

PICA tiles direct-bonded to a composite substructure with RTV filled gaps. PICA tiled system selection, 

while accepted for MSL and not for CEV Orion, is based on system level integration issues and reliability 

requirements. Any tiled system for human mission design will have to overcome concerns similar to the 

ones considered by Orion.

Conclusion – The TPS sizing results presented here, along with the validation testing of the dual heat 

pulse, dual layer concept presented in section 7.1 below, demonstrate the credibility of providing the TPS 

for the 10x29 meter Mid L/D aeroshell (ellipsled) for future human Mars missions. The use of a 5 cm 

honeycomb for the windward TPS installation appears to enable the manufacture of the large heat shield. 

The thermal performance of the PICA atop LI-900 TPS installed with the honeycomb is shown to be 

predictable for the dual heat pulse environment. The dual layer TPS concept, including attachments, 

provides significant mass savings compared to a PICA-only solution. Three caveats are made here: (1) 

As shown in work by the Orion TPS ADP, PICA recession at heat rates below 50 W/cm2

 is not well 

predicted by FIAT, owing to out-of equilibrium gas-solid interactions within the material at those 

conditions. As with Orion, the low heating conditions will be experienced by the mid L/D aeroshell at the 

“tails” of the two flight heat pulses. (2) Four-point positive and negative flexural tests of PICA atop LI900 in ~ 5 cm honeycomb showed encouraging results for test articles fabricated on an aluminum 

substrate. However, further development on the bonding of the honeycomb to realistic composite 

substrates has not been performed to date. (3) The “fencing” caused by non-uniform ablation of the 

honeycomb and PICA at low heat fluxes may be an issue and requires testing and refinement of the 

phenolic impregnated fiberglass honeycomb technology. 
32

4 Simulation Results 

The simulation used to evaluate the EDL-SA architectures is the Program to Optimize Simulated 

Trajectories (POST2) [28], which is a generalized point mass, discrete parameter targeting and 

optimization simulation. POST2 has the capability to target and optimize multiple vehicles near a 

rotating oblate planet and has extensive heritage for simulating ascent, descent, and orbiting trajectories 

including past missions like Shuttle, Genesis, Stardust, Genesis, and Mars Missions including, Pathfinder, 

Odyssey Orbiter, Exploration Rovers, Reconnaissance Orbiter, and the Phoenix Lander. POST2 is also 

used extensively in current missions like Mars Science Laboratory, Launch Abort Systems and Ares. 

Specific models used to tailor production POST2 for the EDL-SA purposes include use of the Mars 

Global Reference Atmosphere Model 2005 (MarsGRAM) [29], a standard accepted throughout the field. 

The entry date was arbitrarily chosen to be November 3, 2010 (a MSL derivative). The atmosphere entry 

altitude is approximately 125 km. The planet model included a gravitational constant equal to 

4.2828376383e13 kg*m2

/s2 with a mean equatorial radius equal to 3396190 m, a mean polar radius of 

3376200 m and a planet rotation rate of 7.088218e-5 rad/sec. The gravity model included a JPL Mars 

85x85 model truncated at 20x20. The terrain model used 1/32nd deg Mars Orbiter Laser Altimeter 

(MOLA) data. 

Other EDL-SA simulation specific models, described in previous sections, include aerodynamic and 

aerothermodynamic models for the HIAD, SIAD and Rigid Aeroshell (Section 3.2) and tension cone 

based on a previous CEMENT study; mass properties, derived from response surface equations for 

HIAD, SIAD and Rigid Aeroshell (Section 3.1); a propulsion model (Section 3.4); various guidance 

routines of varying fidelity tailored for specific flight scenarios (Section 3.3). In addition EDL-SA 

assumes that the trajectories have perfect Navigation knowledge and that technologies required to 

perform precise landings will be leveraged from ALHAT. 

The purpose of the first year of the EDL SA was to examine alternative architectures that would 

provide the aerocapture phase required by DRA5 for the cargo delivery vehicles and the EDL required by 

both the cargo and crewed vehicles. The approach velocity and target orbit were provided by DRA5. To 

summarize: 1) the hyperbolic approach velocity was set at 7.36 km/s; 2) the target orbit was a 1Sol 

(33793km x 250 km); 3) EDL initiates from the 1 Sol orbit; 4) the landing site is at 0 m altitude; 5) the 

touchdown provides 10 m accuracy; and 6) the deceleration profiles remain within those limits set for a 

deconditioned crew. It was assumed for all the architectures that a reaction control system (RCS) would 

be the primary control. To emulate the characteristics of a RCS without having to design a control system, 

a “pseudo-controller” was used where the bank acceleration, maximum bank rate, and bank direction are 

modeled. 

The simulation was used for both the aerocapture and entry, descent and landing analysis. Details of 

nominal reference trajectories, sensitivity studies and Monte Carlo analysis results are presented in this 

section. 

4.1 Aerocapture Simulation Results 

The purpose of aerocapture is to use atmospheric drag to slow the vehicle from a hyperbolic velocity 

relative to the planet into an elliptical orbit. Figure 23 shows the aerocapture trajectory and major events. 

The aerocapture phase (events numbered 2-8) is followed by a periapsis raise burn (9), then a (nominally 

zero) cleanup burn (10) which place the vehicle into the desired post-aerocapture target orbit. Following a 

period of time in orbit, a de-orbit burn is performed in order to establish the desired atmospheric interface 

to initiate the EDL phase, which culminates with touchdown on the Martian surface. 
33

The eight combinations of architectures selected contain three different aerocapture scenarios. The 

first is a mid-L/D rigid aeroshell. The second is a 23 m diameter inflatable aeroshell (HIAD). The third 

configuration is a large HIAD (~55 m), sized to achieve a peak heat rate of 50 W/cm2 (3-σ). The decision 

process for the sizing of the 55 m HIAD is discussed in detail in 4.1.5, in the discussion of the Breakpoint 

Study. 

For consistency in the mass modeling, a ΔV budget is allocated for aerocapture not to exceed 150 

m/s. Also, based on historical practice and to ensure robust Monte Carlo analysis, the skip-out margin, 

defined as the difference between the flight path angle flown for the pass and the flight path angle, which 

results in a flyby, is required to be greater than 1 deg. Peak acceleration results are compared to the 

Human-Systems Integration Requirements (HSIR) [30] limits and checked to ensure that no exceedances 

have been observed. 

Section 3.3.1 describes the HYPAS guidance used to determine the aerocapture trajectories.

Figure 23. Aerocapture Phases 

4.1.1 Aerocapture Fligth Path Corridor Sizing 

Initially, the theoretical lift up-lift down profile was flown, in place of HYPAS, to gain an 

understanding of the expected performance for the three aeroshell designs. Values of skip-out margin can 

be obtained from studying these results. Skip-out margin is the difference between the flight path angle 

flown for an individual case and the flight path angle that causes skip-out. Figure 24 shows the amount of 

available skip-out margin for a range of ballistic numbers and L/D values. The flight path angle is chosen 

such that the maximum g-load is 4-Earth g’s. The figure also highlights the three aeroshell designs studies 

in this analysis according to each design’s L/D and ballistic number combination and the amount of skipout margin available for each design. Historically, a skip-out margin of 1 degree is required to see good 

performance (i.e. no cases lost to skip-out) in the Monte Carlo runs. The 23 m diameter inflatable 

aeroshell has 1.25 degrees, the 55 m diameter inflatable aeroshell has roughly 0.9 degrees, and the rigid 

aeroshell has approximately 1.6 degrees. Therefore, good performance is expected from Monte Carlo 

analysis for all configurations excluding the 55 m diameter inflatable aeroshell. 
34

Figure 24. Available Skip-out Margin for Variations of L/D and Ballistic Number 

4.1.2 Nominal Trajectory Results 

Reference trajectories were developed for the three different aeroshells using a 1 Sol (33793 km x 

250 km) post-aerocapture target orbit. For each of these references, the HYPAS analytical closed-loop 

guidance algorithm was used to generate the trajectories. 

 The simulation inputs include using the MarsGRAM 2005 atmosphere and the aerodynamic model 

described in Section 3.2.3. Additionally, an entry velocity of 7.36 km/s was selected, and the flight path 

angle at entry interface was chosen to meet the desired peak g-load for each case. Figure 25 compares the 

altitude profiles for each of the three configurations studied, and shows that the minimum altitude 

decreases as the ballistic number increases. The 55 m HIAD has the lowest ballistic number, and reaches 

a minimum altitude just below 50 km. Conversely, the rigid aeroshell is the highest ballistic number and 

reaches a minimum altitude of slightly higher than 30 km. 

Figure 26 and Figure 27 show the key trajectory guidance parameters for the 23 m HIAD nominal 

trajectory. These parameters include bank, drag acceleration, and altitude rate profiles. The bank profile, 

which the guidance algorithm commands in order to fly the drag acceleration and altitude rate profile 

reference trajectory, shows good performance in that the bank is consistently being modulated and does 

not saturate the guidance by holding full lift up (0 degrees) or full lift down (180 degrees) during the 

guided portion of the maneuver. Additionally, the drag acceleration and altitude rate profiles follow their 

reference profiles as expected. The drag acceleration profile is trying to pull the vehicle deeper into the 

atmosphere in order to avoid skip-out, whereas the altitude rate profile is attempting to maintain the 

equilibrium glide condition of constantly increasing altitude. However, the actual trajectory does not 

match the reference profiles exactly, as each profile is attempting to achieve competing objectives. 

Similar nominal trajectory data was generated for the 55 m HIAD and rigid aeroshell cases. All of these 

nominal trajectories were designed to a maximum of 3-Earth g’s. 
35

Figure 25. Altitude Profiles for All Configurations 

Figure 26. Nominal Bank Profiles for the 23 m HIAD 

Figure 27. Nominal Drag Acceleration and Attitude Rate Profile for 23 m HIAD 
36

4.1.3 Monte Carlo Results 

Sets of 2000 case Monte Carlo runs were completed for each of the reference profiles, and the 

HYPAS gains were optimized for ΔV performance within each set. The applied dispersions are listed in 

Table 11. 

Table 11. Aerocapture Monte Carlo Dispersion Values 

Dispersion Value

Aerodynamic +/- 10% (normal, 3σ)

Angle of Attack +/- 5% (normal, 3σ)

Entry Flight Path Angle (EFPA) +/- 0.35 deg (normal, 3σ)

Dust Tau 0.1:0.9 + random density perturbations

Mass +/- 5% (normal, 3σ)

Figure 28 through Figure 30 show key Monte Carlo results for each of the three aeroshell designs 

considered in this study. The figures are shown in order of decreasing ballistic number and L/D, with the 

rigid mid-L/D aeroshell having the highest combination of the two, and the 55 m HIAD having the 

lowest. With the exception of one extraneous high point in the Rigid Mid L/D Monte Carlo results, it is

observed that, as ballistic number and L/D decreases, the range of apoapse altitude reached and the 

amount of ΔV required to perform the post aerocapture clean up burn increases. There is a direct 

correlation between the size of the apoapse altitude target miss distance and the amount of ΔV required to 

clean up that miss distance. The larger the miss distance, the larger the clean up burn must be to put the 

vehicle into the target 1 sol orbit. 

The single high apoapse altitude observed in the rigid aeroshell is due to poor tuning in the guidance 

gains. With improved guidance tuning, the maximum ΔV value for that design could be brought below 

the maximum ΔV observed in the 23 m HIAD and 55 m HIAD cases which are well tuned and the 

performance observed is ideal. 

Figure 28. Monte Carlo Results for the Rigid Mid L/D Aeroshell 
37

Figure 29 Monte Carlo Results for the 23 m HIAD, L/D = 0.3 

Figure 30. Monte Carlo Results for the 55 m HIAD, L/D = 0.2 

The trend in larger apoapse altitude and higher ΔV values for the rigid aeroshell and the 55 m HIAD 

can be attributed to a decrease in L/D value. As stated previously, the larger the apoapse altitude, the 

larger the clean-up burn. The lower the L/D, the less lift the vehicle is able to command throughout the 

guided portion, which increases the difficulty the guidance has to achieve the target altitude rate profile. 

This, coupled with the worst of the Monte Carlo dispersions, creates individual cases in which the 

guidance, commanding the maximum amount of lift, is still unable to meet the altitude rate targets and 

exits the atmosphere with a large apoapsis altitude error. Though less than ideal performance is observed 

in Figure 10 for the L/D=0.2 case, the required ΔV is still below the budgeted 150 m/s, therefore, the 

performance has been deemed acceptable. 

4.1.4 Sensitivity Study Results 

Several sensitivity studies were also performed to evaluate the effect of additional parameter changes 

on the post aerocapture clean up ΔV. Nominal trajectories were simulated that parametrically varied 

atmosphere inputs such as season, dust opacity, and time of day over the full or expected range, as well as 

the initial mass and target orbit. Gain scaling (but not tuning) was performed for the mass sensitivity 
38

cases, and no gain changes were made for the atmosphere sensitivities. An examination of the results of

atmosphere and mass sensitivities resulted in no significant variation (less than 8 m/s) or trends in 

variation of ΔV primarily because of vehicle velocity and the fact that the guidance is capable of flying 

out the variations. That was not the case for the sensitivity study that considered various target orbits. 

The choice of target orbit, or post aerocapture apoapsis altitude (ranging from 33,793 km to 500 km) 

was determined to have a significant effect on the ΔV required for the post-aerocapture cleanup 

maneuver. Figure 31 shows the cleanup ΔV that results from the sensitivity that compared the aerocapture 

of a 23 m HIAD vehicle into both a 1 sol and 500 km circular post-aerocapture target orbit. A significant 

difference in mean ΔV, roughly 100 m/s, is observed. 

Figure 31. Delta V Sensitivity to Post-Aerocapture Orbit 

4.1.5 Trades 

Breakpoint Study – Architecture 2 assumes a 23 m HIAD, and due to the heat rate results from the 

aerocapture and entry simulations, the HIAD required heavy ablator TPS material. At the External Peer 

Review a request was made to examine the lower limits in L/D and ballistic coefficient that would enable 

the HIAD to maintain heat rates less than 50 W/cm2

 thus possibly requiring a lighter insulating-only TPS. 

Since the higher heat rates are seen during aerocapture, a study was initiated to determine the aerocapture 

HIAD performance “break point.” 

The range of L/D considered for the study was 0.1 to 0.3, and ballistic numbers ranged from 25 to 50 

kg/m2

. Sixteen specific combinations of L/D and ballistic numbers were considered for a scaled MIAS 

shape inflatable aeroshell configuration that aerocaptured into a 1sol orbit. Results show that performance 

is highly sensitive to L/D, and that poor performance is observed for L/D’s less than 0.25. 

Revisiting the concept of skip-out margin and understanding the amount available for each 

combination can provide insight into aerocapture performance. Figure 32 shows the skip-out margin 

available for the 16 combinations considered. The plot shows that for L/D values less than 0.25, the 

amount of skip-out margin is less than the one-degree required to obtain sufficient performance. 
39

Figure 32. Available Skip-out Margin for Low Combinations of L/D and Ballistic Number 

Using well-tuned guidance gains for a high L/D and ballistic number combination, the gains were scaled 

down and a 2000 case Monte Carlo simulation was performed for each of the sixteen combinations. 

Figure Figure 33 shows post aerocapture ΔV range required to target the 1 sol orbit for each of the 

combinations. 

Figure 33. Delta V Monte Carlo Results 

The key parameter to note is the maximum ΔV value required to clean up large apoapsis miss 

distance for each value of L/D. A large miss in apoapsis altitude requires a corresponding large amount of 
40

ΔV to clean up the orbit. Validating known poor performance at low skip out margins, the L/D=0.15 data 

has several cases that require more than 700 m/s of clean up ΔV, significantly exceeding the budgeted 

amount of 150 m/s. Even for a L/D=0.2, 300 m/s of ΔV is required, exceeding the budgeted amount, 

while the cases with L/D=0.25 and L/D=0.3 show acceptable performance. 

Figure 34 shows the relationship between the Monte Carlo 3-sigma peak heat rate values and HIAD 

diameter. Each line represents a different value of L/D, which is shown to have a relatively small effect 

on the peak heat rate. The data also show that the choice of HIAD diameter (or the effect of ballistic 

coefficient) has a significant effect on the peak heat rate. 

Initial mass and aerodynamic models limited the HIAD diameter to 50 m. The breakpoint study 

considered whether holding to a 3-sigma peak heat rate of less than 50 W/cm2

 would enable use of the 

same models by maintaining diameters below 50 m. Based on results shown in Figure 34, it is clear that 

no combination of ballistic number and L/D considered for the break point study is sufficient to meet both 

the peak heat rate and diameter constraints. Meeting the heating constraint was considered more 

important than meeting the diameter constraint; therefore, the HIAD diameter was allowed to increase to 

55 m. 

Figure 34. Plot of 3 sigma peak heat rate versus diameter for lines of constant L/D

The previous data include dispersed results for a scaled, but un-tuned set of guidance gains. The next 

step in characterizing the breakpoint was to choose a breakpoint set of cases for which to tune gains to 

examine improvements in performance. Gain tuning is a labor intense process, so only 4 of the 16 cases 

were selected for tuning. By examining Figure 30 combinations with a L/D lower than 0.2 exceeded the 

ΔV budget, therefore L/D = 0.2 was established as the L/D break point. Additionally, Figure 34 shows 

that HIAD diameters in the 50 to 55 m range come closest to satisfying heat rate constraint, therefore 

corresponding ballistic numbers of 41.67 and 33.33 kg/m2

 established the ballistic number breakpoint. 

Therefore, the two L/D and ballistic number breakpoints were combined to form the four combinations to 

be optimized using guidance gain tuning. 

The attempt to improve performance by tuning guidance gains was also improved by making a 

change in assumed entry flight path angle dispersion in the Monte Carlo. The previous Monte Carlo 

analysis had included a 0.35 deg variation in entry flight path angle. However it was felt that a reduction 

in the dispersion to 0.25 deg was acceptable and achievable by the time of the actual aerocapture flights. 

The results of these modification for the optimized L/D = 0.2, ballistic number of 33.33 kg/m2

 (or 55 m 

diameter) are shown in Figure 35. The effect of reducing the flight path angle dispersion and optimizing 

the guidance tuning results in a maximum ΔV (with the exception of one case circled in red) below the 
41

budgeted 150 m/s. 

L/D=0.15 cases (not shown) still remained almost 400 m/s over the budget, therefore, L/D=0.2 was 

determined to be the breakpoint in L/D. The final ballistic number breakpoint was selected as 33.33 

kg/m2

, which corresponds to the 55 m HIAD. In summary, the breakpoint of the aerocapture simulation 

using the HYPAS guidance algorithm was determined to be for cases below an L/D of 0.2 and ballistic 

number of 33.33. 

Figure 35. Effect of Flight Path Angle Dispersions for L/D=0.2, Bn=33.33 kg/m2

 

4.2 Entry, Descent and Landing 

4.2.1 Comparison of EDL-SA to DRA5 

EDL-SA Architecture 1 is based largely on and is an extension of the EDL work done for DRA5. 

Due to the scope of a reference mission, DRA5 considered only a single idealized (in-plane, nondispersed) nominal trajectory for Mars EDL. The performance of the theoretical guidance was verified by 

the NPC and Apollo-derived guidance algorithms. Due to lack of dispersion analysis in DRA5 and the 

assumption that a Mars cargo vehicle would have to demonstrate human rated g loads (~3g’s), the entry 

g’s were limited to a constant 2 g’s. However since DRA5, the Constellation Program Human-Systems 

Integration Requirements, HSIR, document has indicated that the entry g level can be as high as 4 or 

greater for short durations. So the entry g’s for EDL-SA has been constrained to a maximum of 4 g’s on 

entry including dispersions. Using linear feed back to maintain a constant 2 g entry, a throttle setting to 

maintain a constant 3g terminal descent and simple mass model, DRA5 provided an initial design point 

for modeling more detailed EDL phases for multiple EDL-SA architectures. 

