December 2009 

NASA/TM-2009-216165 

NESC-RP-06-060 

 

Constellation Program (CxP) Crew 

Exploration Vehicle (CEV) 

Project Integrated Landing System 

John D. Baker 

Jet Propulsion Laboratory, Pasadena, California 

Daniel E. Yuchnovicz/NESC 

Langley Research Center, Hampton, Virginia 

David J. Eisenman and Scott G. Peer 

Jet Propulsion Laboratory, Pasadena, California 

Edward L. Fasanella 

Langley Research Center, Hampton, Virginia 

Charles Lawrence 

Glenn Research Center, Cleveland, Ohio 
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December 2009 

NASA/TM-2009-216165 

NESC-RP-06-060 

 

Constellation Program (CxP) Crew 

Exploration Vehicle (CEV) 

Project Integrated Landing System 

John D. Baker 

Jet Propulsion Laboratory, Pasadena, California 

Daniel E. Yuchnovicz/NESC 

Langley Research Center, Hampton, Virginia 

David J. Eisenman and Scott G. Peer 

Jet Propulsion Laboratory, Pasadena, California 

Edward L. Fasanella 

Langley Research Center, Hampton, Virginia 

Charles Lawrence 

Glenn Research Center, Cleveland, Ohio 
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443-757-5802 

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NASA Engineering and Safety Center 

Technical Assessment Report 

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NESC-RP06-060 

Version: 

1.0 

Title:

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1 of 314 

NESC Request No.: 06-060-E 

 

Constellation Program (CxP) Crew Exploration Vehicle (CEV) 

Project Integrated Landing System 

October 15, 2009 
NASA Engineering and Safety Center 

Technical Assessment Report 

Document #: 

NESC-RP06-060 

Version: 

1.0 

Title:

Crew Exploration Vehicle Integrated Landing System 

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2 of 314 

NESC Request No.: 06-060-E 

Report Approval and Revision History 

Approval and Document Revision History 

NOTE: This document was approved at the October 15, 2009, NRB. This document was 

submitted to the NESC Director on November 2, 2009, for configuration control. 

Approved 

Version: 

Original Signature on File 11/3/09 

1.0 NESC Director Date 

Version Description of Revision Office of Primary 

Responsibility

Effective Date 

1.0 Initial Release John Baker, NESC 

Assessment Lead and 

Daniel Yuchnovicz, 

NESC SEO 

October 15, 2009
NASA Engineering and Safety Center 

Technical Assessment Report 

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NESC Request No.: 06-060-E 

Table of Contents 

Volume I: Technical Assessment Report 

1.0 Notification and Authorization ........................................................................................ 8

2.0 Signature Page ................................................................................................................... 9

3.0 Team List ......................................................................................................................... 10

4.0 Executive Summary ........................................................................................................ 12

5.0 Assessment Plan .............................................................................................................. 19

5.1     Charter and Background ................................................................................................... 19

5.2     Scope ............................................................................................................................... 19

6.0 Description of the Problem and Approach ................................................................... 19

6.1    Phase I - Comparison of LM 604, TS-LRS001, and Apollo LRS – Problem and 

Approach ........................................................................................................................... 19

6.2     Phase II -Integrated LRS Risk Assessment – Problem and Approach ............................. 23

7.0 Data Analysis ................................................................................................................... 37

7.1    Phase I - Comparison of 604, TS-LRS001, and Apollo LRS Data Analysis .................... 37

7.2    Phase II - Integrated LRS Risk Assessment Data Analysis .............................................. 40

8.0 Findings and Recommendations .................................................................................. 121

8.1    Phase I - Comparison of 604, TS-LRS001, and Apollo LRS ......................................... 121

8.2     Phase II - Integrated LRS Risk Assessment ................................................................... 125

9.0 Alternate Viewpoints .................................................................................................... 130

10.0 Other Deliverables ........................................................................................................ 130

11.0 Lessons Learned ............................................................................................................ 130

12.0 Definition of Terms ....................................................................................................... 131

13.0 Acronyms List ............................................................................................................... 132

14.0 References ...................................................................................................................... 133

Volume II: Appendices 

Appendix A. Crew Exploration Vehicle Integrated Landing and Recovery Evaluation .......... 135

Appendix B. Task 2 T&V Assessment .................................................................................... 162

Appendix C. Task 1 Integrated Entry, Descent and Landing Risk Matrix .............................. 299

 
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List of Figures

Figure 3.0-1. CEV Integrated Landing Assessment Team Structure ......................................... 11

Figure 4.0-1. Risk Comparison of the NASA TS-LRS001 and LM 604 Landing and Recovery 

System Baselines .................................................................................................. 13

Figure 4.0-2. Systems Approach to the Integrated Landing and Recovery System Risk 

Assessment ............................................................................................................ 14

Figure 6.1-1. NASA and LM Landing and Recovery Baselines used for Comparison in 

Phase I ................................................................................................................... 20

Figure 6.1-2. Post-Entry Descent and Landing Event Sequence (Functional Decomposition of 

Main Events) for a Successful Landing ................................................................ 22

Figure 6.1-3. Vulnerability Concern Color Code: Concern Colors Are Assigned per Sub-Event 

with the Assumption that All Previous Sub-Events and Sequences Are 

Successful ............................................................................................................. 22

Figure 6.2-1. Systems Approach to the Integrated LRS Risk Assessment ................................. 23

Figure 6.2-2. Phases of the Entry, Descent, and Landing Included in the Integrated Risk 

Assessment ............................................................................................................ 24

Figure 6.2-3. Approach to determining Landing Site Accessibility and Availability ................ 25

Figure 6.2-4. Concept of Operations for an ISS Return to Land-Landing Sites ........................ 26