Other design characteristics similar to both DRA5 and EDL-SA include entry from a 1 Sol orbit 

(33793 x 250 km), use of a 10x30 m rigid aeroshell (for the EDL-SA rigid mid L/D configurations), a 40 

mt useable landed payload, a 3g system thrust-to-weight and an 80 lbf/lbm engine thrust-to-weight. The 

touchdown phase for both studies, similar to Viking, reduced the vehicle velocity to 2.5 m/s for 5 seconds 

prior to touchdown at 0 km above the MOLA areoid at an equatorial landing site which was fixed for all 

EDL-SA architectures to allow for cross and down range dispersion evaluation.

Throughout the DRA5 EDL analysis it was evident that the design and conclusions relied heavily on 

a crude mass model. Initial mass estimates were extrapolated from vehicles orders of magnitude less 

massive than DRA5. Work was done for EDL-SA to incorporate response surface methodology to 
42

increase the fidelity of the structural masses, and thermal analysis was performed to obtain better 

estimates of TPS material mass to the first order. The result of incorporating the structural, TPS and 

propellant models into a single routine, which could be iterated to convergence within the POST2 

simulation, was a mass model that provided traceability to individual components as well as the ability to 

apply standard and growth margins appropriately. EDL-SA recognizes the limitations of the mass model 

and realizes that detailed packaging analysis of the components is required to further increase the fidelity 

of the model. 

Little is known about the ability to control the very large vehicles considered for EDL-SA. DRA5 

used only a “pseudo”, in-plane guidance. EDL-SA analysis incorporated a more flight-like theoretical 

guidance that included analysis of various control methods, namely using bank reversals and direct center 

of gravity control. For bank angle control realistic bank reversals and error bounds were set to provide 

three reversals, nominally. For center of gravity control a mass movement rate of one inch per second is 

assumed. Unlike a flight guidance algorithm, the theoretical guidance has full knowledge of the 

atmosphere and aerodynamics and is much quicker to implement and run in the simulation. The 

performance of the theoretical guidance was verified by both NPC and Apollo-derived guidance 

algorithms. 

4.2.2 EDL Strategy 

DRA5 – To summarize, the DRA5 simulation used a pseudo guidance activated at 1.75 g’s on entry. 

The pseudo guidance uses linear feedback to select the (constant) bank angle to maintain 2 g’s. The 

simulation was allowed to select the time to turn off the guidance and fly full lift up until selecting engine 

initiation. Engines were modeled to hold 3 g’s until the vehicle velocity was 2.5 m/s at which time the 

vehicle holds that velocity for 5 seconds prior to touch down. The entry trajectory is idealized but, as will 

be shown, provided a good approximation of the higher fidelity EDL-SA strategy. 

EDL-SA – The strategy of the EDL-SA exploration class EDL is similar for all architectures and is 

described in detail. Specific design parameters allowed to vary between architectures are noted. The entry 

phases include deorbit, atmospheric entry, pullout, and heading alignment. Descent phases include 

lander/drag device separation and terminal descent. The landing phase details vehicle touch down 

criteria. 

The first step in the exploration mission EDL design (and the primary difference from DRA5) is to 

converge the mass model using an in-plane guidance and to determine the appropriate entry point such 

that the trajectory ends at the preselected landing target. Controlled parameters include the deorbit ΔV 

from apoapsis of a 1 Sol orbit and the altitude at which the separation from the drag device occurs such 

that the vehicle lands within 50 m of the target. The POST2 simulation modifies the trajectory until 

parameters such as dynamic pressures and aero-heating characteristics permit mass model convergence. 

The first EDL phase, deorbit, begins with the POST2 simulation selecting the deorbit ΔV from the 

apoapsis of the 1 Sol elliptical orbit to modify heat rate and dynamic pressure as needed to minimize the 

arrival mass. The second phase, atmospheric entry, begins at a geographic radius of 3255.2 km and the 

atmosphere is activated. The third phase, pullout, performed after the mass model has converged, is 

achieved using the theoretical guidance that employs either a bank reversal strategy based on allowable 

heading error as a function of velocity, another notable difference from DRA5. 

For bank control, if the error, input by the user, is exceeded, a bank reversal is commanded. The 

allowable error may be different for each architecture but will nominally produce 3 reversals. POST2 is 

also allowed to select the initial bank angle, which is approximately 90 deg to allow for margin from 

either the full lift up (bank angle = 0 deg) or full lift down (bank angle = 180 deg) configuration. The 

guidance is initiated at 1.25 g’s for EDL-SA. 
43

The next phase, heading alignment, begins at a specified velocity; the value may vary between 

architectures. Bank angle accelerations of five deg/s2

 and a max rate of 20 deg/s are based in realistic 

flight controllers. During the heading alignment phase, the objective is to drive the azimuth error to the 

landing site target to 0 deg at engine initiation. The velocity at which heading alignment begins as well as 

the allowable errors required to produce three reversals, nominally, may vary between architectures. 

These phases summarize the entry portion of the trajectory. 

The descent begins with the separation phase in which the simulation determines the altitude to 

jettison the entry drag device (rigid aeroshell, HIAD, or SIAD). For DRA5 the event was assumed to be 

instantaneous. However, the lack of higher fidelity vehicle separation models and the desire to more 

realistically model the event, the mass of the decelerator is instantaneously jettisoned and the Lander 

vehicle is allowed to free-fall for a given amount of time to ensure no recontact. Table 12 provides the 

separation times for each of the architectures based on assumptions made about the nature of the 

separation. Details of work done to further characterize the separation or transition events can be found in 

Appendix B. After the allotted time, the terminal descent engines are initiated, commencing the terminal 

descent phase. Terminal descent design is governed by the desire to balance the fuel saved of using 

higher throttle settings (therefore higher g’s) at a low altitude close to the landing site with maintaining 

control authority to fly out errors to allow precise landing capability (within 50 m of the target). 

Therefore the terminal descent is designed to use, nominally, a maximum throttle setting of 80% to reach 

the desired maximum nominal acceleration (2.5 g’s) for only an instant before throttling back to fly a 

linear velocity curve with respect to changes in altitude. 

The final phase, landing, also known as the touch down phase, is based on Viking heritage. Once the 

vehicle reaches 2.5 m/s, the velocity is held constant for 5 second until the vehicle touches down at 0 km 

above the MOLA areoid at -1.18 deg North latitude and 182 deg East longitude. 

Table 12. Architecture Component Separation Times 

Architecture Separation Time (s)

1 Rigid areoshell from descent stage 20

2,4,5,6 HIAD from descent stage 15

7 Rigid aeroshell from SIAD 20

7 SIAD from descent stage 10

8 HIAD and SIAD from descent stage 10

As mentioned previously, certain parameter values like deorbit ΔV, altitude at separation, heading 

alignment initiation and azimuth errors, as well as additional separation and deployment events required 

in Architectures 7 and 8, can vary between architectures. However, the general strategy for each 

architecture remains the same. Figure 36 shows the reference altitude versus velocity plot for each 

architecture. 
44

Figure 36. EDL-SA Exploration Class Altitude Versus Velocity 

4.2.3 Nominal and Monte Carlo Results 

Each exploration architecture was simulated using POST2 according to the strategy and models 

outlined previously. Unlike DRA5, Monte Carlo dispersion analysis (using 2000 cases) was preformed 

on most architectures, including dispersion on aerodynamics, entry state and the atmosphere. Table 13 

shows the nominal value and perturbation for each Monte Carlo input parameter. The aerodynamic 

coefficient multipliers are applied to aerodynamic database output and the atmospheric model 

(MarsGRAM) dispersions and winds are applied to all cases except the nominal. Detailed nominal 

trajectory and Monte Carlo results for the architectures are provided below. 

Table 13. EDL-SA Monte Carlo Dispersions 

Parameter Nominal Value Perturbation Distribution

Nominal Coefficient multiplier 1 +/- 10 % (3σ) Normal

Axial Coefficient multiplier 1 +/- 10 % (3σ) Normal

Angle of Attack (deg) 55.0, -22.2, -22.2, 55.0, -22.2 +/- 5 (3σ) Normal

Engine ISP (sec) 369 +/- 2.5 % (3σ) Normal

Entry FPA (deg) -10.8, -9.9, -9.0, -10.8, -9.8 +/- 0.25 (3σ) Normal

Atmospheric Random Number 1 1-29999 Uniform

Dusttau 0.7 0.1:0.9 Uniform

Architecture 1 – The Architecture 1 entry configuration, selected for its similarity to DRA5, includes 

a mid-L/D rigid aeroshell for aerocapture and hypersonic flight and supersonic retro-propulsion for 

descent and landing. Figure 37 shows the timeline below 30 km. The vehicle flies at 55 degrees angle of 

attack and has a L/D of 0.5. The heading alignment phase begins at 1700 m/s and the nominal terminal 

descent throttle setting is set at 80%. However, for the Monte Carlo analysis, the throttle setting is 

allowed to increase to 90%, which result in several cases that fly a more aggressive descent profile and 

use considerably less propellant than the nominal case. Additional 1 and 99 percentile Monte Carlo 

results, including Mach, dynamic pressure and altitude dispersions at aeroshell separation and terminal 

descent initiation, are included in Table 14. The DRA5 arrival mass estimate of 110.9 mt is remarkably 
45

close to the higher fidelity EDL-SA arrival mass of 110.1 mt shown in Table 20, implying that the 

approximations made for DRA5 adequately represent current knowledge. Because of the use of an 

aeroshell, Architecture 1 is considered to be at a higher TRL than the drag devices considered in other 

architectures. However, several technology challenges remain for the aeroshell including packaging, 

detailed modeling of the separation event and engine initiation at supersonic velocities. 

Architecture 2 – Architecture 2, which uses a HIAD for aeorcapture and entry along with supersonic 

retro-propulsion for terminal descent, was selected to evaluate the mass savings of using a dual use TPS 

HIAD over a rigid mid-L/D areoshell. The 23 m inflatable design is based on the MIAS concept and is 

currently at a lower TRL than the rigid aeroshell. The angle of attack for the vehicle is -22 degrees and 

the L/D is 0.33. Percentile dispersion values are provided in Table 15. The plot of the altitude versus time 

curve for Architecture 2 is shown in Figure 38. Plotted for reference is the similar curve for Architecture 

1. Note that for Architecture 2 the 3-sigma low range to target at terminal descent initiation has been 

reduce from 7 km in Architecture 1 to 2.7 km resulting in reduced timeline margin for terminal descent. 

However, use of the HIAD, based on the mass models developed for EDL-SA, reduces the entry mass 

over Architecture 1 by nearly 25 mt, making it an advantageous alternative to Architecture 1, despite 

technical challenges including HIAD packaging, separation and dual-use TPS. 

 

 Figure 37. Architecture 1 Altitude vs. Time 

 Figure 38. Architecture 2 Altitude vs. Time

1% Mean 99%

Aeroshell Separation

 Mach 2.4 2.7 3.3

 Dyn. Pres., N/m2 1077.7 1283.4 1598.3

 Altitude, km 6.5 7.5 8.9

Terminal Descent Initiation 

 Dyn. Pres., N/m2 1479.2 1721.8 2049.4

 Altitude, km 3.3 4.6 6.9

 Range To Targ, km 6.7 10.9 25.4

 Prop Use, t 12.4 14.2 16.9

1% Mean 99%

HIAD Separation

 Mach 1.6 1.8 2.1

 Dyn. Pres., N/m2 563.3 646.4 780.0

 Altitude, km 5.1 5.7 6.5

Terminal Descent Initiation 

 Dyn. Pres., N/m2 775.3 881.8 1027.2

 Altitude, km 2.5 3.3 4.6

 Range To Targ, km 3.3 4.9 8.5

 Prop Use, t 9.0 10.0 11.4

Table 14. Architecture 1 Results 

Table 15. Architecture 2 Results
46

Architecture 3 – Architecture 3, the all-propulsive entry configuration, was the least complex EDL 

alternative and, using only engines, was considered to be at a relatively high TRL compared to the other 

architectures. However, there are major technology questions concerning the flowfield interactions of the 

rocket plume firing into the atmosphere at hypersonic and supersonic conditions as illustrated in Figure 

39 [31]. Unknown are the impact of the total pressure on the nozzle performance and internal nozzle flow 

(possible flow separation), the flow stability and the resulting dynamics on vehicle control, and the 

resulting drag and aerodynamic heating. 

Figure 39. Rocket Plume Hypersonic Flowfield Interactions 

Because of the unknowns in drag and aerodynamic heating from the diverted forebody flowfield, the 

initial study results, used to determine feasibility, included no drag and no aerodynamic heating. Because 

no estimate of drag was made in this initial assessment, the conservative approach was to assume that no 

drag would be used to decelerate the system to reduce the ΔV. The same flow-diversion assumptions were 

also made for the aerodynamic heating; thus no thermal protection system was used. 

Therefore, Architecture 3 assumed that a nuclear thermal rocket, NTR, performs the Mars Insertion 

burn to place the vehicle into a 1 Sol orbit. The strategy for deorbit and entry from 1 Sol was optimized 

using an instantaneous burn (ΔV=1200 m/s) to transition to an intermediate orbit of 250 x 250 km orbit 

and then a deorbit burn of 186 m/s thereby minimizing the arrival mass and gravity losses. 

The trajectory simulation was initiated at the deorbit condition. Initial thrust-to-weight ratio (T/W) 

was varied from 1.01 to 3.7 Mars g’s, and the altitude at thrust initiation was optimized. At a T/W of 

1.010, the thrust initiation altitude was 250 km, and at T/W = 3.7 the altitude was 50 km. With 

atmospheric interface at approximately 125 km, the cases for T/W greater than 1.8 may not be valid with 

the no-drag assumption because the nozzle and vehicle forebody are not protected with diverted flow 

unless a very low-thrust flow divert could be used. A unique EXAMINE based mass model was 

developed for Architecture 3 and is summarized in Section 3.1.4 and includes no TPS mass. The deorbit 

mass results of the trade study can be seen in the solid lines in Figure 40. No cases exceeded the HSIR 

requirements for human g levels because the maximum deceleration was 3.7 Mars g’s or 1.2 Earth g’s. 
47

The lowest deorbit mass resulted from a system T/W of 1.01 Mars g’s and a total ΔV of 5292 m/s. Using 

the study baseline LOX/CH4 engines with a specific impulse of 369 seconds, the arrival mass was 310 mt 

which is 282% higher than Architecture 1. 

Figure 40. Architecture 3 Arrival Mass and Delta V 

The initial study shows that an all-propulsive entry is feasible, albeit with much higher arrival masses 

than the other architectures considered, and it is noted that the initial study neglected several key 

parameters. Namely that the drag from recirculization of the flow on the forebody, skin friction drag on 

the sides, and base drag on the aft body is not negligible and that the no-heating assumption that utilized 

no TPS may be highly optimistic. Minimal mass margins were carried in this study, which differed from 

the margin policy used for the other Architectures. Therefore, for comparison, a second study was 

performed that accounted for the base drag of the aft body assuming a circular area with a nine-meter 

diameter. The trajectory was designed such that it maintained stagnation heat rates below 25 W/cm2

, 

which would eliminate the need for TPS. A mass margin policy was included that was similar to that used 

to evaluate the other Architectures. Also, because atmosphere drag was include, a larger (though not 

optimized) intermediate orbit of 400 x 250 km was considered. As the initial thrust to weight of the 

vehicle was increased, the altitude at which the engines are ignited also drops, resulting in larger losses 

due to drag. The increases in drag loss result in less propellant required to slow the vehicle and therefore, 

the arrival mass is reduced. The arrival mass corresponding to the trajectories that include drag loss, 

thermal constraints and margined masses are shown in the dashed line in Figure 40. Lower arrival masses 

can be achieved using larger initial thrust to weight. 

The arrival mass can further be reduced by optimizing the intermediate orbit and/or using advanced 

technology or staging. Using LOX/LH2 engines with a specific impulse of 450 seconds, which are the 

same technology engines as the cryogenic DRA5 Mars Transfer Vehicle option, the mass may be reduced 

to 218 mt. This option also requires the same zero-boiloff technology as the Earth-to-Mars transfer 

vehicles. Adding a second LOX/LH2 stage, the mass may be further reduced to 186 mt; however, this 

staging benefit must be traded with the additional cost and reduced reliability of using a second stage with 

a vehicle that is still 169% heavier than Architecture 1. It is noted that there is also no TPS or mass 

margin included in these estimates. No Monte Carlo analyses were performed for Architecture 3. 

Architecture 4 – The entry portion of Architecture 4 is identical to Architecture 2. However, 

consideration was given to Architecture 4 to compare the mass savings of using a single use (entry only) 

TPS on a 23 m HIAD to the dual use (aerocapture and entry) version used in Architecture 2. Therefore, 
48

different from Architecture 2, Architecture 4 uses a rigid areoshell for aerocapture, then, prior to entry, 

the HIAD is inflated. The result is a larger arrival mass of 109.0 mt compared to Architecture 2’s 83.6 mt 

but the entry mass is lighter by almost 8 mt due, in part, to the 4 mt lighter HIAD mass. Because of the 

difference in the entry ballistic coefficient, the architectures have slightly different trajectories but the 

Monte Carlo analysis results are nearly the same. See Figure 41 and Table 16. The component mass 

numbers are compared in Table 20. The increase in launch mass and mission complexity of using a 

separate aerocapture system, in addition to the technical challenges already present in Architecture 2, did 

not make this configuration appear practical. 

 Figure 41. Architecture 4 Altitude vs. Time 

Architectures 5 and 6 – Architectures 5 and 6 were selected to compare the arrival mass savings of 

using a very large HIAD down to subsonic speeds as an alternative to supersonic retro-propulsion, 

considered in Architecture 2 and 4. (This is a contingency in the event that supersonic retro-propulsion 

proved to be an unusable option for an exploration class mission.) Architecture 5, like Architecture 4, 

compared the mass saving of using a single-use TPS HIAD for entry and a rigid aeroshell for aerocapture. 

The simulation optimized the HIAD diameters to be 68 m and 82 m for Architectures 5 and 6 

respectively. While the mass of the aeroshell and smaller HIAD in Architecture 5 did result in a lower 

arrival mass than Architecture 6’s dual use HIAD, there was an issue in the mass model used for both 

architectures that calls all the results into question. The issue is that the TPS mass model used for the 

large HIADs was limited to a 50 m diameter. Extrapolation beyond that diameter resulted in the 

assumption that the areal density of the material was constant, which is not likely to be the case. Also 

there are additional EDL timeline challenges of successfully slowing the vehicle to subsonic speeds with 

adequate altitude margin to initiate the terminal descent engines and land at the target. The timeline for 

Architecture 6 is provided in Figure 42. Architecture 5 is very similar. A low throttle setting of 

nominally 65% is used to slow the vehicle starting at approximately 2.5 km such that the vehicle can 

touch down at the target. The large diameter HIAD flies at the same angle of attack -22 deg and same 

L/D as in Architecture 2. Monte Carlo one and 99% dispersions are also provided in Table 17. However, 

the results indicate that investments in supersonic retro-propulsion might be a more prudent choice over 

very large diameter HIAD systems, which will have packaging and separation technical challenges. 