Figure 6.2-5. SLAAM User Interface ......................................................................................... 28

Figure 6.2-6. Conceptual Automatic Parachute Separation on Landing to Prevent CM Tumbling 

(left) and Simplified LS-DYNA® Model (right) .................................................. 30

Figure 6.2-7. Comparison of Results of Parachute Forces from Two Different Computational 

Methods................................................................................................................. 32

Figure 6.2-8. CM Structural Parts used in the Model ................................................................. 33

Figure 6.2-9. Overview of the Complete Finite Element Model ................................................ 34

Figure 6.2-10. Effect of Rigging Stiffness on CM Rollover (0 degree hang, Parachute Remains 

Attached) ............................................................................................................... 36

Figure 6.2-11. Effect of Rigging Stiffness on Pallet Acceleration (0 degree hang, Parachute 

Remains Attached) ................................................................................................ 36

Figure 7.1-1. Vulnerability Concern Map .................................................................................. 39

Figure 7.2-1. Summary of Risk Drivers for the Nominal ISS Return Mission – Updated from 

the NESC CEV Water Versus Land-Landing Study (Before and After) .............. 43

Figure 7.2-2. Results of the Risk Update .................................................................................... 44

Figure 7.2-3. Comparison of All Options (Uses Generation 2 Airbags) .................................... 49

Figure 7.2-4. Comparisons of Lower Risk Options with Generation 2 Airbags (bottom) ......... 50

Figure 7.2-5. CM Event Sequence Diagram Used to Investigate the Completeness of the CEV 

Project’s T&V Approach ...................................................................................... 51

Figure 7.2-6. Problem geometry: illustration of cross range options, ascending and descending 

passes and 3 orbit ground trace ............................................................................. 53

Figure 7.2-7. Concept of Operations: CM Disposal and Raise Maneuver ................................. 54

Figure 7.2-8. Diurnal wind study results, Edwards 2006 ........................................................... 56

Figure 7.2-9. Seasonal Study: Summer 2006, EAFB ................................................................. 57
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Figure 7.2-10. Percent Capture versus Wind Speed (knots) ......................................................... 59

Figure 7.2-11. Monte Carlo Simulation Flow Chart for Statistical Weather Analysis ................. 64

Figure 7.2-12. Landing delay risk - Chance of abort is 47 percent for the 18 fps baseline case .. 67

Figure 7.2-13. Logarithmic plot of risk of LOC over criteria ...................................................... 68

Figure 7.2-14. Logarithmic plot of risk of LOC over capability ................................................. 68

Figure 7.2-15. Landing delay cumulative probability, 1 site, 18 fps ........................................... 69

Figure 7.2-16. Landing delay cumulative probability, 3 sites, 18 fps ......................................... 70

Figure 7.2-17. Landing delay cumulative probability, 1 site, 18 fps, daytime ............................. 70

Figure 7.2-18. Landing delay cumulative probability, 3 sites, 18 fps, daytime ........................... 71

Figure 7.2-19. Landing delay cumulative probability, 1 site, 13 fps, daytime ............................. 71

Figure 7.2-20. Updated Landing Opportunities versus Risk ........................................................ 73

Figure 7.2-21. Water-landing illustration ..................................................................................... 74

Figure 7.2-22. Water as back up to Edwards on 1/5 ..................................................................... 75

Figure 7.2-23. Landing accessibility for the landing site network, all landing options available, 

180 days (14 days shown for clarity), in hours and days ...................................... 75

Figure 7.2-24. Water as back up to EAFB.................................................................................... 76

Figure 7.2-25. Lunar inclination CONUS landing evaluation ...................................................... 77

Figure 7.2-26. Crew Acceleration Limits for the Apollo Command Module .............................. 78

Figure 7.2-27. CM Roll Torque Capability Under Main Parachutes ............................................ 79

Figure 7.2-28. Rigging Force without Chute Release at a 15 degree Hang Angle ....................... 80

Figure 7.2-29. Effect of Parachute Release Time on CM Roll - 15 degree Hang Angle, CM does 

not rollover when chute released before 1.0 seconds. .......................................... 81

Figure 7.2-30. Effect of Parachute Release Time on Pallet Acceleration 15 degree Hang Angle 82

Figure 7.2-31. 0 Degrees Hang Angle (Rigging Attachment Located to Maintain Hang Angle) 83

Figure 7.2-32. Rigging Force without Chute Release 0 degree Hang Angle ............................... 84

Figure 7.2-33. Effect of Parachute Release Time on CM Roll 0 degree Hang angle, Vehicle rolls 

over for all cases ................................................................................................... 85

Figure 7.2-34. Effect of Parachute Release Time on Pallet Acceleration 0 degree Hang Angle . 85

Figure 7.2-35. Effect of Horizontal Velocity on Rollover 0 degree Hang Angle, Vertical velocity 

= 26 fps, No Roll ................................................................................................... 86

Figure 7.2-36. Effect of Horizontal Velocity on Rollover 15 degree Hang Angle, Vertical 

Velocity = 26 fps, No Roll .................................................................................... 87

Figure 7.2-37. Six Crew Configuration with local X, Y, and Z-Axis .......................................... 88

Figure 7.2-38. +Z Acceleration Tolerance with Ejection Seat Data Omitted .............................. 90

Figure 7.2-39. +Z-Human Tolerance Limits showing Ejection Seat Design Limits .................... 90

Figure 7.2-40. DRI+Z Related to Probability of Spinal Injury for Ejection Seats ....................... 91

Figure 7.2-41. Central Mass Located in the Mid-Thorax (Brinkley Dynamic Response Method 

Model) ................................................................................................................... 96

Figure 7.2-42. CM 604 Seat Attenuation Model .......................................................................... 98