1% Mean 99%

HIAD Separation

 Mach 1.5 1.7 1.9

 Dyn. Pres., N/m2 493.1 557.4 657.6

 Altitude, km 5.3 5.9 6.7

Terminal Descent Initiation 

 Dyn. Pres., N/m2 690.5 781.1 895.9

 Altitude, km 2.7 3.3 4.5

 Range To Targ, km 3.1 4.4 7.2

 Prop Use, t 8.8 9.6 10.9

Table 16. Architecture 4 Results
49

 

 

1% Mean 99%

HIAD Separation

 Mach 0.7 0.7 0.7

 Dyn. Pres., N/m2 93.3 99.5 106.3

 Altitude, km 5.0 5.2 5.4

Terminal Descent Initiation 

 Dyn. Pres., N/m2 199.9 214.1 230.4

 Altitude, km 2.6 2.7 2.9

 Range To Targ, km 0.6 0.7 1.0

 Prop Use, t 6.7 6.9 7.2

 Figure 42. Architecture 6 Altidue vs. Time 

 Architecture 7 – Architecture 7 was selected as an alternative to Architecture 1 assuming that 

supersonic retro-propulsion was an infeasible option for exploration-class missions. Architecture 7 

replaces supersonic retro-propulsion with a 51 m diameter SIAD and subsonic retro-propulsion. The 

nominal timeline compared to Architecture 1 is shown in Figure 43. Monte Carlo statistics are provided 

in Table 18. The entry strategy remains the same. To accommodate wind effects on the large unguided 

SIAD, a cross range offset at entry was included in the simulation. The altitude at engine initiation was 

also maximized to account for the winds during the unguided portion of the trajectory. Therefore, the 

maximum throttle setting during terminal decent was reduced from 80 to 65%, as well as the system 

thrust-to-weight from 3.0 to 2.5 g’s. In addition to the aeroshell packaging and separation technical 

challenges, Architecture 7 has SIAD packaging, deployment, inflation and separation technical 

challenges. In the end, the added complexity of EDL for Architecture 7, in particular the unguided 

portion of the trajectory on the SIAD, results in a mass savings of only 3 mt of Mars arrival mass 

compared to Architecture 1. 

1% Mean 99%

Aeroshell Separation

 Mach 0.7 0.7 0.8

 Dyn. Pres., N/m2 115.3 125.3 133.6

 Altitude, km 4.3 4.5 4.6

SIAD Full Inflation 

 Mach 3.5 3.8 4.1

 Dyn. Pres., N/m2 983.6 1151.0 1406.1

 Altitude, km 12.2 16.0 18.5

SIAD Separation

 Mach 4.4 4.6 4.8

 Dyn. Pres., N/m2 1096.7 1416.6 2188.1

 Altitude, km 12.8 17.5 20.3

Terminal Descent Initiation 

 Dyn. Pres., N/m2 190.3 209.8 225.4

 Altitude, km 2.6 2.7 2.8

 Range To Targ, km 0.0 0.5 1.5

 Prop Use, t 6.9 7.4 8.2

Figure 43. Architecture 7 Altitude vs. Time

Table 17. Architecture 6 Results

Table 18. Architecture 7 Results
50

Architecture 8 – Architecture 8 considered a HIAD plus SIAD configuration as an alternative to a 

single large HIAD from Architecture 6 to reduce the vehicle to subsonic speeds prior to engine initiation. 

The HIAD/SIAD combination included the same 23 m HIAD used in Architecture 2 and 4. The 

simulation determined the size of the SIAD (44 m), and it was deployed nominally at Mach 2.6, such that 

the terminal descent engine initiation would occur subsonicly. The inflatable designs were based on the 

MIAS and IRDT studies. The advantage of a SIAD is the lighter mass due to deployment after peak 

heating, which eliminates the need for TPS, while still providing drag to slow the vehicle to subsonic 

speeds at engine initiation. To compare the mass savings of a HIAD/SIAD system to a larger HIAD only 

system, the results of the mass model suggest that using a SIAD can save approximately 40 mt of HIAD 

mass while costing only approximately 2 mt in SIAD mass for a net savings in Architecture 8 of 

approximately 38 mt over Architecture 6. See Table 20. However note the caveat of extrapolating HIAD 

diameter used in Architecture 6. The mass savings in reality may not be as significant. However, 

comparing use of a HIAD/SIAD combination to using supersonic retro-propulsion in Architecture 2, the 

use of a SIAD only saves approximately 3 mt of arrival mass. Despite the mass advantages, combinations 

of inflatables are arguably at a lower TRL than single inflatable structures, and in this particular 

configuration, also has additional technical challenges associated with the SIAD, increased complexity of 

EDL and a compressed the EDL timeline. See Figure 44 and Table 19. 

1% Mean 99%

Aeroshell Separation

 Mach 0.7 0.8 0.8

 Dyn. Pres., N/m2 131.5 144.1 155.6

 Altitude, km 3.7 3.9 4.0

SIAD Full Inflation 

 Mach 2.4 2.6 2.7

 Dyn. Pres., N/m2 654.7 775.7 958.4

 Altitude, km 8.4 11.9 14.3

SIAD Separation

 Mach 2.9 3.0 3.2

 Dyn. Pres., N/m2 804.6 1008.0 1379.0

 Altitude, km 8.9 12.5 15.0

Terminal Descent Initiation 

 Dyn. Pres., N/m2 195.8 222.0 242.6

 Altitude, km 2.2 2.3 2.5

 Range To Targ, km 1.0 1.1 1.5

 Prop Use, t 6.2 6.4 6.7

 Mass – It is acknowledged that there are many cost functions by which to evaluate the various 

technologies and Architectures in this study. It is difficult to quantify timeline margin when the details of 

transitions are only now being investigated. It is difficult to apply error margins on the mass model 

because there are few data points with which to validate the model. However, mass is the easiest to 

quantify. For this reason, the primary comparisons between architectures have been component masses 

that include standard margins and the values are provided for each Architecture in Table 20. The 

components in the table include a standard 30% mass margin as well as a 15% mass growth allowance. It 

is also recognized that environmental parameters are of interest for defining and identifying requirements 

for various technologies. For that reason Table 21has been included. 

Figure 44. Architecture 8 Altitude vs. Time

Table 19. Architecture 8 Results
51

Table 20. Component Dimensions and Masses using Standard Margins 

Architecture 1 2 3 4 5 6 7 8

HIAD Diameter (m)

Aerocapture/Entry -- 23.0 -- 23.0 67.8 81.9 -- 23.0

SIAD Diameter (m) -- -- -- -- -- -- 44.6 44.3

Aerocapture/Entry BN 

(kg/m2) 396.1 154.0 -- 134.6 20.6 20.4 412.0 152.5

Descent BN (kg/m2) -- -- -- -- -- -- 27.2 40.8

Mass (mt)

Arrival Mass 110.1 83.6 265.2 109.0 133.5 140.5 107.4 80.6

Deorbit Mass 109.2 82.8 188.1 75.1 98.0 139.4 106.3 79.6

 Aeroshell, AS (Total) 28.9 0.0 -- 26.0 27.0 0.0 28.8 0.0

 AS Structure 18.3 -- -- 18.4 19.3 -- 18.2 --

 AS TPS 10.6 -- -- 7.6 7.6 -- 10.6 --

Avionics and Separation 

Structure 2.0 2.1 -- 4.3 7.4 8.5 2.0 1.9

Entry RCS (Total) 10.8 7.1 -- 9.9 14.7 16.8 11.7 7.4

 RCS Dry Mass 5.2 2.7 -- 5.8 9.2 9.9 6.1 2.9

 RCS Propellant 5.5 4.4 4.1 5.5 6.9 5.6 4.5

 HIAD (Total) -- 10.6 -- 6.0 25.7 56.0 -- 10.6

 HIAD Structure -- 6.0 -- 3.1 10.7 22.3 -- 5.9

 HIAD TPS -- 4.7 -- 2.9 15.0 33.7 -- 4.7

 SIAD Mass -- -- -- -- -- -- 6.8 2.1

Descent Stage, DS 28.4 23.8 148.1 22.8 18.7 19.2 18.1 18.6

 DS Dry Mass 12.3 11.7 19.2 11.2 9.9 10.5 10.2 10.4

 DS Propellant 16.2 12.0 128.9 11.6 8.8 8.7 7.9 8.2

Landed Mass 52.3 51.8 74.7 51.2 49.9 50.5 50.2 50.4

Payload Mass 40.0 40.0 40.0 40.0 40.0 40.0 40.0 40.0
52

Table 21. Nominal Environment Conditions 

Architecture 1 2 3 4 5 6 7 8

Aerocapture

Peak Stag Heat Rate (W/cm2) 122.8 74.6 -- -- -- 36.6 -- --

Peak Ames Heat Rate (W/cm2) 335.8 88.8 -- -- -- 36.7 -- --

Total Stag Heat Load (J/cm2) 13563.8 7665.3 -- -- -- 3894.9 -- --

Total Ames Heat Load (J/cm2) 34499.4 7944.4 -- -- -- 3433.4 -- --

Peak Dynamic Pres. (N/m2) 10600.9 4072.6 -- -- -- 1061.6 -- --

Peak Earth G load 3.0 2.9 -- -- -- 3.0 -- --

Entry

Deorbit ΔV (m/s) 15.3 14.0 1307.7 13.9 12.6 12.6 15.1 14.0

Peak Stag Heat Rate (W/cm2) 19.7 27.3 0.7 25.6 9.9 9.9 19.5 27.2

Peak Ames Heat Rate (W/cm2) 119.5 27.0 5.0 24.4 6.8 5.2 117.9 26.8

Total Stag Heat Load (J/cm2) 2270.7 3070.5 102.6 2875.6 1149.5 1143.8 2363.6 3047.4

Total Ames Heat Load (J/cm2) 13297.4 2610.0 661.9 2403.1 723.4 547.7 13833.8 2593.1

Peak Dynamic Pres. (N/m2) 10506.5 4240.1 1823.1 3706.2 572.4 569.2 10961.2 4194.5

Peak Earth G load 3.0 3.0 2.0 3.0 3.0 3.0 5.2 3.0

Descent ΔV (m/s) 797.8 601.0 3701.5 585.9 447.1 435.8 438.2 406.4

4.2.4 Sensitivities Study Results 

A set of sensitivity analyses were run to understand how the unique capability of each architecture, 

such as ballistic number, lift and drag performance and guidance, perform against the following: 

environmental variability, engine characteristics, landing site, aerodynamics and vehicle size. The 

sensitivities from this study are meant to show how each architecture performs against the Figures of 

Merit for EDL-SA and are broken down into three categories: Environment, Performance and Design. 

The sensitivities for Architecture 3, 4, 5 and 9 are not presented because of the all-propulsive design, the 

duplication of EDL design for Architectures 4 and 5 and the late addition of Architecture 9. 

Environmental Sensitivity – The Environmental Sensitivity separately varied the nominal trajectory 

for each architecture with season (Ls), dust opacity, landing altitude above MOLA, latitude and time of 

day at Mars. Table 22 shows the results of the Environmental sensitivity, quantified by total mass 

variation per architecture. This analysis shows that the exploration class vehicle performance, in terms of 

mass, is insensitive to environmental effects. 
53

Table 22. Environmental Sensitivity 

Sensitivity Nominal

Value

Sensitivity

Range

Total Mass Variation (kg)

( %) of total entry mass

 Architecture 1 2 6 7 8

Season (Ls) ~174.5 0:30:360 deg

Dust opacity

MarsGRAM dusttau 0.7 0.1:0.2:0.9

688

(0.7%)

863

(1.1%)

660

(0.7%)

334

(0.4%)

367

(0.5%)

Landing Altitude

Above MOLA 0 km -4:0.2: 2.5 km 3724(3.6%)

2582

(3.4%)

1860

(2%)

1075

(1.1%)

464

(0.6%)

Latitude -1.177 -75:15:75 deg 1555

(1.5%)

2053

(2.7%)

2146

(2.3%)

941

(1%)

890

(1.2%)

Time of Day 5:30 am 0:1.5:24 hours 591

(0.6%)

574

(0.8%)

494

(0.5%)

430 

(0.4%)

226

(0.3%)

Time of Day 5:30 am 0:1.5:24 hours 591

(0.6%)

574

(0.8%)

494

(0.5%)

430

(0.4%)

226

(0.3%)

Performance Sensitivity – The Performance Sensitivity varied engine thrust-to-weight and specific 

impulse and retro-propulsion aerodynamics of each architecture’s nominal trajectory. Table 23 shows the 

Performance Sensitivity results for Architectures 1, 2 and 6 through 8. The thrust-to-weight sensitivity 

showed that, while it is important for powered descent performance, it is insensitive to arrival mass. 

Specific impulse variation for each of the architectures also showed insensitivity in that the low time-onengines result in a minimal effect of specific impulse on arrival mass. Historical investigations to 

supersonic retro-propulsion have indicated engine placement can augment drag performance. However, 

drag only accounts for 2.5% of powered descent ΔV, making the supersonic retro-propulsion 

aerodynamic augmentation insensitive and providing minimal benefit to arrival mass among all 

architectures. 

Table 23. Performance Sensitivity 

Sensitivity Nominal

Value

Sensitivity

Range

Arrival Mass Change

 Architecture 1 2 6 7 8

Engine Thrust 

to Weight

80 

lbf/lbm

50:5:90 -1.1 MT

(-1 %) / 

10(lbf/lbm)

-1 MT 

(-1.4 %) / 

10(lbf/lbm)

-1.3 MT 

(-1.4%) / 

10(lbf/lbm)

-0.9 MT 

(-0.9%)/ 

10(lbf/lbm)

-0.91 MT

(-1.2%) / 

10(lbf/lbm)

Vehicle Thrust 

to Weight

3 g’s 2:0.25:4 0.73 MT 

(0.7%) / g

0.86 MT 

(1.1%) / g

-0.91 MT 

(-1.0%)/g

0.63 MT 

(0.7%)/g

1.1 MT 

(1.5%)/g

Specific 

Impulse

369 sec 355:2.5:37

5

-0.61 MT 

(-0.6 %) / 

10 sec

-0.43 MT 

(-0.6%) / 

10 sec

-0.31 MT 

(-0.3 %) / 

10 sec

-0.24 MT 

(-0.2 %) / 

10 sec

-0.25 MT 

(-0.3 %) / 

10 sec

Supersonic 

Aerodynamic

Augmentation

Ca=0 Ca=0:2 -0.1 MT 

(-0.1 %) / 

10% Ca

-0.17 MT 

(-0.22 %) / 

10% Ca

N/A N/A N/A
54

Design Sensitivity – The Design Sensitivity focused on the effect of varying the divert maneuver, 

extra propellant, payload and lift-to-drag (L/D). Table 24 shows the results of the Design Sensitivities for 

each architecture. The L/D sensitivity looked at increasing L/D while maintaining the same vehicle 

ballistic number, which by increasing L/D up to 10%, has minimal sensitivity to arrival mass. The divert 

maneuver is included to reduce the risk of aeroshell far-field recontact after separation and uses only beta 

control (modified gravity turn) to perform. For architectures with a low engine ignition velocity, the 

trajectory does not have the necessary energy to make large diverts. For large diverts, Architectures 6 and 

8 are not optimal, but provide the most sensitivity to arrival mass. Architectures 1 and 2 carry more 

energy at terminal descent ignition allowing for larger diverts, making them least sensitive with arrival 

mass. However, the EDL design of Architecture 7 makes it very difficult for it to perform any kind of 

divert and was therefore not analyzed. Payload is modeled as landed usable mass on the Mars surface and 

does not include any of the EDL system mass. The payload sensitivity showed that mass growth is a 

linear relationship of arrival mass versus payload mass and showed similar linearity between each 

architecture. 

Table 24. Design Sensitivity 

Sensitivity Nominal

Value

Sensitivity

Range

Arrival Mass Change

 Architecture 1 2 6 7 8

Divert 

Maneuver

0 km 0:0.5:3.0 -0.1 MT

(-0.1 %) / 

0.5 km

-0.3 MT 

(-0.4 %) / 

0.5 km

-4.1 MT 

(-4.4 %) / 

0.5 km

N/A -2.5 MT 

(-3.4 %) / 

0.5 km

Extra 

Propellant

0 MT 0:0.5:5.0 1.9 MT 

(1.84%) / 

MT Prop

1.93 MT 

(2.51 %) / 

MT Prop

2.72 MT 

(2.95%) / 

MT Prop

1.74 MT 

(1.8 %) / 

MT Prop

1.74 MT 

(2.4 %)/ 

MT Prop

Payload 40 MT 10:5:60 1.72 MT 

(1.67 %) / 

MT 

Payload

1.79 MT 

(2.33 %) / 

MT 

Payload

2.29 MT 

(2.48 %) 

/ MT 

Payload

1.54 MT 

(1.59 %) 

/ MT 

Payload

1.71 MT 

(2.34 %) 

/ MT 

Payload

L/D 0.51 Rigid 

0.33 Inflat

75%:125

%

-0.782 MT 

(-0.76 %) / 

10% Lift

-0.485 MT 

(-0.63 %) / 

10% Lift

N/A N/A N/A

Summary 

In summary, the impracticality of designing and carrying separate systems for aerocapture and entry 

has rendered the Architecture 4 and 5 unlikely candidates for further study. Decreasing engine throttle 

settings and igniting engines at higher altitudes has mitigated the small range-to-targets at subsonic engine 

initiation in Architectures 6, 7 and 8. The results have been included in the nominal and Monte Carlo 

analysis; however, the sensitivities studies have not been updated to reflect the performance of the new 

configuration. The more significant issue is accurate mass modeling of large HIADs (>50 m diameter). 

The assumptions made in this study, namely to assume constant areal density for diameters larger than 50 

m is probably too conservative. Current mass models perhaps unrealistically penalize those architectures, 

namely 5 and 6, although packaging and inflation issues remain. The compressed EDL timeline that 

results from using subsonic retro-propulsion makes Architectures 6, 7, and 8 less likely candidates for 

exploration-class missions but should not be ruled out for robotic missions. Finally, Architecture 1 and 2, 

with their relatively less complex EDL, indicate that the HIAD, mid L/D aeroshell and supersonic retropropulsion are technologies to be considered in more detailed study for exploration class missions. 
55

5 Packaging 

5.1 Preliminary Packaging 

A preliminary packaging analysis was performed to aid in stage sizing and configuration transition 

analyses. Engines, deceleration devices, propellant tanks, and major structural elements were packaged 

for each architecture approach. Figure 45 shows the packaging for Architecture 2, and includes a stowed 

HIAD on the far right side that would be deployed on-orbit prior to entry. All required elements can 

easily be accommodated within a 20 m length, rather than the allotted 30 m. The propellant tanks are 

only a small portion of the required volume, which is dominated by the engines and the habitat and 

airlock. Significant volume remains for packaging subsystems and cargo (e.g., rovers). 

Figure 45. Architecture 2 Packaging 

5.2 HIAD Packaging 

HIAD packaging studies were initiated as a part of EDL-SA’s conceptual development of the 23 m 

HIAD TPS. These studies have been carried forward into the EDL TDP Flexible Ablator TPS activities. 

At the outset of the packaging study summarized below, discussions were held with personnel at GRC 

[32], who are conducting preliminary design studies of the Ares V shroud. Inputs from these discussions 

(presentation packages/references of Titan IV shroud designs [33]) and concepts for the flexible ablator 

were accounted for in the packaging presented below. 

Several packaging assumptions are made. First, the barrel of the Ares V shroud will be a right 

circular cylinder of 10 m outer diameter and will extend upward from the base toward the nose with an 

overall shroud length of 30 m. Second, the GRC design allowed an annular zone 1 m thick between the 

payload and the inner diameter of the shroud for acoustic blankets. It was assumed that the HIAD TPS 

could be placed between the inner diameter of the shroud and payload other than the HIAD TPS, serving 

a partial dual-use as the flexible TPS and as an acoustic blanket. Thirdly, based on the FIAT HIAD sizing 

presented in Section 3.5.1, and early trials of making silicone impregnated silica felt flexible TPS, it was 

assumed that the TPS thickness would be 3.8 cm thick and could be folded in a radius of 15.2 cm. 

Figure 46 depicts the first attempt at the HIAD TPS packaging where the 23 m diameter “Apollo” 

flexible ablator (shown in grey) was folded 50 times in a uniform, “umbrella” fashion. The central rigid 

TPS (shown in red) was assumed to be 7.5 m in diameter. For this scheme, the HIAD would be contained 
56

in the annular volume inside the shroud with an inner diameter of 7.5 m and an outer diameter of 9 m. 

The required length of the annular volume would be 10 m. 

Figure 46. First Attempt at HIAD TPS Packaging 

 

Figure 47 depicts an improved packaging scheme, also in an “umbrella” fashion, except there are only 

six folds. The three views shown depict the packaged HIAD wrapped around the payload volume, 

covering the payload as a blanket. The sketch in the upper left corner shows the 23 m diameter “Apollo” 

shape in grey while the star-shaped figure, colored in yellow, depicts the initiation of the folding. Here the 

cylindrical volume for the HIAD stowage has an inner diameter of 8 m and an outer diameter of 9 m. The 

difference between the 10 m outer diameter of the Ares V shroud and the 9 m outer diameter of the HIAD 

stowage volume accounts for shroud structure and additional acoustic blanketing to that of the dual use 

acoustic shielding provided by the flexible ablator. Again, the red object in the drawing depicts the solid 

nose TPS for the HIAD. In this scheme, the HIAD TPS is contained in a cylindrical volume of 8 m inner 

diameter, a 9 m outer diameter and an overall length of 9 m. 

Importantly, these studies indicate that the HIAD TPS can serve, to some extent, a dual use for 

acoustic shielding for launch loads to the payload. They also provide early concepts for HIAD packaging 

that accounts for the allowable bending radius of the flexible ablator. 

It is anticipated that the EDL TPS Flexible Ablator group will be able to more accurately define 

HIAD TPS ablator thicknesses and folding radii during FY10 and FY11. As this information becomes 

available, improved packaging will be forthcoming. In the future, study should be extended to increase 

the understanding of the dual use of HIAD TPS for shielding payload against acoustic loads. Finally, it is 

important that the possibly detrimental effects of acoustic launch loads on flexible TPS should also be 

evaluated and accounted for in future designs. 
57

Figure 47. Second Attempt at Packaging the 23 m “Apollo” HIAD 
58

6 Figures of Merit 

The EDL-SA Figure of Merit (FOM) assessment activity fulfills the requirements of the EDL-SA 

project to assess technologies using a process that has heritage and utilizes the expertise of the EDL 

community. The outcome from this process gives decision-makers a systematic perspective of the tradeoffs for each of the candidate technologies needed to enable an exploration class Mars mission which can 

be used to form a defensible technology development roadmap. An overview of the FOM assessment 

process and the results of the assessment are presented in this section. 

The assessment of the eight EDL architectures and technologies that are identified within the 

architectures was based on evaluation and relative scoring of FOMs. Thirteen FOMs were used from five 

FOM categories: Safety and Mission Success, Performance and Effectiveness, Programmatic Risk, 

Affordability and Life Cycle Cost and Applicability to Other Missions. Each FOM score was derived 

from proxy parameters and discriminating factors based on three classes of data source: Expert opinion, 

simulation and analysis and past studies. Figure 48 shows all of the FOM categories, FOMs and the data 

sources for each category. Detailed definitions of the FOMs and their proxy parameters are given in 

Appendix A. 

Figure 48 FOMs, Sub-FOMs and Sources Used to Evaluate

6.1 Assessment Process 

 The first step in the process includes defining the FOMs, and proxy parameters to be used. An 

external review of these metrics was conducted with solicited input from technology program managers. 

The EDL-SA analysis team developed reference architectures that would be used to assess the 

performance of the technologies integrated into an EDL system. In parallel to the simulation and analysis 
59

tasks, a separate team held a series of TIMs to address non-performance related areas such as risk and 

cost. The data was then used to build a detailed FOM analysis worksheet that was designed to integrate 

the technical and programmatic factors specific to each technology in the assessment. After the 

worksheet based tool was developed, programmatic weights were collected from a variety of project 

managers and decision makers. The results of the survey were used as part of a sensitivity analysis. The 

EDL-SA FOM assessment process is illustrated in Figure 49. 

The Technology Managers Meeting was held in May 2009 with the managers of the relevant technology 

programs: James L. Pittman (ARMD Fundamental Aeronautics Program/Hypersonics Project), Frank Peri 

(ESMD/Exploration Technology Development Program), Samad Hayati (SMD/Mars Technology 

Program) and David Anderson (SMD/In-Space Propulsion Program). A series of FOM Technical 

Interchange Meetings (TIMs) were held the collect data from subject matter experts. Table 25 lists the 

particular TIMs and their participants. Results from the July 2009 Transition Mechanisms TIMs are 

summarized in Appendix B. 

Figure 49. FOM Assessment Process 

Table 25. Contributors to the FOM Assessment 

Technology Program Managers Mtg. (May 2009)

Dave Anderson Samad Hayati Frank Peri Jim Pittman

Transition Mechanisms TIMs (July 2009)

Doug Adams Allen Chen Ian Clark Karl Edquist

Richard French John Gallon Carl Guernsey Christian de Jong

Devin Kipp Ashley Korzun Chris Madsen Tom Rivellini

Miguel San Martin Steve Sell Adam Steltzner Aron Wolf

Chuck Campbell Chirold Epp Steve Hoffman Don Pettit

Mike Tigges
60

Programmatic Cost & Risk TIM (Nov. 2009)

Robin Beck Anthony Calomino Chuck Campbell Neil Cheatwood

Karl Edquist Stephen Hughes Charles Player

Applicability to Other Missions TIM (Nov. 2009)

Dave Anderson John Dankanich Samad Hayati Tom Rivellini

Adam Steltzner Mike Wright

FOM Weighters (Dec. 2009–Jan. 2010)

Neil Cheatwood Doug Craig Robert Dillman Bret Drake

Dana Gould Samad Hayati James Reuther Dale Thomas

Ethiraj Venkatapathy Henry Wright Thomas Zang

Figure 50 shows an example of the process used to evaluate of the Programmatic Risk FOM category. 

The Programmatic Risk FOM category is made up of two FOMs, a) Technology Development Risk and 

b) Programmatic Cost and Schedule Risk. The proxy parameters for the Technology risk FOMs are: 

number of technologies required; TRL’s of each technology and RD3 scores of each technology. The 

proxy parameter for the Programmatic Cost and schedule FOMs is: Level of EDL system design 

maturity/uncertainty for a given technology approach. 

In order to quantify and justify the evaluation of each proxy parameter, the Risk TIM panelists were 

asked to rate the areas on the far right of Figure 50. This information was used to evaluate the proxy 

parameters and establish a relative score for each of the FOMs. The information from the FOMs was 

used to evaluate the FOM categories. A tool was developed to store the information from the TIMs and 

to roll up values from the proxy parameters into relative values for the FOMs. 

Figure 50. Programmatic Risk FOM Process Flow with Supporting Information from Program Risk TIM 
61

6.2 Figure of Merit Weighting 

In the last phase of the assessment process, input from programmatic experts was used to assign 

weights to each FOM and FOM category. A survey was sent to 11 managers, EDL subject matter experts 

and systems engineers that used the Analytical Hierarchy Process (AHP) to do a pairwise comparison of 

each FOM and FOM category. Figure 51 shows the results of the average FOM category weights based 

on the inputs from the surveys. 

Safety and Mission Success, Affordability and Performance Effectiveness are the top three FOM 

categories based on the programmatic inputs collected and each of the three categories had the highest 

weight in at least one survey. The Applicability to other Missions FOM category was assigned the lowest 

weight in all but one survey (where it was ranked second to last). Programmatic risk was consistently 

ranked neither highest nor lowest priority. 

Figure 51. Average FOM Category Weights 

6.3 Assessment Results 

An assessment of both architectures and EDL technologies was performed based on the process 

described above. Both the architecture and technology-based approaches were designed to inform the 

relative benefits of EDL technologies as well as configurations of those technologies. The observations 

and conclusions from each assessment are included below. 

6.3.1 Architecture Assessment Results 

The following is a list of observations and conclusion based on the results of the architecture 

assessment: 

• The architectures that utilize Dual Pulse Rigid TPS (Architecture 1 and 7) ranked high overall 

due to high scores in the safety, performance, and programmatic risk categories. 

• The all propulsion architecture (Architecture 3) was very sensitive to the weights on all of the 

FOM categories. This architecture had very high scores in the safety, performance, and 

programmatic risk category but had a very low score for affordability due to the considerably 

higher mass compared to the other architectures. 

• The architectures that utilize dual pulse ablative inflatable heat shields (Architecture 2, 6 and 8) 

scored relatively low on average. However Architecture 2 and 8 had high scores in affordability 

and applicability to other missions and therefore the overall score of these architectures was very 

sensitive to the weights placed on these two FOM categories. During the sensitivity study 
62

Architecture 2 and 8 ranked highest overall when either affordability or applicability to other 

missions was valued significantly more than the other FOM categories. 

• The architectures that utilize single pulse heat shields (Architecture 4 and 5) were consistently 

ranked low due to low scores in safety (due to the complexity of integrating two heat shields), 

performance, programmatic risk and affordability. 

• The architectures that utilize supersonic retro propulsion (Architecture 1, 2 and 4) scored high in 

safety (due to lower complexity compared to inflatables) but low in affordability (due to higher 

mass) compared supersonic aerodynamic decelerators. 

• The architectures that utilize supersonic inflatable decelerators (Architecture 7 and 8) scored high 

in affordability (due to lower mass) but low in safety (due to dynamic deployment and integration 

complexity) compared to supersonic retro propulsion. 

Figure 52. Architecture Score from FOM Weigth Surveys and Sensitivity Studies 

Figure 52 illustrates the range of total FOM assessment scores for each of the first 8 architectures 

considered as part of the study. The overall FOM score is on the vertical axis. The black bars show the 

range of scores encountered during the sensitivity study (where a high weight was assigned to one FOM 

category and low weights assigned to the other four categories). The gray bars represent the range of 

scores encountered based on the results of the FOM weight surveys and therefore represent results based 

on actual decision makers value systems. The middle tick represents the total score based on the average 

FOM weights. The architectures that have larger vertical bars are more sensitive to the FOM weights. 

Based on the mean FOM weights, Architecture 1 (rigid dual pulse TPS, and SRP) ranked highest overall. 

However, no specific configuration scored higher than all other architectures in either the sensitivity study 

or the FOM weight survey study and therefore no architecture was identified as a clear choice based on 

the information available at the time of the study. 
63

6.3.2 Technology Assessment Results 

For direct assessment of technologies, the same data collected as part of the architecture assessment 

could be applied directly, with the exception of the Performance and Effectiveness Category. For that 

category, technology-based performance metrics were needed. Since the performance data was available 

for all of the architectures an approach was defined such that the technology performance was based on a 

comparison between two similar architectures. Performance metrics for technology evaluation were 

assessed based on comparing reference architectures and their relative performance scores. 

 (a) (b) 

(c) 

Figure 53. Technology Score Ranges from FOM Weight Surveys and Sensitivity Study (a) Supersoinc 

technologies, (b) TPS technologies (c) Hypersonic technologies 

Figure 53(a-c) summarize the results of the technology based assessment. The vertical axis is the 

total integrated score for each technology. The black bars show the range of scores encountered during 

the sensitivity study (where a high weight was assigned to one FOM category and low weights assigned 

to the other four categories). The gray bars represent the range of scores encountered based on the results 

of the FOM weight surveys and therefore represent results based on actual decision makers value systems. 

The middle tick represents the total score based on the average FOM weights. 
64

SRP scored slightly higher on average than the SIAD configuration integrated with a rigid aeroshell. 

However the SIAD integrated with a deployable heat shield showed an advantage over the SRP 

alternative. The dual pulse TPS options scored higher on average than the single pulse alternatives 

primarily due to reduced mass and complexity. The rigid aeroshell scored higher on average than the 

deployable ablative aeroshell due to safety performance and programmatic risk. However, in none of the 

technology areas did a single technology rank consistently highest in either the sensitivity study or the 

FOM weight survey study and therefore no clear set of preferred technologies was identified as part of 

this assessment based on the information available at the time of the study. 

The conclusions of the technology based assessment are the same as the conclusions of the 

architecture based assessment summarized above. 

6.4 Summary 

Figure of Merit assessment was performed on eight exploration-class Mars EDL architectures along 

with assessment of the technologies used in the architectures. Based on relative values of the proxy 

parameters and programmatic weighting of the FOMs, Architectures 1 and 7 consistently ranked high. 

Architecture 3 varied greatly between worst and best architecture due to high sensitivity to FOM weights. 

The score for Architectures 2 and 8 were sensitive to the weights on affordability and applicability to 

other mission FOM categories. 

The technologies required for the highest scoring architectures include: Dual pulse rigid TPS, SRP, 

drag SIAD, lifting SIADs, and lifting HIADs. The same technologies scored high in the technology 

assessment. Due to the sensitivity to decision maker values and the uncertainty in the proxy parameter 

inputs, no specific technologies are recommended for elimination from further consideration as a result of 

this assessment. 
65

7 Viability Assessments 

7.1 Dual-Layer TPS 

Dual heat pulse Proof of Concept or Validation Testing - The objective of this viability testing was 

to demonstrate that a PICA atop LI-900 dual-layer TPS concept could be manufactured in a very large, 10 

x 29 m configuration, and could endure heating rates of more than factor of ten greater than those seen by 

the Shuttle during Earth entries from the Space Station. The Orion TPS ADP developed an approach [34] 

to use large (~ 5 cm) honeycomb attached to the substructure and filled with blocks of TPS material. The 

composition of these TPS material blocks could be changed to address the local heating environments, 

which vary widely across the heat shield. This approach for manufacturing of the dual-layer TPS for the 

10 x 29 m mid-L/D vehicle was adopted because previous arc jet testing sponsored by the Orion TPS 

ADP [27] showed it to be viable for heating rates at levels appropriate for both pulses, aerocapture and 

EDL, predicted for Architecture 1. 

Figure 54 (a) illustrates the large honeycomb TPS manufacturing concept and post arc jet test articles 

where the material filling the honeycomb was PICA only. The PICA is bonded into the honeycomb with 

RTV, a bonding and gap-filling method use on the Shuttle and MSL. The test conditions for Figure 54 

(b) clearly indicate that the thermal performance of PICA in honeycomb is viable at entry environments 

slightly exceeding the peak heating rate values predicted for the aerocapture portion of the mid-L/D 

vehicle shown in Figure 22 above. The same conclusion can be made from Figure 54 (c) for the lower, 

second, entry heating pulse also shown in Figure 22 above. At both the aerocapture and entry peak 

heating conditions, the honeycomb and PICA recede at approximately the same rate, and “fencing” of the 

honeycomb does not seem to be an issue. 

 (a) (b) (c) 

Figure 54. (a) curved panel covered with ~5cm honeycomh filled with PICA blocks; (b) post tested stagnation 

iso-q model after exposure to 620 W/cm2 for 30 seconds; and (c) post tested swept cylinder model after 

exposure at 140 W/cm2 for 25 seconds 

Viability testing of the dual-heat-pulse, dual-layer concept was performed by simulating the 

aerocapture, on-orbit cool-off, and subsequent entry in the ARC arc jet complex. First, the aerocapture is 

simulated in the ARC Interactive Heating Facility (IHF) at 438 W/cm2

 (hot wall) and 28 kPa pressure for 

75 seconds. The charred model was then stored for several months at room temperature to simulate cool 

down in Mars orbit. Lastly, the model is exposed to 153 W/cm2 (hot wall) and 5 kPa for 109 seconds in 

the ARC Aerodynamic Heating Facility (AHF) to simulate the out-of-orbit entry heat pulse and 

understand any failure modes at burn-through into the LI-900. Figure 55 below depicts the iso-q duallayer test article design. The model was designed with an outside ablator layer 1.9 cm thick atop a 1.9 cm 
66

thick layer of LI-900. Five models were made; two had no bonding between the iso-q faces of the PICA 

and the LI-900 and three had a very thin layer of RTV at the boundary of the ablator and the insulator. 

Each model had the LI-900 press fitted against the inner face of the PICA to ensure thermal contact. 

Figure 56 specifies the test thermal cycle sequence that is the best approach currently available to 

simulate the dual-heat-pulse maneuver. The flow in the arc jet test impinges on curved surface of the test 

article (left to right) and the shape of the model tends to make the surface heating constant (iso-q). Two 

IHF tests were run. The first was run for 75 seconds duration to correspond directly to the first pulse. It 

was predicted that the test would leave 0.63 cm of charred PICA by the ablative analysis tool FIAT [25]. 

The actual measured value was 0.61 cm. The second test allowed the model to remain in the stream until 

the PICA receded to the LI-900. Burn-through was predicted by FIAT to occur at 108 seconds while the 

actual burn through was measured at 107 seconds. Figure 57 shows the first post-tested model with 0.61 

cm of charred PICA remaining. 

Figure 55. Iso-Q Model Design of PICA atop LI-900. Dimensions in inches. 

Figure 56. Test Thermal Cycle to Simulate Dual Heat Pulse 
67

Figure 58 shows the burned-through model that was removed from the stream after 107 seconds. As 

can be seen, the LI-900 slumped. This is considered to be a failure mode and shows that for the flight 

design, there must be sufficient PICA remaining after the first pulse to prevent burn-through of the PICA 

during entry from orbit as discussed in the analysis section above. 

A dual-layer model tested at the first pulse IHF condition with RTV bonding between the PICA and 

the LI-900 was stored for several months and then tested for a second heating pulse in the AHF arc jet. 

The second heat pulse level was 153 W/cm2 (hot wall) and 5 KPa. FIAT predicted that it would take 84 

seconds for the PICA to burn-through to the LI-900. The actual time to burn-through was observed to be 

83 seconds. After burn-through the test run continued for an additional 25 seconds, during which time the 

LI-900 slumped. The slumping was gradual, indicating that this failure mode occurs “gracefully.” 

Figure 59 shows clips taken from the movie made during the AHF run at 10, 50, 82, 83, 84 and 93 

seconds, respectively. These views are taken with a mirror at an oblique angle so the glowing model face 

appears elliptical in the figure. The front face of the model undergoes uniform recession from insertion 

until 82 seconds later. At 83 seconds, a brighter glowing is seen at the center of the model, and a second 

later, the PICA has been penetrated. At 84 seconds, one can “see” into the LI-900 region. As time 

progresses, the LI-900 slumps as shown in the clip taken at 

93 seconds. Again the, failure is “graceful” and even the 

spherical dome of PICA did not break off during the 

completion of the run. 

The data clearly shows the viability of building TPS for a 

10 x 29 m mid-L/D aeroshell for a future human Mars 

mission: It has been demonstrated that PICA in a large 

honeycomb is manufacturable, and that thermal performance 

for the dual pulse environment is predictable. The pre-test 

simulations of the arc jet tests demonstrate that the FIAT 

model [25] predicts the thermal performance of the PICA 

atop LI-900 amazingly well for the dual, constant, heat pulse 

arc jet conditions that simulate the real dual-heat-pulse flight 

case. Predictions of recession and time for burn-through for 

both the first and second heat pulses are accurate to within a 

second of those observed. Importantly, charred PICA from 

the first pulse, performed beautifully, as predicted for the 

second, out-of-orbit heat pulse. Four-point positive and 

negative flexural tests of PICA atop LI-900 in ~ 5 cm 

honeycomb showed encouraging results [27] for test articles 

fabricated on an aluminum substrate. 