Figure 7.2-43. Configuration for X Acceleration (eyes in/out DRI) ............................................ 99

Figure 7.2-44. Acceleration Profile for Low Eyes In/Out DRI (Vn=33.3 fps) ............................ 99

Figure 7.2-45. Acceleration Profile for Moderate Eyes In/Out DRI (Vn=41.67 fps) ................ 100
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Figure 7.2-46. Acceleration Profile for High Eyes In/Out DRI (Vn=58.33 fps) ........................ 100

Figure 7.2-47. Configuration for Y Acceleration (sideways DRI) ............................................. 101

Figure 7.2-48. Acceleration Profile for Low Sideways DRI (Vn=16.66 fps) ............................ 101

Figure 7.2-49. Acceleration Profile for Moderate Sideways DRI (Vn= 25 fps) ........................ 102

Figure 7.2-50. Acceleration Profile for High Sideways DRI (Vn= 33.3 fps) ............................. 102

Figure 7.2-51. Configuration for Z Acceleration (spine DRI) .................................................... 103

Figure 7.2-52. Acceleration Profile for Low Spine DRI (Vn= 25 fps) ....................................... 103

Figure 7.2-53. Acceleration Profile for Moderate Spine DRI (Vn= 29.2 fps) ........................... 104

Figure 7.2-54. Acceleration Profile for High Spine DRI (Vn= 33.3 fps) ................................... 104

Figure 7.2-55. Location of Head Acceleration ........................................................................... 105

Figure 7.2-56. Location of Pelvis Acceleration .......................................................................... 105

Figure 7.2-57. Location of Thorax Acceleration ........................................................................ 106

Figure 7.2-58. Location of Lower Lumbar Spine Force ............................................................. 106

Figure 7.2-59. Location of Lower Lumbar Spine Moment ........................................................ 107

Figure 7.2-60. Location of Neck Joints ...................................................................................... 107

Figure 7.2-61. Location of Upper Neck Moment ....................................................................... 108

Figure 7.2-62. Location of Lower Neck Moment ....................................................................... 108

Figure 7.2-63. Hybrid III ATD Constrained with Five Point Harness ....................................... 109

Figure 7.2-64. Hybrid III ATD Response at Intermediate Time Steps ....................................... 110

Figure 7.2-65. Modern Race Car Seat with Lateral Supports (left) and Apollo CM seats 

(right) .................................................................................................................. 112

Figure 7.2-66. Airbag Concept shown in a CEV Six-Seat Configuration .................................. 113

Figure 7.2-67. Airbag Sketch of Adjustable Length Linkage .................................................... 113

Figure 7.2-68. Comparison Among Padded and Constrained Crew Member for Sideways 

Loading ............................................................................................................... 114

Figure 7.2-69. Hybrid III ATD Response at Intermediate Time Steps for Side Impact (No Side 

Padding) .............................................................................................................. 115

Figure 7.2-70. Hybrid III ATD Response at Intermediate Time Steps for Side Impact (Thin Side 

Padding) .............................................................................................................. 116

Figure 7.2-71. Acceleration Profile for Maximum Drag Abort .................................................. 118

Figure 7.2-72. Hybrid III ATD Response at Intermediate Time Steps for Maximum Drag 

Abort ................................................................................................................... 119

Figure 7.2-73. Acceleration Profile for Maximum Drag Abort, Brinkley Dynamic Response 

Method and FE ATD Chest Acceleration ........................................................... 120

 
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List of Tables

Table 7.1-1. Areas of Comparison between the Apollo ELS and the LM and NASA CM 

LRSs ...................................................................................................................... 38

Table 7.2-1. PLOC for Risk Driver (Tabular) ........................................................................... 45

Table 7.2-2. Comparison of L/D 0.30 and 0.35 for EAFB (Edwards) landing site, all options 

available (day and night), 180 days (first 14 days shown) ................................... 52

Table 7.2-3. Landing accessibility for NASA’s landing site network, all landing options 

available, 180 days. ............................................................................................... 55

Table 7.2-4. Landing site availability for Edwards with day landing only constraint .............. 55

Table 7.2-5. 5 knot wind threshold (8.4 fps) ............................................................................. 58

Table 7.2-6. 10 knot wind threshold (16.8 fps) ......................................................................... 58

Table 7.2-7. 15 knot wind threshold (25.3 fps) ......................................................................... 58

Table 7.2-8. 20 knot wind threshold (33.7 fps) ......................................................................... 58

Table 7.2-9. Gust elimination, No wind threshold, EAFB 2006 ............................................... 59

Table 7.2-10. Offset Example, 20 knot threshold, no pre- or post-offset, forecast data ............. 60

Table 7.2-11. Offset Example, 20 knot threshold, 6-hour-pre- and 3-hour-post-offset, forecast 

data ........................................................................................................................ 61

Table 7.2-12. 2006 Winds, Reported in Knots, at Representative Landing Sites ....................... 62

Table 7.2-13. Raise maneuver classification ............................................................................... 62

Table 7.2-14. Availability at EAFB without Raise Maneuvers .................................................. 63

Table 7.2-15. Availability at Carson Sink, NV. EAFB, and the Utah Test Range without Raise 

Maneuvers ............................................................................................................. 63

Table 7.2-16. Results from the Monte Carlo Weather Analysis ................................................. 66

Table 7.2-17. Monte Carlo Results Incorporating Increased Capability of Generation 2 

Airbags .................................................................................................................. 72

Table 7.2-18. Brinkley Dynamic Response Limits (NASA HSIR Specification for the CEV 

Project) .................................................................................................................. 93

Table 7.2-19. Dynamic Response Limits from Brinkley and Moser [Ref. 8] ............................. 93