Two caveats are made here: (1) As shown in work by 

the Orion TPS ADP, PICA recession at heat rates below 50 

W/cm2

 is not well predicted by FIAT as explained above. As 

with Orion, the low heating conditions will be experienced at 

the “tails” of the two flight heat pulses and (2) Fencing of the 

honeycomb at low heat fluxes may be an issue, requiring 

testing and refinement of the phenolic impregnated fiberglass 

honeycomb. The technology for this test series represents the 

first formulation of the new honeycomb technology [27]. 

Figure 58. Post tested model at the IHF 

run condition - Burned-through model 

after and overtest of 107 seconds at the 

IHF condition showing a failure mode of 

the dual layer PICA atop LI-900 TPS 

concept

Figure 57. Post tested model at the IHF 

run condition - Post tested model after 75 

second entry in the IHF
68

Figure 59. : Clips from a movie taken during the second heat pulse simulating entry from out-of-Mars orbit 

after aerocapture 

7.2 IAD Controllability 

HIAD structural mass models were developed for zero angle of attack and zero bank angle rates. In 

order to assess the HIAD concepts, the EDL-SA project initiated a contract with the Computational Fluid 

Dynamic Research Corporation (CFDRC) [47] to perform high-fidelity analysis of several HIAD 

configurations at nominal bank angle maneuver rate of 5 deg/s2

. The primary objective was to perform 

computational aero-structural simulations to support the EDL-SA assessments of IAD controllability and 

structural integrity during EDL. Analyses were performed with various design parameters, and the initial 

results indicate that the Von Mises stress levels on the gores and toroids remained below the material 

yield stresses with a comfortable margin. 
69

8 Technology Investment Recomendations 

The purpose of the first year of the EDL-SA task was to identify the Entry, Descent and Landing 

(EDL) technology investments that NASA needs to make in order to successfully land large payloads at 

Mars for both robotic and human-scale missions. The primary effort was focused on improving models 

of those missions, including mass, TPS, structural, guidance, and propulsion. The models were 

incorporated into aerocapture and EDL simulations to assess the overall effects on performance. 

Following are some key recommendations are made based on the results. A list of the TDP 

recommendations is provided in Table 26. 

The lowest arrival mass is associated with use of deployable/inflatable decelerators. Rigid aeroshells 

offer a more traditional solution, despite a 35% arrival mass penalty compared with an inflatable 

decelerator. Supersonic retro-propulsion reduces the sensitivity to environmental variability compared 

with supersonic/subsonic aerodynamic decelerators. However, both inflatables and supersonic retropropulsion need significant development before concerns about their controllability can be allayed. 

Precision landing requirements have never been demonstrated. Aerocapture is enabling (in the absence of 

NTR) and has never been demonstrated. The all-propulsive architecture warrants further investigation 

despite its 200%+ arrival mass penalty because of its high ratings on the Safety and Mission Success and 

Programmatic Risk FOMs. Moreover, significant engineering work will be needed to ensure reliable 

mechanisms are developed for transitioning between the system configurations during EDL. 

An initial study for robotic-class missions showed that landing 1.5 mt at 0 km MOLA on Mars 

requires extension of current aerodynamic decelerator technology. Within the scope of parachute 

technology, a reefed, supersonic Ringsail parachute is the recommended option. However, a flight 

qualification program is required. 

Table 26. TDP Recommendations 

Technology Area TDP Content

Rigid Aeroshells Tools and processes for generating aero/aerothermal databases & mass 

models; rigid, dual heat-pulse capable TPS; structures; rigid aeroshell 

shapes for aerodynamic performance and controllability

Supersonic Retro-Propulsion Aero-propulsion interaction propulsion for supersonic deceleration—

tools, controls, and configurations. Works for high supersonic initiation 

through touchdown.

Deployable/Inflatable 

Decelerators

Tools & processes for generating aero/aerothermal databases & mass 

models for flexible entry/aerocapture vehicles; flexible HIAD materials, 

including flexible ablators and insulators, flexible aeroshell shapes for 

aerodynamic performance, structural strength and controllability

Precision Landing Sensors, navigation and controls and their integration for precision 

landings with hazard avoidance in atmospheres

All-propulsive System studies of open issues for hypersonic phase and staging 

Aerocapture Development Requirements for an Aerocapture Technology Validation Flight Test 

Supersonic Retro-Propulsion 

Flight Test Program

Flight demonstration (TRL=6) of controllability from initiation to 

simulated touchdown of supersonic retro-propulsion descent system.

Deployable/Inflatable 

Decelerator Flight Test 

Program

Flight demonstration (TRL=6), including controllability of Deployable, 

Inflatable Aerodynamic Decelerator

Aerocapture Flight Test Flight demonstration (TRL=6–7) in upper Earth atmosphere

Parachute Flight Test 

Program

Flight testing of a supersonic Ringsail parachute, including reefing and 

deployment of a large (>21.5 m diameter) parachute at Mach >2.0
70

9 Exploration-Class Study Open Issues 

There is no shortage of additional work that could be done on the Exploration-class study. These can 

be placed in three broad categories—increased fidelity of the models and simulations, additional trade 

studies, and alternative Figures of Merit. 

Clearly, greater fidelity in the component models would reduce the uncertainty in the simulation 

output, but a systematic uncertainty analysis should first be applied to determine which component 

models are responsible for the greatest uncertainty in the results. Even though substantial refinements 

were made in the mass models post-EPR, this is probably the one model whose further refinement would 

most increase the credibility of the results. At the simulation level, the biggest gap is the development of 

realistic flight system configurations, especially for the transition events. 

Of the various additional work that might provide further discrimination between technologies, the 

following were recommended by the EPR: 

• Compare the development costs of a packaging and inflation of a very large HIAD with low 

heating and no thermal protection technology development to a small one with a flexible 

ablative TPS development. 

• Determine the areal density at which the HIAD no longer useful. 

• Develop and document breakpoints for each of the technologies and the associated TRL/ 

development risk at each breakpoint, as a means to help discriminate between options and 

architectural paths. 

• Investigate on-orbit assembly of a rigid aeroshell and rigid deployable concepts to compare to 

the HIAD and slender-body aeroshells. 

• Define L/D breakpoints for hypersonic system. Determine the minimum L/D required for the 

flight systems. Identify advantages of L/D’s in the range of ~0.8 to 1. 

• Consider the use of control surfaces on both the rigid and non-rigid systems. 

• Consider approaches in which the 80 mt are placed on the surface by more than 2 EDL 

systems. Assess how this trade affects technology development risk. Consider options in 

which a single EDL system is developed to land 80/x mt (where x is the # of landings) and 

options in which multiple systems are developed (perhaps one 40 mt lander and either two 20 

mt landers or four 10 mt landers). 

• The proposed 3-g limit may provide information sufficient for a relative comparison of 

selected architectures; however, in future studies consider include higher g-load sensitivity 

analyses in order to assess a more realistic benefit of each decelerator technology.

• Consider a more detailed failure mode based assessment for each technology to inform 

programmatic risk, development cost, and the reliability metrics used in the FOM assessment 

including considerations of margins, performance overlap at staging interfaces and lack of 

sensitivity to component performance uncertainties and increases in payload 

• Consider performance metrics specifically tailored to the function of each technology (i.e. ΔV 

normalized by component mass for decelerators) 

• Develop detailed models of the separation of each drag device from the descent stage. 

• Perform a detailed packing analysis and a subsequent update to the mass models used for 

each technology and subsystem and to consider deorbit from a 500 km circular orbit. 
71

10 Large Robotic-Class Study 

Future robotic missions to Mars are expected to require payload masses on the order of one to two 

metric tons. Thus far, robotic missions to Mars have used heritage from the deceleration technologies 

developed during the Viking era in the 1960s. The payload mass on the Viking landers was 590 kg. Every 

Mars mission after Viking has implemented incremental performance enhancements in aeroshell and 

parachute designs to obtain modest increases in landed mass. However, the Mars Science Laboratory 

(MSL), a mission using an Atlas V to deliver payload of 950 kg pushes the limits of the Viking 

decelerator technology (a 21.5 m parachute deployed at Mach ~2.3). Therefore, future robotic missions 

attempting to land payloads larger than 1 mt on Mars must extrapolate the current decelerator technology 

beyond the tested and validated flight regime and invest in a new, costly technology development 

program. 

The robotic-class system analysis performed in this document focused on delivering 1500 kg to the 

surface of Mars using various parachute designs on an MSL-like entry vehicle. The parachute type, size, 

and number were varied to achieve the desired payload performance. To minimize the parachute design 

modifications from current technology, the MSL entry capsule’s lift-to-drag ratio was increased from 0.24 

to 0.3 by increasing the angle of attack during the entry phase to taking advantage of the high heat tolerant 

properties of the PICA heatshield. Single-stage, two-stage, and unreefed and reefed parachutes were 

explored for two different parachute types, the Disk Gap Band (DGB) and Ringsail. In addition, a systems 

engineering analysis was performed to assess the ability of the current 4.5 m MSL aeroshell to adequately 

package a 1500 kg payload. A complete overview of the ground rules and assumptions for the study can 

be found in Appendix A. 

A survey of past test flights was performed to determine the aerodynamic modeling for each 

parachute design. The current MSL parachute is a 21.5 m diameter DGB but a Ringsail design is known 

to exceed DGB performance in the subsonic flight regime. Although popular for subsonic applications, a 

Ringsail parachute had only a few supersonic fight tests, making the determination of its supersonic 

aerodynamic characteristics difficult. Ultimately, a Ringsail model was adopted for use in this study. A 

reefed parachute is a parachute or cluster of parachutes that opens in stages to improve the probability of a 

successful deployment especially with large diameter designs. Since it was expected that parachutes 

significantly larger than 21.5 m would be required to land a 1500 kg payload, a reefed chute design option 

was selected. 

10.1 Packaging Feasibility 

Mass scaling equations were used to scale the structural elements of the MSL-type configuration for 

each point design. An additional 30% contingency was applied based on JPL design principles. It was 

determined that the packing density resulting from integrating the 1500 kg payload into the MSL 4.5 m

aeroshell was exceedingly large. The only recourse was to increase the aeroshell diameter to 4.7 m, which 

still complied with ground and launch vehicle restrictions. In addition, the number of engines on the MSL 

descent stage would have to increase from 8 to 12 in order to maintain control authority similar to that of 

MSL during the powered descent phase. Finally, an additional 300 kg of propellant (700 kg total) was 

required to land the larger payload mass. The Apollo entry guidance [23] was used to steer the vehicle to 

the landing site. 

10.2 Parachutes 

The parachute type, size, and number were varied to achieve the desired payload performance. Figure 

60 shows each parachute configuration considered for this study: single-stage, two-stage, and reefed 
72

parachutes. Each parachute arrangement was studied for two different parachute types: Disk-Gap-Band 

(DGB) and Ringsail. Models derived from flight test data for the DGB parachute were used for this 

study[37]. 

Figure 60. Parachute systems considered. The systems were implemented using DGB or ringsail parachutes 

(adapted from Ref. NASA TN D-6469). 

10.2.1 Modeling Overview 

This section provides a description of the parachute performance models used for the robotic-class 

mission analysis of the EDL-SA study. The Disk-Gap-Band parachute and a Ringsail parachute were 

evaluated for use in three different implementation strategies. 

Single-stage parachute system -A single, large parachute is deployed at supersonic conditions and used 

to decelerate the entry vehicle to subsonic conditions. This is the parachute concept used by previous 

Mars missions with the exception that larger parachutes (D0 > 21.5 m) are deployed at higher Mach 

numbers (Mach > 2.1) than have previously been employed at Mars.

Two-stage parachute system - A large (~21.5 m) parachute is deployed at supersonic conditions (~ Mach 

2.3) and used to decelerate the vehicle to lower supersonic conditions (~ Mach 1.5). At that point, a 

second, larger parachute of the same type (i.e. DGB or Ringsail) is deployed by the first parachute and 

used to decelerate the vehicle to subsonic conditions. This approach essentially uses the first parachute as 

a very large pilot parachute.

Reefed parachute - A single, large parachute is deployed in a reefed state at supersonic conditions 

(~Mach 2.3). The parachute is disreefed at a lower Mach number and used to decelerate the vehicle to 

subsonic conditions.

Each of the parachute arrangements described above utilize a set of performance models to simulate 

their behavior during deployment and descent. The performance models are of two types: those that 

simulate deployment and those that simulate drag performance. The equations and inputs used for each 

model type are described in subsequent sections. 

10.2.2 Deployment Modeling 

Time to inflation initiation - The deployment sequence is considered as two separate events: 

deployment initiation (t0) to inflation initiation (tSI) and inflation initiation to full inflation (tFI). For the 

case of the single-stage parachute, the reefed parachute, and the first stage of the two-stage parachute, 

deployment initiation corresponds to the moment of mortar fire. For the second stage of the two-stage 

parachute, deployment initiation corresponds to the moment when the first stage parachute begins to 

extract the second parachute. 
73

The time duration between deployment initiation and inflation initiation is estimated as being the time 

required for the parachute bag to travel a given distance d. With the exception of the second stage 

parachute, this time is calculated from an assumed mortar velocity vm as follows: 

 

The distance d was selected to be 10 times the diameter of the entry vehicle, or 47 m. The mortar 

velocity was assumed to be 35 m/sec. These values yield a time from mortar fire to inflation initiation of 

1.343 seconds. 

In the case of the second stage parachute, an estimate of the parachute extraction time is calculated by 

considering the difference in ballistic coefficients between the entry vehicle and the first stage parachute 

as follows:

 

Using the equation above and data from an initial set of Monte Carlo trajectories, estimates of the tSI - 

t0 values were calculated and were seen to fall between approximately 0.6 and 0.85 seconds. The final set 

of trajectories utilized a more conservative value of 1.0 second. 

Time to full inflation - Following inflation initiation the parachute undergoes a growth in area until 

reaching its maximum value of S0. The time required for this to occur, tinf, is estimated using the following 

relation: 

 7 

For DGB parachutes, a value of the inflation factor Kinf of 0.02 sec/m was used. An estimate for the 

Kinf value for a supersonic Ringsail parachute was developed based on flight data from three separate 

tests. Those data are summarized below. 

Table 27. Summary of supersonic ringsail parachute inflation times 

D0 (m) tinf* (sec) Kinf (sec/m) Reference

9.51 0.36 0.03785 [38]

16.6 0.63 0.03795 [39]

12.2 0.75 0.06148 [40]

 *tinf is measured from line stretch to peak load 

Averaging the three calculated values of Kinf shown in Table 27 yields a value of 0.04576 sec/m, 

which was subsequently used during analysis. 

Parachute Area Growth - During the inflation process the parachute is undergoing a rapid growth in 

size. The growth in parachute diameter is modeled as occurring linearly with time, in a manner similar to 

Equation 7. This yields the following relation for the ratio of parachute area to nominal area as a function 

of time. 

8 
74

During initial deployment of a single stage, or first stage parachute the parachute area, S, grows from 

a value of 0 to a value of S0 within time tinf. During initial deployment of a reefed parachute, the parachute 

is considered to inflate to its targeted value of canopy area, SR. Disreefing is set to occur at a fixed time, 

tdr, after mortar deployment. Based on early Monte Carlo runs, a value of 9.55 seconds for tdr, was used. 

Area growth during disreefing is modeled as before, using Equation 8. 

Parachute Peak Loads - The peak loads occurring during parachute inflation are calculated using a 

one-equation method provided in [41]. The equation is as follows: 

 9 

Where q∞ is the dynamic pressure at the start of inflation or disreefing, and Cx is the opening force 

coefficient. For the DGB parachute, a Cx value of 1.456 was used based on models used for the Mars 

Science Laboratory simulations. For the Ringsail parachute, a Cx value of 1.1 was used based on 

examination of the available supersonic flight data and data provided in [41]. In particular, deriving a 

value of Cx from the available supersonic flight data yielded values of 1.0619, 1.0725, and 2.1610 while 

[41] suggested a value of 1.1 based on subsonic flight data. The large Cx value of 2.1610 was discounted 

due to the difficulty in extracting a Cx value from the published data set. 

10.2.3 Parachute Drag Performance 

Modeling of the parachute drag performance was done using two different CD0 vs. Mach performance 

curves for each parachute type, a low drag curve and a high drag curve. The low drag curve was used to 

size the parachute while the high drag curve was used during calculation of the peak opening loads. 

Figure 61. High and Low Drag Coefficient Curves Used for DGB Modeling 

DGB Nominal Drag Performance - Disk-Gap-Band parachutes have been successfully deployed on 

several Mars missions. A relative extensive database of drag versus Mach performance has been acquired 

through sevearal Earth-based flight tests and wind tunnel tests [37] [42]. Thus, there is a strong precedent 

for modeling the performance of a supersonic DGB parachute that was drawn on for the present study. In 

particular, the drag coefficient vs. Mach models used for modeling Mars Science Laboratory trajectories 

was used. The high and low drag performance curves are shown in Figure 61. 

Ringsail Nominal Drag Performance - Modeling of the performance of a supersonic Ringsail 

parachute was considerably more difficult than for the DGB due to the lack of historical data. 

Specifically, only four successful supersonic flight tests (out of five attempts) of a Ringsail parachute 
75

have taken place and those are over a limited range in Mach numbers. A summary of the deployment 

conditions for the flight tests is provided in Table 28. A summary of the available flight test data is shown 

in Figure 62. Also provided in Figure 62 are explanations of many of the phenomena observed in the 

data, as they are described in the references. From this data, high and low drag performance curves were 

established. These curves are shown in Figure 63. The shape of the curves was designed to roughly 

mimic that of the DGB curves in Figure 61. Specifically, the inclusion of a decrease in the drag 

performance around a Mach number of 1.0 was based on a similar trend for DGB parachutes. For the low 

drag curve, the supersonic performance was based on the data from [40]. It is worth noting that the 

Ringsail configuration used during all 5 supersonic tests is somewhat different than that typically used for 

subsonic applications, and as a result the subsonic performance was quite low [43]. Therefore, it was 

assumed that a modern day application of a supersonic Ringsail parachute would employ considerable 

modifications to the design used in prior testing. Thus, the supersonic performance seen in [40] was 

considered to be a lower bound. The subsonic performance seen in the four successful flight tests was 

discounted and the lower drag model instead used a more typical Ringsail subsonic drag coefficient of 

0.85. The higher drag model assumed an approximately two times increase in drag coefficient at higher 

Mach numbers. 

Table 28. Summary of Supersonic Ringsail Test Conditions 

D0 (m) Deploy 

Mach

Deploy 

q∞ (Pa) Comments Reference

26.0 1.15 280 Deployed in reefed state with large fluctuations, disreefed 

at ~Mach 0.8, initially inflated behind blunt body

[44]

[45]

9.51 1.39 527 Slight flutter in skirt after opening, minor damage during 

deployment

[38]

[45]

12.2 1.64 436 Full inflation never achieved, failure attributed to excessive 

geometric porosity of crown

[46]

[45]

16.6 1.61 555 Minor damage due to drogue bag recontact, initially 

inflated behind blunt body 

[39]

[45]

12.2 2.95 440 Large areal oscillations until ~Mach 1.5, minor canopy 

damage due to parachute bag and mortar lid recontact [40]
76

Figure 62. Summary of supersonic ringsail flight test drag data. Note: Caption refers to NASA technical 

document references and whether the data was derived from accelerometer or tensiometer measurements.

Figure 63. Summary of supersonic ringsail flight test data along with high and low drag performance models. 

Note: Label text refers to NASA technical document numbers from which the data originated. 

Reefed Parachute Performance - Performance of reefed parachutes used the same drag curves as the 

unreefed parachutes. However, the DGB parachute was scaled by a factor of 0.42, while the Ringsail used 

a scale factor of 0.784. 
77

10.3 Monte Carlo Simulation Results 

The parachute solutions in Table 29 were computed using 8000 Monte Carlo simulations. Dispersions 

in aerodynamics, atmospheric density, winds as well as attitude error and off-nominal entry conditions 

were introduced in the 8000 simulations. A description of the dispersions can be found in Appendix A. 

The configuration in Table 29 is considered feasible if the timeline margin is greater than 15 s. The 

timeline margin is the time between radar acquisition and descent stage ignition at two km above ground 

level (10s of radar operation to guarantee altitude solution plus 5 s for engine priming). The diameters 

shown in Table 29 are the minimum diameters required to obtain feasible solutions (i.e., solutions with 

timeline margin >15s) in the Monte Carlo simulations. Figure 64 shows data from a Monte Carlo 

simulation using the reefed parachute configuration. Note how the 3σ high for parachute deploy Mach is 

less than 2.5. 

Table 29. Supersonic parachute performance results. 

Parachute deploy Mach numbers correspond to 3σ high values 

Parachute DGB Ringsail

Single Parachute

Diameter (m) 32.5 31.5

Deploy Mach 2.5 2.5

Two-stage Parachute

1st Diameter (m) 2nd Diameter (m) 21.5 41.5 21.5 41.0

1st Deploy Mach 2nd Deploy Mach 2.5 1.5 2.5 1.5

Reefed Parachute

Reefed Diameter (m) Disreefed Diameter (m) 21.5 >34.0 21.5 34.0

Reefed Deploy Mach Disreefed Deploy Mach 2.5 2.0 2.5 2.0

In Table 29, the single and two-stage parachute solutions have lower probability of success due to the 

large loads. The single parachute configuration is deployed at Mach ≤ 2.5, and the diameters needed to 

land 1500 kg are 32.5 m and 31.5 m for the DGB and Ringsail parachute types respectively. The diameter 

of the first parachute in the two-stage parachute configuration is held constant to 21.5 m (i.e., to maintain 

MSL heritage) and is deployed at Mach ≤ 2.5. The second chute in the two-stage parachute configuration 

is deployed at Mach ≤ 1.5 and the diameters of the second chutes are 41.5 m and 41 m for DGB and 

Ringsail respectively. As seen in Table 29, the diameter for a reefed Ringsail parachute configuration is 

34 m. The DGB reefed parachute diameter will be larger than 34 m since the aerodynamic performance of 

a reefed DGB parachute is worse than that of a reefed Ringsail parachute. 

All parachute solutions in Table 29 are capable of landing 1500 to the surface of Mars using a scaled 

MSL-type vehicle. Given the certain investment in a flight test, the recommended technology that shows 

the most promise for success and ultimate capability is a single 34 m reefed supersonic Ringsail. Reefing 

adds an element of stability, which enables larger diameters over an unreefed parachute. Other notable 

advantages of using the reefed Ringsail parachute include reduced mass and complexity while increasing 

flexibility to tailor the drag deceleration as needed. Also the 34 m reefed supersonic ringsail is considered 

to be at TRL-4 today. 
78

Specific parameters for the recommended reefed parachute configuration are shown in Table 30 

below. Also, Table 31 compares this reefed chute configuration with MSL. 

Figure 64. Monte Carlo simulation results using 8000 dispersed cases for the reefed parachute configuration. 

Table 30. Supersonic reefed parachute performance results. 

Chute deploy Mach numbers correspond to 3σ high values.

Parameter Value Comments

Entry Mass 5604 kg ~2 mt heavier than MSL

Entry Lift-to-Drag Ratio 0.3 Heating, stability and packaging issues still need 

to be analyzed

Peak g-load 13 g’s MSL typically ≤ 15g’s

Peak Heat Load 9600 J/cm2 MSL typically ≤6200 J/cm2

Parachute Deployment Mach # 

(nominal; 3-sigma high = ~2.5) 2.28 No new technology but flight test required

Reefed Diameter 21.5 m No new technology but flight test required

Disreef Mach # 1.8 (2.0) No new technology but flight test required
79

nominal (3-sigma high)

Disreefed Diameter 34 m No new technology but flight test required

Number of Descent Engines 12 4 additional engines over MSL. More engines 

required for 2≤T/W≤3

Engine Characteristics 3046.9 N/eng

220s Isp

MSL engine characteristics

Time Between Radar 

Acquisition and Ignition

15.6 s Meets 15s minimum requirement

Total Propellant Load 624.0 kg 225 kg additional propellant over MSL

 

Table 31. Comparison of MSL and the scaled reefed ringsail configuration. 

Sub-System MSL* Scaled Reefed Ringsail 

Configuration**

Entry Mass 3350 kg 5600 kg

Aeroshell Diameter 4.5 m 4.7 m

Parachute DGB 21.5 m Reefed Ringsail 21.5 m / 34 m

Parachute Deploy Mach ≤2.3 ≤2.5

Lift-to-drag ratio 0.24 0.30

Number of MLEs 8 12

Propellant 400 kg 625 kg

Time on Radar >20 s >15 s

Rover Mass 900 kg 1500 kg

Launch Vehicle Atlas V (541) Delta IV-Heavy

* Masses of MSL obtained from MEL 08/15/2008

**Applied a post-scaling factor of 30% to all sub-systems since design is appreciably 

different from MSL

10.4 Robotic Class Summary 

Single, two-stage and reefed parachutes capable of landing 1.5 mt at 0 km MOLA were explored in 

this study. All parachute solutions require supersonic high altitude BLDT (Balloon Launched Decelerator 

Test). The Monte Carlo results in this document indicate that a single 34 m reefed supersonic Ringsail is 

the most promising configuration. 
80

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Optimize Simulated Trajectories (POST2), Vol. II Utilization Manual.” Version 1.1.1G, May 2000. 

29. Justus, C. G. and Johnson, D. L., “Mars Global Reference Atmospheric Model 2001 Version (MarsGRAM 2001): Users Guide” NASA TM 2001-210961 April 2001 

30. Kumar, "Issues on Human Acceleration Tolerance After Long-Duration Space Flights" NASA 

Technical Memo 104753, Oct 1992. 

31. Finley, P. J., "The Flow of a Jet from Body Opposing a Supersonic Free Stream," Journal of Fluid 

Mechanics, Vol. 26, No. 2, 1966, pp. 337-368.

32. Kirvanek, T. and Lepsch, R., GRC Ares V Shroud Studies, September 22, 2009. (Unpublished) 

33. Hughes, W.O., and McNelis, “Acoustic Testing of the Cassini Spacecraft and Titan IV Payload 

Faring”, NASA TM 107474, 67th Shock and Vibration Conference, Monterey, CA. November, 18-22, 

1996.

34. Zell, P. T., “The Block-Ablator in a Honeycomb Heatshield Architecture,” International Planetary 

Probe Workshop, Barcelona Spain, June 12-18, 2010. 

35. Personal communication, M. Blosser and C. Poteet, LaRC. (Unpublished) 

36. Mendeck, G. F. and Carman, G. L., “Guidance Design for Mars Smart Landers Using The Entry 

Terminal Point Controller,” AIAA 2002-4502 Atmospheric Flight Mechanics Conference and 

Exhibit, Monterey, CA, August 2002. 

37. Bendura et al., “Flight Tests of Viking Parachute System in Three Mach Number Regimes II: 

Parachute Test Results,” NASA Technical Note TN D-7734, 1974. 

38. Preisser et al., “Flight Test of a 31.2-Foot-Diameter Modified Ringsail Parachute Deployed at a Mach 

Number of 1.39 and a Dynamic Pressure of 11.0 Pounds Per Square Foot,” NASA Technical 

Memorandum TM X-1414, 1967. 
82

39. Whitlock et al. “Performance of a 16.6-Meter-Diameter Modified Ringsail Parachute in a Simulated 

Martian Environment,” NASA Technical Memorandum TM X-1500, 1968. 

40. Eckstrom, C. V., “High-Altitude Flight Test of a 40-Foot-Diameter (12.2-Meter) Ringsail Parachute at 

a Deployment Mach Number of 2.95,” NASA Technical Note TN D-5796, 1970. 

41. Knacke, T. W., “Parachute Recovery Systems Design Manual,” Naval Weapons Center Technical 

Publication NWC TP 6575, 1991. 

42. Steinberg et al., “Development of the Viking Parachute Configuration by Wind Tunnel Test,” AIAA 

Paper 1973-454, 1973. 

43. Ewing, E. G., “Ringsail Parachute Design,” Air Force Flight Dynamics Laboratory Technical Report 

AFFDL-TR-72-3, 1972. 

44. Whitlock et al., “Performance of a 26-Meter-Diameter Ringsail Parachute in a Simulated Martian 

Environment,” NASA Technical Memorandum TM X-1356, 1967. 

45. Whitlock, C. H. and Bendura, R. J., “Inflation and Performance of Three Parachute Configurations 

From Supersonic Flight Tests in a Low-Density Environment,” NASA Technical Note TN D-5296, 

1969.

46. Eckstrom et al., “Flight Test of a 40-Foot-Nominal-Diameter Modified Ringsail Parachute Deployed 

at a Mach Number of 1.64 and a Dynamic Pressure of 9.1 Pounds Per Square Foot,” NASA Technical 

Memorandum TM X-1484, 1967. 

47. Sheta, E., Venugopalan, V., Tan, XG, Liever, P. and Habchi, S., “Aero-Structural Assessment of an 

Inflatable Aerodynamic Decelerator,” CFDRC Report: 8927/6, April 2010 (CFD Research 

Corporation, Huntsville, AL) 
83

A Appendix A. DRMs, GR&As and FOMs 

A.1 Design Reference Missions 

A Design Reference Mission (DRM) describes the objectives and top-level requirements for a 

representative future mission for the purposes of facilitating choices of architectures and technologies for 

a class of missions. The EDL-SA Study encompasses two such classes—Exploration Class missions and 

Large-Robotic Class missions. This section provides a summary of the respective DRMs for these two 

mission classes that will be used in the EDL-SA Study. 

A.1.1 Exploration-Class Design Reference Mission 

Figure A-1 illustrates the Exploration-class DRM used in the 2007 Mars Design Reference 

Architecture 5 study (“Human Exploration of Mars Design Reference Architecture 5.0”). For this 

architecture, pre-deployed assets are used—two large cargo vehicles arrive first, and are placed in orbit 

using propulsion or aerocapture. The vehicle carrying the Mars ascent vehicle undergoes EDL shortly 

thereafter, while the vehicle carrying the habitat lander remains in orbit. A crewed system arrives 2 years 

later, is propulsively captured into orbit, from which it rendezvous with the orbiting habitat lander. The 

habitat lander with the crew then undergoes EDL. This not only reduces the single EDL mass size, but 

also reduces the mission risk. 

Figure A-1. Exploration-Class Design Reference Missions 

Some of the key parameters of the systems for the overall architecture are provided in Table A-1 for 

both nuclear thermal rocket (NTR) and chemical propulsion options. 
84

Table A-1.Key Parameters for the Exploration-Class DRM 

Parameter NTR Reference Chemical Option

Total Crew Flight Duration (approx. days) * ~900 ~900

Crew Transit time LEO-Mars (approx. days) ~180 ~180

Crew Mars Stay Time (approx. days) ~540 ~540

Crew Transit time Mars-Earth (approx. days) ~180 ~180

Total Initial MTV Mass in LEO (IMLEO) (t) ** 825 1252

Crew Vehicle Mass 333 534

Inter-Planetary Transportation (t) 282 483

Crew Transit Payload (t) 51 51

Cargo Vehicle Mass (mt each) 246 359

Inter-Planetary Transportation (t) 144 257

Mars Surface Payload (t) 36 36

Propulsive Lander (wet, t) 23 23

Aeroshell Mass (t) 43 43

Launch Data †

Ares-I Launches (crew) 1 1

Ares-V Launches (cargo) 7-9 10-12

Launch Campaign Duration (days) 300 390

* Trip times are average durations across the synodic cycle

** All mass data exclusive of Project and Program reserves

† Number of launches dependent on launch vehicle selected

Figure A-2 illustrates the 3 Mars Transfer Vehicles (MTVs) needed for the NTR option. The two 

cargo landers embarking in the first landing cycle require two Ares V launches each. The crewed lander 

embarks on the second cycle; it requires three Ares V launches and one Ares I launch. For purposes of the 

EDL-SA Study, the Exploration-class DRM is the EDL portion (including the aerocapture phases) of that 

DRM (modified, as discussed below). 

The aeroshell concept used in DRA5 for the Mars landers was a dual-use Ares V shroud, referred to 

as an ellipsled. This is illustrated in Figure A-3. Further details on this DRM are available in the DRA 5 

report (“Human Exploration of Mars Design Reference Architecture 5.0 Executive Summary”). 

The particular DRM employed in the EDL-SA study is similar. The primary difference is the use of a 

more conservative margin policy. The details are given in Section A.2.4. 
85

Figure A-2. Mars Transfer Vehicles 

Figure A-3. Outer Mold Line of a Mars Lander Aeroshell 

A.1.2 A.1.2 Large-Robotic Class Reference Mission 

The Large-Robotic-Class DRM to be used for the FY 09 robotic study has been designed to land an 

equivalent 1.5 metric ton payload on the surface of Mars using a scaled Mars Science Laboratory (MSL) 

configuration (see Figure A-4). A 1.5 metric ton payload configuration accommodates all the currently 

planned robotic missions to Mars within the next 20 years as well as utilizes technology that could be 

applied to future manned missions to Mars. This feed forward technology is embedded in the performance 

enhancements to the MSL design that enable more deceleration forces to be applied to the vehicle at 

lower mach numbers. This would include, but not be limited to, Inflatable Aerodynamic Decelerators 
86

(IAD) and supersonic retro-propulsion, both of which are scalable to much larger configurations. 

Figure A-4. Large-Robotic-Class Design Reference Mission.

It is of high interest for near term robotic missions to have the ability to deliver the payload to the 

surface with very high landing accuracy: within 100 m of the intended target. Pin-Point Landing (PPL) 

during powered descent is employed as the next logical state-of-the-art step to achieve this level accuracy, 

amongst other navigational enhancements (see Figure A-5). Note that PPL will be studied in FY10 and 

beyond in favor of a simple gravity turn in FY09 for ease of establishing an integrated analysis process 

across multiple centers. 

The landed accuracy of the Large-Robotic-Class DRM is compared with that of MSL in Figure A-5 It 

is expected that an additional 400 kg of propellant will be required to “fly-out” typical delivery errors at 

ignition of the powered descent stage. This extra propellant combined with the 1.1 metric ton of landed 

mass achieves the desired 1.5 metric ton equivalent capability. Note for FY09, a landed mass of 1.5 

metric ton will be delivered to the surface using a gravity turn only. 

Note for FY09, a landed mass of 1.5 metric ton will be delivered to the surface using a gravity turn 

only. To further challenge the system, the DRM will assume a landing at a relative high altitude of 0 km 

MOLA during an arrival opportunity with unfavorably low atmospheric density. Unfavorable 

atmospheres are seasonally related and it has been determined that an Earth departure date in 2022 would 

provide sufficiently stressful Mars arrival atmospheric conditions. In addition, 2022 demands much 

higher Earth departure energy, or C3, that will also put a challenge on the assigned launch vehicle. Table 

1 provides a summary of the general assumptions that will be utilized for the robotic DRM. 

The similarities of this DRM to MSL include (1) a 4.7 m, 70 deg cone angle aeroshell, (2) the Apollo 

entry guidance, (3) the Sky Crane architecture, and (4) a Descent stage propulsion system. Some of the 

differences are (1) a pin-point landing strategy, (2) a three-axis stabilized cruise stage for improved 

attitude uncertainty (cruise balance mass removed), (3) a velocity trigger chute deploy strategy, (4) TPS 

thickness scaled up to account for increased heating, (5) propellant mass and tanks scaled for increased 
87

propellant, (6) more engines added to the Skycrane configuration, and (7) remaining structure scaled as a 

function of acceleration loads and increased payload mass. 

Figure A-5. MSL and Large-Robotic-Class Mission Landing Strategies

Table A-2. Key parameters of the large-robotic-class DRM. 

Assumption Value Comments

Configuration 4.7 m aeroshell

12 MLE’s

700 kg prop load

Scaled MSL design: 4.7 m, 70 deg cone 

aeroshell, more MLE’s deeded to maintain T/W, 

more propellant needed to all a direct 

consequence of 1.5 t payload

Entry Mass 5400 - 5600 kg Dependent on scaled configuration.

Entry Altitude 125 km

Entry Speed 5.6 km/s Dependant on arrival trajectory (direct descent 

assumed).

Entry Flight Path Angle -15.5 deg Defined to maximize altitude at heat shield 

jettison. Optimal is defined as optimal balance 

between system impacts (e.g. heat shield 

thickness) and altitude at heat shield jettison.

Entry Guidance Apollo MSL Apollo guidance. No pin-point landing; 

MSL gravity turn strategy

Opportunity 2022 Most challenging Earth departure C3 & Martian 

Atmosphere (Type I/II transfer assumed).

Site Latitude 0 deg Near equatorial

Site Altitude 0 km MOLA

Launch Vehicle Delta IV - Heavy Atlas V 551 most likely not capable.
88

A.2 EDL-SA Study Ground Rules and Assumptions 

This section lists the Ground Rules and Assumptions (GR&As) employed in the EDL-SA Study. 

A.2.1 General Ground Rules and Assumptions 

* Margins will be applied to all technologies and systems as described in Section A.2.4. 

* Subsystem performance parameters, e.g., engine Isp, engine T/W, vehicle inert mass fraction, will 

be based upon historical data and trends. 

* The atmosphere model used for this study will be MarsGRAM 2005. 

* Turbulent flow onset will be estimated using the Reθ = 200 criterion. 

* Representative guidance algorithms will be developed. Representative guidance algorithms will 

be used for final FOM evaluations. 

* POST2 will be used for simulations. 

* The nominal landing site altitude will be 0 km MOLA. 

* The global landing precision will be within 50 m of the desired touchdown point. 

* Landing site altitude sensitivities will be evaluated for -1 km. MOLA to 2.5 km. MOLA. 

* Peak heating and total heat load will be determined using methods provided by the TPS 

designers. 

* Hazard avoidance technology will be provided by the ALHAT project. Hazards could be predeployed assets or natural hazards. 

* The thermal protection system is uncompromised at Mars arrival. 

A.2.2 Exploration Mission Ground Rules and Assumptions 

* The shed drag device must never pass any closer than 2 km. to the target at any point in its 

trajectory (“keep-out zone” criterion). 

* The Ares V shroud will be capable of launching a 10 m diameter by 33 m length Mars entry 

aeroshell 

* Separation events will be modeled as a 10-sec. interval with no deceleration. 

* Simulations will include a method (e.g., divert maneuver, fly away) to ensure the “keep out zone” 

criterion is met. 

* Crewed and Cargo entries will be made from Mars orbit. 

* Aerocapture will be used to insert the cargo vehicles in the desired parking orbit. One sol and 500 

km. circular orbits will be used to provide sensitivities. 

* Deceleration limits for crewed entry will use the relevant Entry limits from CxP 70024, “Human 

Systems Integration Requirements.” 

* The nominal landed payload mass will be 40 mT. 

* Payload mass sensitivities will be evaluated for 20–50 mT. 

* First human landing will be no later than 2033. 

* Only those technologies capable of being at or above TRL 6 by 2025 will be considered. 

* First human landing will be no later than 2033. Technology Development Roadmaps will 

establish TRL 6 capability by 2025. 

A.2.3 Robotic Mission Ground Rules and Assumptions 

* The nominal landed payload mass will be 1.5 mT. 

* Mars opportunity will be 2022. 

* Interplanetary mission design will be direct to entry (e.g. no aero/propulsion capture into a Mars 

parking orbit). 