Table 7.2-20. Vertical Velocity and Brinkley Dynamic Response Method Injury Landing 

Vertical Velocity (fps)/Brinkley DRI ................................................................... 95

Table 7.2-21. Hybrid III Response and Brinkley Dynamic Response Method Injury Levels ZAxis (Spinal) Excitation ...................................................................................... 110

Table 7.2-22. Comparison Among ATD Configurations Using Input Acceleration Profile that 

Produces High Brinkley Dynamic Response Method Injury in Lateral (Sideways) 

Direction ............................................................................................................. 116
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 Restriction Distribution until Report Approval by NRB and Export Control Review 

NESC Request No.: 06-060-E 

Volume I: Assessment Report 

1.0 Notifi cation and Authorization 

This assessment was initiated out-of-board by the authority of the NASA Engineering and Safety 

Center (NESC) Director on October 13, 2006. Julie Kramer White, Crew Exploration Vehicle 

(CEV) Chief Engineer requested a risk comparison of the Integrated Landing System design 

developed by NASA and the design developed by Contractor– referred to as the LM 604 

baseline. Based on the results of this risk comparison, the CEV Chief engineer requested that the 

NESC evaluate identified risks and develop strategies for their reduction or mitigation. The 

assessment progressed in two phases. 

A brief Phase I analysis was performed by the NESC team to compare the CEV Integrated 

Landing System proposed by the Contractor, as defined in the LM 604 baseline, against the 

NASA TS-LRS001 baseline with respect to risk (safety and reliability). 

During the Phase II effort, the NESC team further examined the areas of critical importance to 

minimizing the overall landing risk, including risk to the crew and the Crew Module (CM) 

during a nominal land-landing. The areas studied included: 

Task 1: Landing System Risk Assessment - Estimate the risk to the crew during landing 

using the LM 606 baseline and postulated landing attenuation system configurations. 

Task 2: Landing System Test and Verification (T&V) Approach - Assess the CEV 

Project’s T&V approach to determine the level of residual risk. 

Task 3: Landing Site Accessibility and Availability - Evaluate implications of relying on 

wind limits to ensure that horizontal velocity at touchdown remains within the CM design 

capability. 

Task 4: CM Roll Control in Preparation for Landing – Assess the requirements and 

capability of the CM roll control concept to orient the CM for an acceptable landing. 

Task 5: Investigate Parachute Effects during Landing - Investigate parachute induced 

effects on CM stability during landing near or beyond the CM’s horizontal velocity 

capability. 

Task 6: Crew Protection System Enhancements – Develop tools to evaluate enhanced 

crew restraints and develop restraint concepts that would enhance crew safety during 

high-g impacts or CM tumbling 

The key stakeholders for this assessment were Ms. Julie Kramer White at the Johnson Space 

Center (JSC), and Mr. Christopher Johnson, JSC, who served as the CEV Project Office liaison. 
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NESC Request No.: 06-060-E 

2.0 Signature Page 

Submitted by: 

Team Signature Page on File – 12/1/09 

_____________________________________ __________________________________ 

Mr. John D. Baker Date Mr. Daniel E. Yuchnovicz Date 

Significant Contributors: 

_____________________________________ __________________________________ 

Mr. David J. Eisenman Date Mr. Scott G. Peer Date 

_____________________________________ __________________________________ 

Dr. Edward L. Fasanella Date Dr. Charles Lawrence Date 

 

Signatories declare the findings and observations complied in the report are factually based from 

data extracted from Program/Project documents, contractor reports, and open literature, and/or 

generated from independently conducted tests, analysis, and inspections. 
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3.0 Team List 

Name Discipline Organization 

Core Team 

John Baker Team Lead JPL 

David Eisenman Deputy Lead JPL 

Daniel Yuchnovicz NESC SEO LaRC 

Scott Peer Systems Engineering JPL 

Charles Lawrence Landing Simulation GRC 

Ed Fasanella Landing & Recovery LaRC 

Consultants 

Brian Bairstow Systems Engineering JPL 

George Chen Systems Engineering JPL 

 Courtenay Clifford Safety and Mission Assurance JSC 

Chris Johnson NASA CEV Project System Engineering JSC 

Wayne Lee Systems Engineering JPL 

Blake Putney Risk Assessment Valador 

Malcolm Cohen Human Factors 

NASA, ARC 

(Retired) 

Phil Glynn Peer Review SAIC

T.K. Mattingly Peer Review 

Space Planning and 

Analysis, Inc. 

Jerry McCullough Peer Review SAIC 

Dave Shemwell Systems Engineering 

 Space Planning and 

Analysis, Inc. 

Milt Silvera Peer Review SAIC 

Al Tabiei SAIC 

University of 

Cincinnati 

Bob West Peer Review SAIC 

Support Personnel 

Lisa McAlhaney Program Analyst LaRC 

Linda Burgess Planning & Control Analyst LaRC, ATK 

 Christina Cooper Technical Writer LaRC, ATK 
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Figure 3.0-1. CEV Integrated Landing Assessment Team Structure 

Front row: David Eisenman, Brian Bairstow, Chris Johnson, Blake Putney, Christina Cooper, Mal Cohen, Scott Peer, John Baker 

Second Row: Dave Shemwell, Wayne Lee, Ed Fasanella, Bob West, Debora Briggs, Chuck Lawrence, Dan Yuchnovicz 

Third Row: Phil Glynn, Jim McMichael, Jeff Cyphert, T.K. Mattingly 
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4.0 Executive Summary 

This assessment was performed in two phases. A brief Phase I analysis was performed by the 

Water versus Land-Landing Team (NESC Assessment Number 06-020-E) to compare the Crew 

Exploration Vehicle (CEV) Integrated Landing System proposed by the Contractor, as defined in 

the LM 604 baseline, against the NASA TS-LRS001 baseline with respect to risk (safety and 

reliability). A phase II effort examined the areas of critical importance to the overall landing 

risk, evaluating risk to the crew and to the CEV Crew Module (CM) during a nominal landlanding.