* Entry vehicle will be an MSL, 4.7 m entry vehicle with scaled elements as necessary for g-load, 

heat load, chute inflation loads, payload mass, etc. 
89

* Entry phase will utilize Apollo guidance through a velocity trigger chute deploy. 

* Descent & landing phase mission design will be an MSL descent stage with skycrane. 

* Landing accuracy: Simple Gravity Turn: <=10 km from the target landing site. 

* Landing accuracy: Pin-Point Landing: <= 100 m from the target landing site. 

* Capability will be assessed to a 0 km MOLA altitude landing site at 0 deg latitude. 

* Launch vehicle will be a Delta-IV heavy class or smaller. 

A.2.4 Margin Policy 

The design margins utilized in most previous Mars architecture/EDL studies have been quite 

optimistic. Indeed, some studies have carried no margin at all, and others have applied margins only to 

mass. The EDL-SA Study will use an integrated margin approach that is derived from current practice in 

the early phases of flight projects, e.g., development projects in the Constellation Program and robotic 

missions in the Mars Program. This margin approach will provide an appropriate level of realism to the 

study results. 

Margin is composed of both consumable margin and design margin. Consumable margin is margin 

that exists in excess of the current estimate. Design margins provide design requirements that increase the 

robustness of the design; it is tied to subsystem requirements, generally performance parameters—

structural integrity, thermal limits, TPS thickness, etc. This type of margin is not consumed during the 

course of the project. 

In this study, the consumable margin is defined by 

 Margin = Allocation - Current Best Estimate. 

This is sometimes given in percentage terms: 

Allocation is the amount of a resource that has been allotted to a subsystem, and Current Best 

Estimate is the amount of a resource that is available or needed without any additional contingency. 

Targets for consumable margin are based on the life-cycle stage of the design and the maturity of the 

subsystem. This type of margin is released through the life of the project. Table A-3 defines the 

subsystem quantities to which margin is applied and the amount of margin that is utilized in the EDL-SA 

Study. 
90

Table A-3. EDL-SA Subsystem Margins 

Parameter Description Value

Dry Mass Each configuration will have a minimum Dry Mass Margin. The 

minimum dry mass margin is applied across the complete 

operational scenario of each configuration. Note that technologies 

with a TRL <6 will have a Dry Mass Margin increment added -

listed elsewhere.

30%

Propellant Mass Propellant mass margin is used to account for the margin in Delta V 

and overall margin in propellant mass. The propellant mass margin 

value is derived assuming a 20% DV margin (equates to about a 

15% propellant margin) plus a direct 20% propellant mass margin. 

All tankage will be sized based on the fully margined propellant 

mass, with the fully margined dry mass. Propellant mass margin 

does not include any ullage or unusable volume assumptions.

35%

Thrust All integrated thrust modules will have a minimum thrust margin 

applied after initial sizing. Note that the thrust shall be sized based 

on the fully margined mass.

25%

Skipout The entry flight path angle will maintain a minimum margin from 

skipout.

≥1 degree

Structural Integrity Using appropriate structural factors of safety (defined elsewhere), 

each element will maintain a positive margin greater than the 

specified value.

≥ 0.2

TPS Thickness All TPS (ablative or insulative) will have a minimum thickness 

margin applied. Note that the thickness margin is in addition to the 

Dry Mass margin.

30%

Inflation Mass Where an inflatable decelerator is used, and an on-board gas supply 

or gas generator is used, then the mass available for inflation will 

have a minimum gas supply mass margin applied. If the system uses 

a continuous monitoring and fill system, then the margin is applied 

for the complete gas mass consumed. Does not apply to parachutes.

30%

For TPS the margin policy mirrors that used for Orion TPS development. See “Mars EDL/SA 

Thermal Protection System (TPS) Margin Management Plan” for the details. 

Since this study assesses the potential benefits of technologies for landing large payloads on the 

surface of Mars, the lack of maturity of newer technologies needs to be reflected in the analysis. Hence, 

this study will report the simulation results both with and without a “TRL Increment”—the lower the 

TRL, the higher the increment—to guard against over-optimism in the forecasts of immature 

technologies. The TRL-based biasing of the margins is a penalty that is added to the baseline margin 

value given in Table A-4. In addition to this additional mass margin for low TRL technologies, Monte 

Carlo simulation variables (section 4.5) will be assessed and on a case-by-case basis, a wider variability 

(indicative of performance uncertainty) will be applied to some parameters. An example would be 

increased aerodynamic uncertainties for systems that do not have established aerodynamic databases. 

These penalties will be documented in the Peer Review charts and Final Report, and the results will 

manifest themselves in the proxy variables for the Figures of Merit. 
91

Table A-4. TRL Margin Increments 

TRL No. Mass Margin Increment

9 0

8 0

7 0

6 0

5 20%

4 25%

3 30%

2 40%

1 50%

A.2.5 Exploration Mission Monte Carlo Parameters 

The simulations underlying the system analyses performed in this study utilize Monte Carlo 

procedures to account for various uncertainties in the system and its environment. The uncertainties 

include both aleatory (variability) and epistemic (lack of knowledge) ones. We adopt the customary 

approach of describing the uncertainties by probability density functions (PDFs). The PDFs represent the 

degree of belief about the distribution of the uncertain quantities—not only the nominal (most likely) 

values, but also their distribution between their plausible upper and lower limits. Of course, these 

subjective choices are informed by the statistical data that does exist and the team’s experience with 

Monte Carlo simulations of previous Mars missions (MER, Phoenix, MSL). Conservative adjustments 

were made because of lack of detailed information (experimental data, high-fidelity analyses) on current 

concepts. 

Table A-5 and Table A-6 list the parameters that are treated probabilistically in the ellipsled and 

inflatable (MIAS) simulations, respectively. For each uncertain parameter, the rationale for the choice of 

distribution is included. The aerodynamic databases used in the simulations are based upon engineeringlevel tools. The uncertainties in these data are modeled by multipliers on the aerodynamic coefficients. 

The model for the uncertainty in the Mars atmosphere is the model contained in MARSGRAM; the range 

for the random number seed is provided in the tables. Atmospheric density perturbations are applied to all 

cases except the nominal, and winds are applied to all cases except nominal. For now, the de-orbit DV 

considers only maneuver execution errors. As the EDL-SA Study matures, more variables will be treated 

probabilistically in the simulations. For example, once more accurate entry states are obtained; 

perturbations to them will be included. 

Table A-5. Mid-L/D Rigid Monte Carlo Parameters 

Parameter Nominal Perturbation Distribution Rationale

Mass Variation (kg) 0 ±10 Normal Not currently used. Sensitivity to 

ballistic coefficient needs to be 

quantified as the study matures – i.e. 

those technologies with less 

sensitivity would be favored over 

those with a high sensitivity. (also 

this sensitivity will be nonlinear in 

many instances). To minimize the 

number of Monte Carlo scripts that 

must be certified, this capability has 

already been added – but will not be 

exercised until later in the study.
92

Lift Coefficient 

Multiplier 

1 ±10 % Normal Larger than would be expected to 

provide sensitivity analysis. The 

sensitivity characteristics of the 

technology are important 

considerations. Also the 

aerodynamics are uncorrelated, to 

further magnify the sensitivities. 

Many technologies will look good at 

their design point – one major 

discriminator will be sensitivity.

Drag Coefficient 

multiplier

1 ±10 % Normal See Lift Coefficient Multiplier 

discussion.

Angle of Attack 

(deg)

55 ±5 Normal Angle of attack will vary the 

ballistic coefficient, and will 

amplify the ratio of Mass/CL to 

Mass/CD, to provide sensitivity 

analysis

Engine Isp (sec) 369 ±2.5 % Normal This was an estimate for LOX/CH4 

engines which have not been built. 

This parameter is expected to be a 

secondary sensitivity parameter for 

most of the technologies to be 

considered. If it becomes a primary 

driver, this will be reevaluated.

Deorbit ΔV (m/s) 14.978167 ±0.279* Normal In lieu of having initial states 

generated for this study (time and 

resources that would not be a driver 

in selecting technologies), this 

parameter is used as an analog to 

provide the entry g’s and heat pulse 

variation that would be expected 

from this type of mission.

Atm Random # 1 1-29999 Integer Uniform In lieu of having initial states 

generated for this study (time and 

resources that would not be a driver 

in selecting technologies), this 

parameter is used as an analog to 

provide the entry g’s and heat pulse 

variation that would be expected 

from this type of mission.

Dusttau 0.7 0.1:0.9 Uniform This determines the dust loading and 

thus the density and wind profiles 

that the vehicle will experience. 

This range provides large 

variability, but would not include 

dust storms.

Notes

* ΔV perturbation selected to produce a +/- 0.25 deg variation in entry Flight Path Angle

Preliminary Lift and Drag Coefficient multipliers act on aerodynamic tables provided by Dr. Alan 

Wilhite for the ellipsled shape. The tables were generated using a program called APAS. For the high 

supersonic to hypersonic region, a modified Newtonian technique is used to calculate the pressure 

distribution on surfaces as a function of the incident angle to the flow. The pressures are integrated to get 

the aerodynamic coefficients. However, it cannot accurately predict trim angles (generally errors are on 
93

order of 4 to 5 deg), but the approximations are sufficient for analysis of the current fidelity. Note that 

this method will not work once packaging requirements are imposed. Aerodynamic data provided in the 

MIAS reports is used for the inflatable configuration. 

Table A-6. Low-L/D Inflatable Monte Carlo Parameters 

Parameter Nominal Perturbation Distribution Rationale

Mass Variation (kg) 0 +/- 10 Normal Not currently used – Sensitivity to 

ballistic coefficient needs to be 

quantified as the study matures – i.e. 

those technologies with less 

sensitivity would be favored over 

those with a high sensitivity. (also 

this sensitivity will be nonlinear in 

many instances). To minimize the 

number of Monte Carlo scripts that 

must be certified, this capability has 

already been added – but will not be 

exercised until later in the study.

Lift Coefficient 

Multiplier 

-1** +/- 10 % Normal Larger than would be expected to 

provide sensitivity analysis. The 

sensitivity characteristics of the 

technology are important 

considerations. Also the 

aerodynamics are uncorrelated, to 

further magnify the sensitivities. 

Many technologies will look good 

at their design point – one major 

discriminator will be sensitivity.

Drag Coefficient 

multiplier

1 +/- 10 % Normal See Lift Coefficient Multiplier 

discussion.

Angle of Attack 

(deg)

27.5 +/- 2 Normal Angle of attack will vary the 

ballistic coefficient, and will 

amplify the ratio of Mass/CL to 

Mass/CD, to provide sensitivity 

analysis.

Engine Isp (sec) 369 +/- 2.5 % Normal This was an estimate for LOX/CH4 

engines which have not been built. 

This parameter is expected to be a 

secondary sensitivity parameter for 

most of the technologies to be 

considered. If it becomes a primary 

driver, this will be reevaluated.

Deorbit ΔV (m/s) 14.0 +/- 0.279* Normal In lieu of having initial states 

generated for this study (time and 

resources that would not be a driver 

in selecting technologies), this 

parameter is used as an analog to 

provide the entry g’s and heat pulse 

variation that would be expected 

from this type of mission.

Atm Random # 1 1-29999 Integer Uniform This is used by the atmosphere 

program to determine the variability 

in the density and wind profiles. 

The range is that allowed by the 
94

Mars atmosphere program used in 

the simulation. Also, other 

atmospheric parameters have been 

set, based on MSL experience, that 

would further stress the system.

Dusttau 0.7 0.1:0.9 Uniform This determines the dust loading 

and thus the density and wind 

profiles that the vehicle will 

experience. This range provides 

large variability, but would not 

include dust storms.

Notes

* DV perturbation selected to produce a +/- 0.25 deg variation in entry Flight Path Angle

** Sign of lift changed so that 0 deg bank angle is lift up – MSL convention as compared to AFE convention

Lift and Drag Coefficient multipliers act on MIAS aerodynamic tables (MIAS Final Presentation, 

March 2002). 

MarsGRAM density and wind perturbations are applied to all cases except the nominal 

Monte Carlo analysis is not planned for the all-propulsive configuration at this time. 

A.2.6 Robotic Mission Monte Carlo Parameters 

As with the Exploration mission, the simulations underlying the system analyses performed in the 

robotic study utilize Monte Carlo procedures to account for various uncertainties in the system and its 

environment. Table A-7 lists the parameters that are treated probabilistically in the robotic simulations. 

For each uncertain parameter, the rationale for the choice of distribution is included. The model for the 

uncertainty in the Martian atmospheric density is the dust tau model contained in MARSGRAM. 

Atmospheric density perturbations are applied to all cases except the nominal, and winds are applied to all 

cases except nominal. 

Table A-7. Robotic Monte Carlo Parameters 

Parameter Nominal Perturbation Distribution Rationale

Entry State Delivery 

and Knowledge 

Error

Position, 

Velocity as 

necessary to 

impact 

desired 

landing site.

Per Entry States 

File

Normal Entry states file generated assuming 

a MSL dsn only attitude control 

system. Nominal entry state 

determined to be consistent with the 

Earth/Mars transfer trajectory and 

desired landing location for the date 

and time assumed.

Attitude Knowledge 

Error

Angle of 

Attack = 

30deg.

Bank = 70 

deg.

0.25 deg about a 

varying rotation 

axis.

Direction of 

rotation axis 

computed 

uniformly in all

directions (i.e. 

spherical 

distribution).

Again, a dsn only attitude control 

system is assumed. 

Lift/Drag 0.3 +/- 0.04 Normal Accounts for uncertainty in trim 

angle due to c.g. location 

uncertainty directly from detailed 

MSL analysis.

Angle of Attack

(deg)

-19.5 +/- 2 Normal The nominal angle of attack will 

vary to achieve the desired nominal 

lift/drag ratio. Perturbation value is 
95

typical variation from the ideal due 

to stability oscillations about the 

nominal trim angle.

Winds (m/s) 0 

(horizontal); 

0 (vertical)

Horizontal: 25 

m/s

Vertical: 20 m/s

Horizontal: 

constant 

magnitude with 

direction 

uniform from 0 

to 360 deg.

Vertical: normal 

distribution in 

direction and 

magnitude. 

Represents constructed engineering 

winds that encompass most wind 

conditions on Mars as compared to 

global circulation models contained 

in MarsGRAM.

Dusttau 0.33 0.1:0.75 Uniform This determines the dust loading 

and thus the density and wind 

profiles that the vehicle will 

experience. This range provides 

large variability, but would not 

include dust storms.

A.3 EDL-SA Study Figures of Merit 

The technology recommendations from the EDL-SA Study will be based on Figures of Merit (FOMs) 

that assess the impact of the technologies upon key performance, cost and risk metrics for the missions. 

A.3.1 Description of Figures of Merit 

In general, desired attributes of FOMs are that they be 

* as mutually independent as possible (or allow modeling of dependencies, 

* as non-redundant as possible (no double bookkeeping), 

* as complete as possible (include all potentially discriminating attributes between alternatives), 

* as few as possible (not include non-discriminating attributes), 

* as quantifiable as possible (or have proxy variables that are quantifiable, 

* as well-defined as possible (allow common understanding for evaluation, 

* grouped in a hierarchy (enable ease of evaluation/scoring roll-up). Need to define categories. 

*

The following two subsections define the categories and definitions of the figures of merit. In many 

cases, direct assessment of the particular FOMs is not possible due to the lack of a detailed system 

definition in early-phase design or because there is no definitive measure for the FOM. In these cases 

proxy variables are evaluated instead. The proxy variables are ones that are believed to have a strong 

correlation with the desired FOM. 

A.3.2 Categories for Figures of Merit 

The categories chosen for the FOMS are the following: 

Performance and Effectiveness - Provide Figures of Merit associated with performance and 

effectiveness, determining the degree to which a technology option effectively meets mission needs and 

improves mission performance 

Affordability and Life-Cycle Cost – Provide Figures of Merit associated with system life-cycle 

costs, determining the degree to which a mission concept or technology option reduces technology and 
96

system development and recurring costs 

Programmatic Risk – Provide Figures of Merit associated with technology and programmatic risk, 

determining the degree to which a technology option affects these risks. 

Applicability to Other Missions– Provide Figures of Merit associated with extensibility, 

determining the degree to which a mission concept or technology option could be used for other missions 

or customers. 

Safety and Mission Success - Provide Figures of Merit associated with safety and mission success, 

determining the degree to which a technology option ensures safety and reliability for all mission phases 

A.3.3 Definitions of Figures of Merit 

Table A-8 provides the definitions for the FOMs Exploration-class missions. The proxy variables for 

the FOMs used in the EDL-SA Study for Exploration-class missions are provided in Appendix B. 

FOMs for large-robotic-class missions are expected to differ from these to a minor extent. 

Table A-8. Definitions of Figures of Merit 

Category Figure of Merit Definition

Likelihood of Loss of 

Mission

Likelihood of a critical failure occurring during the MOI and 

Safety and Mission EDL phases resulting in loss of one or more mission objectives

Success Likelihood of Loss of 

Crew

Likelihood of a critical failure occurring during the MOI and 

EDL phases resulting in loss of one or more crew members

Sensitivity to Usable 

Payload Mass to Surface

Ability of the technology approach to successfully land usable 

hardware, cargo, fluids, and crew beyond that required for the 

mission.

Sensitivity to Surface 

Elevation

Ability of the technology approach to provide the capability to 

land at sites above the required surface elevation.

Sensitivity to Landing 

Precision

Ability of the technology to land closer to the desired landing 

site than required for the mission

Performance and 

Effectiveness

Sensitivity to 

Environmental 

Variability

Range of atmospheric density, dust and wind profiles that can be 

tolerated.

Technology 

Development Risk

Likelihood of Mars EDL architecture technology development 

Programmatic Risk activities to exceed schedule and budgetary constraints. 

Programmatic Cost and 

Schedule Risk

Likelihood of DDT&E activities for Mars EDL system to 

exceed planned development schedule or planned development 

budget for a given technology approach. 

Technology 

Development Cost

Cost to develop required technologies to TRL 6.

Advanced Development 

Cost

Cost to develop required technologies from TRL 6 to TRL 8.

Affordability and 

Life-Cycle Cost

System Life-Cycle Cost EDL-related system life-cycle cost (e.g., DDT&E, production, 

processing, launch, mission operations).

Applicability to Mars 

Robotic Missions

Applicability of technologies, systems, and operations to Mars 

robotic missions. 

Applicability to 

Other Missions

Applicability to Other 

Planetary Missions

Applicability of technologies, systems, and operations to other 

planetary entry missions.

A.3.4 Scoring for Figures of Merit 

There are many alternatives for scoring and rolling-up (weighting) the figures of merit. Particular 

choices for these have not yet been made. Study results will include sensitivities to weighting choices. 
97

A.3.5 FOM Proxy Parameters 

Table A-9. Safety and Mission Success Proxy Parameters 

Category Figure of Merit Definition Proxy Parameters

System/subsystem functional redundancy (e.g. 

engine-out capability).

Number and complexity of major EDL 

architecture elements/systems.

Number and complexity of interfaces between 

major EDL architecture elements/systems.

Number and complexity of in-space assembly 

operations.

Number of Mars configuration transitions 

required.

Likelihood of Loss 

of Mission

Likelihood of a 

critical failure 

occurring during 

the MOI and EDL 

phases resulting in 

loss of one or more 

mission objectives

Number and type of system risks and mission 

hazards.

Number and type of abort/safe haven options 

per mission phase.

System/subsystem functional redundancy (e.g. 

engine-out capability).

Number and complexity of major EDL 

architecture elements/systems.

Number of interfaces between major EDL 

architecture elements/systems.

Number and complexity of in-space assembly 

operations.

Number of Mars configuration transitions 

required.