Phase I Results Summary 

The Phase I analysis was performed to compare the CEV Integrated Landing System proposed 

by the contractor Lockheed Martin (LM), as defined in the LM 604 baseline, with the NASA TSLRS001 baseline with respect to risk factors potentially leading to severe injury or loss of crew 

(LOC). Due to the early stage of development for both systems, insufficient detail exists for 

either a numerical risk analysis or reliability calculation, thus the following approach was 

utilized: 

• Focus on potential design reliability - mass, volume, cost, and complexity were 

secondary considerations unless obvious concern existed. 

• Compare the two designs with the proven Apollo Earth Landing System (ELS). 

• Use existing design and performance data from NASA and LM CEV teams. 

• Do not incorporate preliminary requirements changes from the CEV Project’s 

Requirement Analysis Cycle-3 (RAC-3) because the NESC task was run in parallel 

with RAC-3 activities. 

• Develop a functional description for each baseline (functional block/event sequence 

diagram). 

• Quantify/evaluate the relative risks of each function. 

The end result of the comparison/evaluation is summarized in Figure 4.0-1. A list of ten 

concerns (C1–C10) and associated NESC recommendations were generated. The major 

concerns included: 

• C1 - LM 604 - Use of a single drogue to extract all the main parachutes was viewed 

as a critical single point failure. 

• C2, C4 LM 604 - 600 lbm confluence retro-rocket pack mounted at the confluence 

point on the main parachute harness is viable, but will require an extensive 

development and validation program. Primary concerns include deployment without 

contacting CM retrorocket plume impingement, parachute risers, and retrorocket pack 

(200 lbm) re-contact with CM after touchdown. 

• C3 - LM 604 - The use of horizontal wind velocity limits for LM 604 was unproven 

for effectiveness. Insufficient data had been produced to demonstrate that relying on 

operational wind limits without horizontal retrorockets was a viable design option. 
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NESC Request No.: 06-060-E 

C5 - LM 604, TS-LRS001 – The reliability and robustness of the apex/Forward Bay 

Cover (FBC) separation system needs to be improved to ensure high reliability and 

adequate clearance with the CM during the mission descent phase. Failure to deploy 

the FBC would result in LOC. 

Figure 4.0-1. Risk Comparison of the NASA TS-LRS001 and LM 604 Landing and Recovery 

System Baselines 

The findings were presented to Julie Kramer White and the CEV Project Configuration Control 

Board (CPCB) on December 2006. Key team members of the NESC Water versus LandLanding Assessment were used to perform this task and were critical in influencing the CEV 

Project and Contractor decision to adopting the NASA TS-LRS001 design as the baseline 

configuration. 

Phase II Results Summary 

A systems approach was required to assess the overall risks since the Landing System design can 

affect multiple CEV subsystems and will operate in a range of landing environments. Figure 

4.0-2 summarizes the six tasks and their relationships. 
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Figure 4.0-2. Systems Approach to the Integrated Landing and Recovery System Risk Assessment 

Phase II Study Task Areas 

Task 1: Integrated Risk Analysis Summary 

The purpose of the Task 1 risk assessment was to identify the risk drivers of different CM 

landing configurations and provide overall risk comparisons of the alternatives. Several 

spacecraft landing configuration baselines were examined including: the LM 606, Zero Based 

Vehicle (ZBV): and a water-lander based on the ZBV with minor modifications, with/without 

vertical / horizontal retrorockets, roll control, and air bags. 

The NESC recommendations included: 

Improve the FBC release (Apex Cover) to minimize interference with drogue parachute 

deployment. 

Increase the horizontal velocity attenuation to protect against under predicted horizontal 

winds at land-landing site, increase stability of the CM during landing, and increases 

landing availability. 

Improve crew occupant protection systems as functional redundancy to any landing 

attenuation system failures. 
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Task 2: Landing System Test and Verification (T&V) Summary 

This task was designed to investigate the completeness of the Landing System T&V plan. A 

landing event sequence diagram was developed and used as the basis of the investigation. The 

major NESC recommendations included: 

 Explicitly add a test program to exercise end-to-end performance of the Landing and 

Recovery System (LRS) with emphasis on interactions between subsystems. 

 Formulate a working group charged with the responsibility for verification of end-to-end 

LRS performance with emphasis on subsystem-to-subsystem and subsystem-to-CM 

interactions. 

 Adopt a strategy of developing specific LRS tests based on evaluation and verification 

needs as opposed to forcing the verification plan to conform to already existing “all 

encompassing” tests. 

 Develop T&V scenarios to test the FBC release, parachute extraction and deployment, 

main heat shield jettison (if required), and active controls used for impact attenuation or 

control during descent, landing, and airbag inflation (if used). 

Task 3: Landing Site Accessibility and Availability Summary 

The terrestrial landing site availability from Low Earth Orbit (LEO) task addressed the 

interaction of continental United States (CONUS) landing accessibility and availability. 

Accessibility is the number of times a given landing site can be physically reached from the 

initial orbit. Availability is defined as the number of times opportunities can be utilized in the 

presence of operational constraints. This study covered accessibility based on orbital mechanics, 

landing site location and CM lift to drag ratio (L/D). Availability topics studied included 

day/night, ascending/descending orbit pass constraints, surface wind conditions and thresholds, 

gust conditions, and Service Module (SM) disposal footprints. 

The task also determined that water-landings were a viable back-up to CONUS-based landlandings and could also be used as standalone landings. Suggestions included choosing a water 

based location that takes advantage of ascending and descending pass opportunities (e.g., off the 

coast of California). In addition, because of the flexibility of water-landings – through ship 

movement – additional opportunities for vehicle return become apparent. There were no obvious 

CONUS-based land-landing opportunities for the 28.5 degree inclination orbit. CONUS sites are 

accessible if an inclination change orbit maneuver is performed. 

The major NESC recommendations included: 

 Implement improved wind forecasting at the landing sites because current wind 

forecasting is not sufficiently accurate at the relatively low velocity constraints that are 

important to CM stability. 

 Improve crew protection to guard against misforecast winds or gusts, and to avoid over 

constraining wind placards. 

 Develop a flight test program that can support International Space Station (ISS) schedules 

and be prepared to capitalize on performance envelope expansion opportunities as they 
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emerge. The initial flight test program will be paced by limited opportunities to satisfy 

initial land-landing wind placards. 

 Preplanned water-landing sites can backup missed or unavailable land-landing site 

opportunities. 

Task 4: CM Roll Control in Preparation for Landing Summary 

The use of roll control can optimize the orientation of the crew seat attenuation system for landor water-landings. CEV Project tests show that the CM Reaction Control System (RCS) has 

sufficient torque to meet the + 30 degree roll requirement with current parachute harness / riser 

designs. The NESC recommendations from the NESC CEV Water versus Land-Landing 

Assessment endorsed the use of roll control, and the recommendation remains applicable 

because: 

 Some form of roll control is useful to orient the CM Z plane with the direction of travel to 

maximize crew safety. 

 There is limited cabin volume and minimal human impact acceleration tolerance in the Y 

axis (lateral impacts). 

 During water-landings, the CM can be oriented to minimize crew impact accelerations. 

This supports meeting reduced impact acceleration requirements for de-conditioned crew. 

Task 5: Main Parachute Release Times during Land-Landing Summary

The effectiveness of releasing the parachute from the CM at different release times post 

touchdown was examined for a vertical and horizontal landing velocity of 26 and 37 fps, 

respectively. A 37 fps horizontal velocity was used since this velocity represents one of the 

more extreme conditions and is most likely to cause CM rollover. 

For landing conditions where there is a horizontal wind, retaining the parachutes has a 

detrimental effect on CM stability since the drag force on the parachutes can pull the vehicle 

over. The effect of rigging and parachute flexibility has minimal effect on the acceleration and 

roll response so that the trends reported here should be applicable to most parachute system 

designs. Peak accelerations occur early at touchdown where the parachutes have negligible 

effect. Rollover which occurs after touchdown is not significantly affected by the parachutes 

since the parachutes are either slack or have minimal tension, thus applying little or no forces on 

the CM. 

Some form of automated parachute release should be a requirement since in the presence of 

horizontal winds with an attached parachute may cause the CM to be dragged and tumble. This 

is true for both water and land-landing events. 

 An automated system should be required since the release may be required to occur 

within 0.50 seconds of touchdown, which is not sufficient time for a crew operated 

manual release. 
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 An automated release system would be a critical function that must ensure no 

unplanned release events. 

Task 6: Crew Protection System Enhancements Summary 

This task explored the utility of using Finite Element Model (FEM) techniques as a method for 

improving overall understanding of the effects and effectiveness of advanced or modified crew 

restraint systems. While in many aspects simplified, the results of this task indicate that the 

technique is useful and traceable to appropriate and well anchored approaches. This task also 

explored the specific case of incorporating lateral supports to improve the crew’s tolerance to 

high Y-axis accelerations. Simulations indicated that substantial reductions in neck moments, 

cited as being most responsible for severe injury for these sorts of crashes, while showing 

increases in some other forces throughout the body. In particular these simulations indicate that 

substantial improvement in crew safety can be effected, with the caution that care must be taking 

in the detailed design to avoid transferring injurious loads to other parts of the body. Concepts 

were also developed that would allow implementation of suitable restraints. 

The major results of Task 6 included: 

 The Finite Element Analysis (FEA) modeling approach shows promise as a tool for seat 

designers. In this task, lateral supports were studied and could be used to improve crew 

survivability for landings with increased Y-axis accelerations. FEA models can provide a 

design engineering means to understand the effect of these restraints. In addition to 

developing tools, it is important that a practical means exist for implementing design 

solutions. 

 Lateral Seat Support - The use of rigid lateral supports of the type found currently in race 

cars were not directly applicable since these designs would likely be impractical during 

crew ingress and egress. In order for lateral restraints to be used in space flight 

applications, a practical means for implementation with reasonable crew ingress, and 

rapid egress must be developed. In addition the restraint system must be flexible in its 

ability to accommodate crew members of different sizes. During this subtask, several 

lateral restraint concepts were examined. 

 Assuming that the potential for unplanned landings on land cannot be eliminated for a 

CM configured for water penetration or other compromised landing configurations, 

additional crew protection is desirable. 

 The Brinkley Dynamic Response Method criteria define an acceptable environment for 

crews restrained in a specific way. However, it does not provide insight into additional 

or modified crew restraint systems. 

 The use of FEA and anthropomorphic test devices (ATD) tests can allow designers to 

evaluate the effectiveness of alternate restraint systems that can provide crews with 

Brinkley Dynamic Response Method levels of protection in substantially harsher 
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environment. Additional evaluation of industry standard injury criteria should be 

initiated to ascertain the applicability of these criteria to crew protection. 

 Based on the work performed to date, further development and adoption of these 

contemporary design tools and techniques is recommended. 

 Practical methods for implementing additional crew protection appear realistic and 

should be developed. 