Safety and 

Mission Success

Likelihood of Loss 

of Crew

Likelihood of a 

critical failure 

occurring during 

the MOI and EDL 

phases resulting in 

loss of one or more 

crew members

Number and type of system risks and mission 

hazards.
98

Table A-10. Performance and Effectiveness Proxy Parameters 

Category Figure of Merit Definition Proxy Parameters

Sensitivity to 

Usable Payload 

Mass to Surface

Ability of the 

technology 

approach to 

successfully land 

usable hardware, 

cargo, fluids, and 

crew beyond that 

required for the 

mission.

Sensitivity of Mars arrival system mass to 

increases in usable surface payload mass.

Sensitivity to 

Surface Elevation

Ability of the 

technology 

approach to 

provide the 

capability to land 

at sites above the 

required surface 

elevation.

Sensitivity of Mars arrival system mass to 

increases in the landing site surface elevation.

Sensitivity to 

Landing Precision

Ability of the 

technology to land 

closer to the 

desired landing site 

than required for 

the mission. 

Sensitivity of Mars arrival system mass to 

decreases in the distance between the final 

landing site and the specified landing site.

Performance and 

Effectiveness

Sensitivity to 

Environmental 

Variability

Range of 

atmospheric 

density, dust and 

wind profiles that 

can be tolerated.

Based on Monte-Carlo analyses, the 3-sigma 

ranges of density, dust and wind profiles that 

can be tolerated and meet the specified landing 

requirements.

Table A-11. Programmatic Risk Proxy Parameters 

Category Figure of Merit Definition Proxy Parameters

Number of technologies required.

TRLs of technologies.

Technology 

Development Risk

Likelihood of Mars 

EDL architecture 

technology 

development 

activities to exceed 

schedule and 

budgetary 

constraints. 

RD3 scores of technologies (e.g., number and 

type of large-scale integrated ground 

demonstrations required, number and type of 

required flight tests, including scale of Earth and 

Mars atmosphere tests).

Level of EDL system design 

maturity/uncertainty for given technology 

approach 

Number and level of complexity of Mars EDL 

architecture systems 

Number of interfaces between major Mars EDL 

architecture elements/systems 

Programmatic 

Risk

Programmatic Cost 

and Schedule Risk

Likelihood of 

DDT&E activities 

for Mars EDL 

system to exceed 

planned 

development 

schedule or 

planned 

development 

budget for a given 

technology 

approach. 

Number of configuration transitions required 
99

Table A-12. Affordability and Life Cycle Cost Proxy Parameters 

Category Figure of Merit Definition Proxy Parameters

Number of technologies required.

TRLs of technologies.

Technology 

Development Cost

Cost to develop 

required 

technologies to 

TRL 6.

RD3 scores of technologies (e.g., number and 

type of large-scale integrated ground 

demonstrations required, number and type of 

required flight tests, including scale of Earth and 

Mars atmosphere tests).

Total gross mass of Mars arrival systems.

Number of technologies required.

TRLs of technologies.

Advanced 

Development Cost

Cost to mature 

required 

technologies from 

TRL 6 to TRL 8.

RD3 scores of technologies (e.g., number and 

type of large-scale integrated ground 

demonstrations required, number and type of 

required flight tests, including scale of Earth and 

Mars atmosphere tests).

Total gross mass of Mars arrival systems. 

Total (length + diameter) of Mars arrival 

systems (stowed).

Total dry mass of Mars arrival systems.

Number and level of complexity of Mars EDL 

architecture systems.

Affordability and 

Life Cycle Cost

System Life-Cycle 

Cost

EDL-related 

system life-cycle 

cost (e.g., 

DDT&E, 

production, 

processing, launch, 

mission 

operations). Number of interfaces between major Mars EDL 

architecture elements/systems.

Table A-13. Applicability to Other Mission Proxy Parameters 

Category Figure of Merit Definition Proxy Parameters

Applicability to 

Mars Robotic 

Missions

Applicability of 

technologies, 

systems, and 

operations to Mars 

robotic missions. 

Level of hardware commonality and traceability 

between Mars Exploration-class mission and 

Mars Large-Robotic-class mission EDL 

technologies, systems, operations, 

infrastructure, and approaches. Applicability to 

Other Missions

Applicability to 

Other Planetary 

Missions

Applicability of 

technologies, 

systems, and 

operations to other 

planetary entry 

missions. 

Level of hardware commonality and traceability 

between Mars EDL and other planetary mission 

technologies, systems, operations, 

infrastructure, and approaches.
100

B Appendix B: EDL Transition Events 

B.1 Background 

During the FOM assessment both integrated architectures and individual technologies were 

considered. For the architectures assessment, the events that involved transitioning from one decelerator 

technology to another were identified as drivers for technical and programmatic risk as well as 

performance. Because of the perceived importance of these events to the architecture FOM assessment 

and the relative immaturity of the design concepts, a series of working meetings were organized with 

NASA EDL experts to develop design concepts and an understanding of the risk/performance trade-offs 

for key transition scenarios in the reference exploration class EDL architectures. The following is a 

summary of the results of the working meetings that were integrated into the FOM assessment. 

B.2 Methodology 

To collect the information needed to assess the relative benefits and challenges associated with each 

transition event, a list of key experts in the fields relevant to EDL were asked to participate in two 

meetings. The first meeting was focused on the design of the EDL transitions and the technical mission 

risks, and the second meeting was focused on the human roles and human safety considerations for each 

transition concept developed during the first meeting. 

The first meeting was held at JPL and technical experts were invited that had flight mission 

experience from various past Mars missions as well as MSL. The objectives of the JPL meeting are listed 

below: 

The EDL system transitions are defined, including the approaches to, and the subsystems for, 

accomplishing those transitions 

The EDL sequences of (non-transition) events are defined, including the approaches to, and the 

subsystems for, accomplishing those events 

The major ground and flight tests required for EDL system technology development and advanced 

development are defined at the top level 

The second meeting was held at JSC and technical experts were invited that had worked on Shuttle, 

ISS, and Orion. The objectives of the JSC meeting are listed below: 

The EDL system transitions developed at JPL are critiqued and refined from the perspective of 

crewed missions 

Critical EDL sequences of (non-transition) events for crewed missions are identified, along with the 

key constraints from the crew perspective 

The major ground and flight tests required for EDL system technology development and advanced 

development, particularly for abort modes, are defined at the top level 

B.3 Summary of Results 

The transitions in the reference were categorized by major transition and a total of 4 transition types 

were discussed: 

Rigid Aeroshell to SRP 
101

Rigid Aeroshell to Drag SIAD 

Inflatable Aeroshell to SRP 

Inflatable Aeroshell to Lifting SIAD 

For each transition type, a family of transition options was developed. The following figures give an 

overview of the different strategies developed for each type of transition and the color codes indicate the 

preference based on risk and performance. Green indicates the baseline option and orange indicates the 

second choice. 

B.3.1 Rigid Aeroshell to SRP 

Figure B-1 summarizes the family of options for the rigid aeroshell to SRP transition. 

Figure B-1. Rigid Aeroshell to SRP Transition Options 

All of these have the nose “into the wind” at transition except for option 1C2b which has a variant in 

which the aeroshell first rotates to put the tail “into the wind”. The major transition strategies include: 

1A. Front Exit: The rigid aeroshell would pitch nose first into the velocity field and then the nose 

would be ejected allowing the descent stage to exit out of the front of the aeroshell. Additional drag 

augmentation such as retrorockets and small parachutes could be used to help insure a clean 

separation between the descent stage and the aeroshell. Another concept discussed included having 

the RCS maneuver the disposed aeroshell after separation to insure that the aeroshell did not recontact the descent vehicle or any sensitive ground assets. 

1B. Rear Exit: The rigid aeroshell would pitch nose first into the velocity field and then the descent 

stage would then slide out of the rear of the aeroshell with engines firing. The firing of the descent 

engines would help ensure the separation of the two bodies and in addition the RCS on the aeroshell 

may be use to help ensure proper disposal of the aeroshell after separation. 

1C. Side Exit: This family of options includes several configurations of the descent stage within the 

aeroshell and for each option the descent stage would exit through the side of split pieces of the 

aeroshell. Charges would separate split the aeroshell into at least two pieces and those pieces would 

either be hinged or completely separate. Drag devices or retrorockets could be used to aid in the 
102

separation. 

The preferred option defined for this transition event is described in more detail in Figure B-2. 

Figure B-2. Baseline Rigid Aeroshell to SRP Transition Option 

B.3.2 Inflatable Aeroshell to SRP 

Figure B-3 summarizes the options developed for the inflatable aeroshell to SRP transition. 

Figure B-3. Inflatable Aeroshell to SRP Transition Options 
103

The major transition strategies for this transition include (see Figure B-3):

2A. Thrust Thru: This option would include holding on to the flexible aeroshell until touch down or 

very near to touch down. The aeroshell may have to be dropped to avoid interactions between the 

descent engines and the ground. 

2B. Slide Thru: These options include the descent stage sliding through the center of the inflatable 

aeroshell and then dropping off the nose cap in front of the vehicle before the descent stage engines 

are started. 

2C. Deflation: This option would include deflating the aeroshell and repackaging it such that the 

ballistic coefficient is high enough so the released material would fall in front of the descent stage. 

2D. Parachute: This option would include using a larger deployable decelerator in the rear of the 

descent stage to give enough separation force to ensure a clean separation. However, the size of the 

parachutes would need to be very large in order to reduce the ballistic coefficient of the descent stage 

enough to ensure a clean separation. 

The preferred option is described in more detail in Figure B-4. 

Figure B-4. Baseline Inflatable Decelerator to SRP Option 

Hypersonic Inflatable Aeroshell to Supersonic Inflatable Decelerator - Figure B-5 summarizes 

the family of options for the hypersonic inflatable aeroshell to larger inflatable lifting supersonic 

decelerator transition. 
104

Figure B-5. Inflatable Hypersonic Aeroshell to Supersonic Inflatable Transition Options 

The transition strategies include the following: 

5A. Staged Rings: This includes a staged IAD deployment starting at Mach number of 6-10 with the 

final rings deployed at Mach 2-3.

5B. Single Ring /Flexible Cone: This includes a single IAD deployment event. The inflatable outer 

torus would be attached to the Hypersonic IAD by a flexible fabric that would provide the 

aerodynamic surface used for deceleration. 

The preferred option is described in more detail in Figure B-6. 
105

Figure B-6. Baseline Hypersonic Inflatable Decelerator to Supersonic Inflatable Decelerator Option 

Rigid Aeroshell to Supersonic Inflatable Decelerator - Figure B-7 summarizes the options 

developed for the rigid aeroshell to inflatable decelerator transition. 

Figure B-7. Rigid Aeroshell to Inflatable Aerodynamic Decelerator Options 

The transition strategies include the following: 

7A. Rear Exit: This would include inflation of the aerodynamic decelerator out the back of the 

aeroshell and using the drag force to pull the descent stage away from the aeroshell. Rockets and the 

Aeroshell RCS may also insure a clean separation. 

7B. Top Exit: This would include ejecting the back of the aeroshell (or in the case of an entry 

without a backshell TPS no ejection would be required) and then inflating the aerodynamic 

decelerator. The drag force would then pull the descent stage up out of the top of the aeroshell as 

shown in the figure. 
106

The preferred option is described in more detail in Figure B-8. 

Figure B-8. Baseline Rigid Aeroshell to Inflatable Aerodynamic Decelerator 

Initial Failure Modes Assessment - During the Transitions meeting an initial set of generic EDL 

failure modes was developed. The following table describes the failure modes, description and notional 

mitigation strategy

Table B-1. EDL Failure Modes 

ID Phase Fault

Occurs

Phase 

Failure 

Occurs

Failure Mode Title Failure Mode Description

Notional 

Mitigation 

Strategy

FM.1 Aerocapture/

Entry

Aerocapture/

Entry

Local heating 

augmentation due to 

surface 

features/roughness

Bondline over-temperature condition 

exceeded due to heating augmentation 

caused by a surface protuberance/cavity

Design 

Margins

FM.2 Aerocapture/

Entry

Aerocapture/

Entry

Hypersonic flow 

impingement on descent 

stage

Uncertainty in predicting afterbody flow 

behind a "convertible" rigid entry 

vehicle or the HIAD configuration 

result in exceeding allowable 

temperatures of payload and/or descent 

stage systems

Design 

Margins

FM.3 Descent Descent Propulsion flow 

impingement on descent 

stage

Uncertainty in predicting propulsion 

exhaust flow impingement on descent 

stage results in exceeding allowable 

temperatures of payload and/or descent 

stage systems

Design 

margins

FM.4 Cruise Aerocapture/

Entry

Loss of aeroshell/TPS 

structural integrity due to 

combined loads

Thermal-Structural loading and/or load 

cycling during mission results in 

cracking/ripping of structural hardware

Design 

margins

FM.5 Cruise Aerocapture/

Entry

Natural (in-space) 

environment degrades 

In-space ionizing radiation, solar flare 

events, long duration vacuum exposure, 

Abort, flight 

Test
107

inflatable structure 

strength properties

etc. degrades the strength properties of 

deployable structures

FM.6 Cruise Aerocapture/

Entry 

In-space ionizing radiation, solar flare 

events, long duration vacuum exposure, 

etc. degrades the strength properties of 

deployable structures

Abort, flight 

Test

FM.7 Aerocapture/

Entry

Aerocapture/

Entry

Local heating 

augmentation due to 

control surfaces/thrusters

Bondline over-temperature condition 

exceeded due to heating augmentation 

caused by control surface or thruster

Design 

Margins

FM.8 Aerocapture/

Entry

Aerocapture/

Entry

Unexpected aerothermal 

environment

Uncertainty in predicting the 

aerothermal environment during entry 

results in bondline over-temperature 

condition

Design 

Margins

FM.9 Aerocapture/

Entry

Aerocapture/

Entry

Unexpected thermal 

response to aerothermal 

environment

Uncertainty in predicting the TPS 

material response during entry results in 

bondline over-temperature condition

Design 

Margins

FM.10 Unknown Aerocapture/

Entry

Deployable TPS/carrier 

structure damaged during 

assembly/storage/deploy

ment

Deployable TPS/support structure 

physically damaged due to natural 

environment, packing loads, storage 

loads, launch loads or deployment event

Inspection/R

epair

FM.11 Cruise Aerocapture/

Entry

In-space hypervelocity 

impact damage to 

payload, and/or support 

systems

MMOD damage punctures storage 

tanks, deflates inflatable structures, 

damages TPS, etc.

Inspection/R

epair/Abort

FM.12 Aerocapture/

Entry

Landing Unexpected aerodynamic 

coefficients of entry 

vehicle

Uncertainty in prediction of 

aerodynamic performance of entry 

vehicle results in unexpected trajectory

Design 

Margins

FM.13 Reconfigurati

on

Descent Re-contact with 

jettisoned elements

Hardware jettisoned during reconfiguration events impacts and 

damages payload or support systems

Flight Test

FM.14 Descent Landing Descent engine does not 

start

Main descent engines fail to start up as 

expected due to mechanical system 

failure

Redundancy

FM.15 Descent Landing Payload unstable after 

touchdown due to engine 

transient shut down

Engine does not shut down as expected 

after descent stage contact with surface 

resulting in payload instability

Crew 

Controls

FM.16 Descent Landing Descent stage instability 

during descent

Forward jettisoned hardware disrupts 

free stream flow seen by the descent 

stage resulting in unexpected 

aerodynamic moments acting on the 

descent stage

Design 

Margins

FM.17 Descent Landing Unexpected propulsion 

performance

Uncertainty in main engine performance 

results in unexpected trajectory (e.g. 

caused by forward jettisoned hardware 

or unexpected supersonic shock 

interaction)

Design 

Margins

FM.18 Descent Landing Unexpected aerodynamic 

coefficients of drag 

device

Uncertainty in aerodynamic 

performance of deployed drag devices 

results in unexpected trajectory

Design 

Margins

FM.19 Descent Landing Aerodynamic drag 

devices not deployed 

Aerodynamic drag devices not properly 

inflated or deployed

Redundancy

FM.20 Descent Landing Descent stage instability 

close to surface features

Main propulsion plume interaction with 

distributed surface features creates nonsymmetric pressure distribution on 

descent stage which results in loss of 

vehicle stability

Crew 

Controls

FM.21 Reconfigurati

on

Descent Loss of vehicle structural 

integrity due to pyro 

shock

Vibro-acoustic shock from pyro devices 

used during reconfiguration events 

result in structural damage to the 

payload or support systems

Design 

Margins
108

FM.22 Cruise/Aeroca

pture/Entry

Aerocapture/

Entry

TPS gap/seam thermal or 

mechanical failure

Thermal or mechanical failure in or 

around a gap/seam in TPS (e.g. crack in 

seam due to combined loads, bondline 

under seam exceeds temperature limit)

Design 

Margins

FM.23 Landing Landing Unstable landed 

configuration due to 

unexpected surface 

features

Descent stage touch down on unstable 

or non-flat surface feature resulting in 

loss of vehicle stability on the surface

Crew 

Controls

FM.24 Landing Landing Unstable landed 

configuration due to 

trenching caused by main 

engines

Descent stage touch down on unstable 

or non-flat surface features caused by 

main engine trenching resulting in loss 

of vehicle stability on the surface

Surface Prep

FM.25 Descent Landing Debris damage during 

terminal descent

Debris kicked up by main engines 

damages the payload, support systems 

or other landed assets

Design 

Margins

FM.26 Aerocapture/

Entry/Descent

Aerocapture/

Entry/Landi

ng

Gas bladder overpressure Entry heating or inflation management 

system failure result in ripping/tearing 

of inflatable structures

Design 

Margins

FM.27 Descent Descent Loss of descent stage 

structural integrity due to 

propulsion induced 

vibro-acoustic loads

Uncertainty in the vibration and 

acoustic environment during powered 

descent result in structural damage to 

the payload or support systems

Design 

Margins

FM.28 Descent Descent Snap loads from 

Aerodynamic decelerator 

deployment damage 

descent stage and/or 

decelerator itself

Unexpected impulse forces during 

aerodynamic decelerator deployment 

results in damage to payload or support 

systems

Design 

Margins

FM.29 Descent Descent Loss of control 

surface/thruster 

functionality

Loss of commanded control capability 

during aerocapture, entry or descent 

results in unexpected trajectory

Redundancy

FM.30 Descent Landing Dynamic instability of 

descent stage caused by 

aerodynamic decelerator 

deployment

Non-symmetric deployment of 

aerodynamic decelerator results in 

dynamic instability of the descent stage

Design 

Margins

FM.31 Descent Landing Loss of flight computer 

functionality during 

AEDL

Flight computer fails to function during 

descent

Redundancy
REPORT DOCUMENTATION PAGE Form Approved

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Entry, Descent and Landing Systems Analysis Study: Phase 1 Report 

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 6. AUTHOR(S)

Dwyer Cianciolo, Alicia M.; Davis, Jody L.; Komar, David R.; Munk, Michelle M.; 

Samareh, Jamshid A.; Williams-Byrd, Julie A.; Zang, Thomas A.; Powell, Richard 

W.; Shidler, Jeremy D.; Stanley, Douglas O.; Wilhite, Alan W.; Kinney, David J.; 

McGuire, M. Kathleen; Arnold, James O.; Howard, Austin R.; Sostaric, Ronald R.; 

Studak, Joseph W.; Zumwalt, Carlie H.; Llama, Eduardo G.; Casoliva, Jordi; 

Ivanov, Mark C.; Clark, Ian; Sengupta, Anita 

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NASA senior management commissioned the Entry, Descent and Landing Systems Analysis (EDL-SA) Study in 2008 to 

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Systems analysis; Mars landing; Payloads; Parachutes; Decelerators; Entry; Descent; Landing; Aerocapture; Technology 

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