Space flight inherently poses risk to the crew during dynamic flight phases. This assessment was 

performed to address specific risk areas identified during the previous study on Water versus 

Land-Landings (NESC Assessment Number 06-020-E) involving the descent and landing phase 

of the mission. New findings and recommendations for the CM parachute architecture, occupant 

protection system, landing attitude, and operational approaches were identified and implemented. 

The Apollo CM remains as a solid reference, which when combined with the NESC team 

experience, allowed the understanding and then improvement on what Apollo had accomplished. 

Finally, additional lessons learned from Apollo Program experience were made to reduce risk as 

part of the T&V program. Given the state of maturity of the CEV Project at this time, those 

recommendations have not yet been implemented. 
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5.0 Ass essment Plan 

5.1 Charter and Background 

The CEV Project Chief Engineer requested that an integrated risk assessment and various 

associated studies be performed on the CEV Integrated Landing System – referred to as the LM 

604 baseline. The Chief Engineer requested the independent study as preparation for the CEV 

Project’s decision to baseline the LRS. 

A brief Phase I analysis was performed by the Water versus Land-Landing Team (NESC 

Assessment Number 06-020-E) to compare the CEV Integrated Landing System defined in the 

LM 604 baseline, against the NASA TS-LRS001 baseline with respect to risk factors potentially 

leading to severe injury or LOC. This assessment covers CM and crew recovery, but does not 

include post-landing operations. 

The Phase I effort resulted in the requestor asking the NESC to perform an in-depth analysis of 

potential risk to the crew in the Integrated Landing System of the LM 606 baseline as described 

in the following section. 

5.2 Scope 

The scope of this task is limited to an evaluation of potential risks found in the LM CEV 

Integrated Landing System as described in LM 606 baseline. Initially, the NESC 06-020-E 

Water versus Land-Landing study evaluated the Integrated Landing System risks to the proposed 

LM 604 CEV baseline. As the CEV design has matured and many design assumptions have 

changed, the NESC was requested to investigate areas of potential risk in the LM 606 baseline. 

The Phase II study focused on areas of risk associated with the flight terminal phase for the CM 

and developed strategies for reducing or mitigating inherent risks. Risk was evaluated in terms 

of risk to the crew during a nominal land-landing. Additionally the CEV Integrated Landing 

System testing approach was evaluated. 

6.0 Description of the Problem and Approach 

The problem description and approach for the Phase I portion are presented in Section 6.1 and 

the Phase II problem description and approach are presented in Section 6.2. Note: each of the 

problems and approaches has a corresponding data analysis section and paragraph number in 

Section 7.0. 

6.1 Phase I - Comparison of LM 604, TS-LRS001, and Apollo LRS – 

Problem and Approach 

The CEV Project Office requested a risk comparison of the LRS proposed by LM with the 

design developed internally by NASA as shown in Figure 6.1-1. 
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Figure 6.1-1. NASA and LM Landing and Recovery Baselines used for Comparison in Phase I 

The primary focus for the Phase 1 task was reliability. Mass, volume, cost, and complexity were 

considered only if an obvious concern existed. The CEV Project’s RAC-3 results were not 

factored into the assessment because this task was run in parallel with RAC-3 activities. A 

qualitative risk assessment was performed because insufficient detail existed for a quantitative 

risk analysis or reliability prediction. 

This assessment relied on existing design and performance data from the NASA (TS-LRS001) 

and LM604 CEV teams primarily from the September – October 2006 timeframe. There were 

no existing formal design documents. Mr. Jim Corliss, Chief Engineer for the CEV Landing 

System Advanced Development Project, provided presentation material regarding landing 

systems from the weekly CEV Government Equipment and Materials (GEM) Office project 

control panel status meetings at JSC. LM data were culled from presentation packages and from 

technical interchange meetings (TIMs) to discuss the status of the LRS design. Additionally, the 

NESC assessment team visited the LM Denver site in October 2006 for a TIM. 

The paradigm for the evaluation procedure was as follows: Entry, Descent, and Landing (postentry) was decomposed into five major sequences, all of which must succeed: 

• Parachute deployment 

• Heat shield unlock or jettison 

• Terminal descent 

• Retrorocket firing 

• Touchdown 
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Each of the five sequences was further decomposed into a set of sub-events or conditions, all of 

which must succeed for the overall sequence to be successful. Each sub-event was evaluated for 

design vulnerabilities with the potential to cause severe injury or LOC. Four categories of design 

vulnerabilities were considered in the qualitative evaluation process. The design vulnerabilities 

were: 

 Brittleness (sensitivity to design assumptions) of performance to flight dynamics, 

environmental factors, or operating conditions. 

 Potential to be a common-cause failure source or susceptibility to a common-cause 

failure. 

 Lack of redundancy coupled with potential reliability issues. 

 Steep drop-off in performance under off-nominal conditions. 

A color code was assigned to each sub-event based on the level of concern relative to the list of 

vulnerabilities. The colors assigned to the sub-events assume that all previous sub-events and 

sequences are successful. Vulnerability concerns by color codes were: 

 Green = low concern (function deemed to be robust). 

 Yellow = moderate concern (design concept may be adequate but robustness 

improvements should be considered). 

 Red = active concern (recommend proactive measures to increase robustness and/or 

change the design). 

 White (no color) = function not applicable to the design. 

Color code assigned to each sub-event based on level of concern relative to the list of 

vulnerabilities as shown in Figure 6.1-3. 
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Figure 6.1-2. Post-Entry Descent and Landing Event Sequence (Functional Decomposition of Main 

Events) for a Successful Landing 

Figure 6.1-3. Vulnerability Concern Color Code: Concern Colors Are Assigned per Sub-Event with 

the Assumption that All Previous Sub-Events and Sequences Are Successful 
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6.2 Phase II -Integrated LRS Risk Assessment – Problem and Approach 

The Phase II study focused on determining the areas of risk associated with the terminal phase of 

flight for the CM and developed strategies for reducing or mitigating risks inherent in this phase. 

A systems engineering approach was used to investigate the overall landing risk by investigating 

multiple areas of risk and suggested strategies for reducing or mitigating the risks. 

Analysis of the problem by the NESC assessment team and working with the CEV Chief 

Engineer resulted in six individual tasks that encompassed specific, but related portions of the 

overall landing system concept. The tasks were developed to assess the risk to the crew of 

operating various vehicle configurations in the anticipated nominal descent and nominal landlanding environments, and investigate specific landing risk mitigation strategies. The overall 

systems approach is depicted in Figure 6.2-1. 

Figure 6.2-1. Systems Approach to the Integrated LRS Risk Assessment 

The tasks and approach are summarized in the following sections. 

6.2.1 Task 1: Entry, Descent, and Landing System Quantitative Risk Assessment Problem 

and Approach 

The NESC assessment team’s objective was to estimate the risk to the crew during landing using 

the LM 606 LRS baseline. The team’s approach was to update the Integrated Risk Matrix from 

the CEV Water versus Land-Landing Study, which provided a quantitative risk assessment of 
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landing. The analysis began with undock through entry, descent, and landing, and included 

various landing configurations (e.g., with and without CM horizontal retrorockets) and site 

conditions. 

Figure 6.2-2. Phases of the Entry, Descent, and Landing Included in the Integrated Risk 

Assessment 

6.2.2 Task 2: Landing System Test and Verification (T&V) Problem and Approach 

The NESC assessment team assessed the CEV Project’s T&V approach with respect to design 

qualification and certification. The original request was to provide an independent assessment 

regarding the CEV Project’s T&V strategy for the CM LRS with respect to design qualification 

and certification. The team found that it was difficult to evaluate strategy since the CEV Project 

T&V planning is still in initial stages. Historically, T&V plans are not developed until the 

Preliminary Design Review (PDR) time frame or after. Early work on T&V planning is also 

problematic because design requirements are in an evolutionary state. 

The team’s opinion was that historical experience should be used to make recommendations 

because the LRS design has heritage from Apollo and Mars exploration programs – while still 

recognizing that the CEV has unique requirements. 

The team restructured the Task 2 objective to develop a recommended set of T&V activities 

based on the team’s experience with similar systems and available knowledge of the efforts to 

date. The goals were to: 

• Develop a set of activities that would comprise a comprehensive entry, descent, and 

landing T&V plan. 

• Provide Project with sufficient T&V information for use as a guide during their T&V task 

planning process. 

• Utilize event-tree methodology to expose the items that must be verified in the absence of 

requirements. 

For each activity, the following was provided: 

• A brief, qualitative description of the activity. 

• Whether the activity should cover exhaustive, bounding, or representative scenarios. 
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• External variables that should be examined. 

• Fidelity of the interfaces to other subsystems or components that interact with the prime 

target of verification. 

6.2.3 Task 3: Landing Site Accessibility and Availability Problem and Approach 

The NESC assessment team’s objective of this analysis was to identify significant trends in 

availability and develop a relative measure of sensitivity for those of special interest. The goal 

was to identify factors that warrant additional attention, not to make probability assessments or 

recommend design or operations criteria. Further, the team was to evaluate implications of 

relying on wind limits to ensure that horizontal velocity at touchdown remains within the CM 

design capability. 

The team addressed the operational implications of implementing various landing scenarios for 

CM return from an ISS orbit. Two quantifiable attributes, accessibility and availability, have 

been defined to guide this evaluation (Figure 6.2-3). 

Figure 6.2-3. Approach to determining Landing Site Accessibility and Availability 

Accessibility is defined as the number of times a given landing site can be physically reached 

from the initial orbit. It reflects the classic Return to Earth (RTE) targeting problem, achieving 

an inertial trajectory that will transition to an earth based coordinate system at a specific time of 

landing and geographic location. This evaluation identifies opportunities to land at a given site 

during a period of interest. 
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• External variables that should be examined. 

• Fidelity of the interfaces to other subsystems or components that interact with the prime 

target of verification. 

6.2.3 Task 3: Landing Site Accessibility and Availability Problem and Approach 

The NESC assessment team’s objective of this analysis was to identify significant trends in 

availability and develop a relative measure of sensitivity for those of special interest. The goal 

was to identify factors that warrant additional attention, not to make probability assessments or 

recommend design or operations criteria. Further, the team was to evaluate implications of 

relying on wind limits to ensure that horizontal velocity at touchdown remains within the CM 

design capability. 

The team addressed the operational implications of implementing various landing scenarios for 

CM return from an ISS orbit. Two quantifiable attributes, accessibility and availability, have 

been defined to guide this evaluation (Figure 6.2-3). 

Figure 6.2-3. Approach to determining Landing Site Accessibility and Availability 

Accessibility is defined as the number of times a given landing site can be physically reached 

from the initial orbit. It reflects the classic Return to Earth (RTE) targeting problem, achieving 

an inertial trajectory that will transition to an earth based coordinate system at a specific time of 

landing and geographic location. This evaluation identifies opportunities to land at a given site 

during a period of interest. 
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iii. inclination 51.63 degree 

b. Edwards Air Force Base (EAFB) wind/gust data for 2006 

2. Multiple site weather from National Oceanic and Atmospheric Administration (NOAA) 

for 2004 [ref. 14]. Geographic data and orbital tracks were derived from Analytical 

Graphics, Inc’s Satellite Toolkit (STK) orbital model 

3. Day-time are initially defined as 0600-1800 local time 

4. Landing opportunities are assumed possible whenever the ISS model flies through a 

cylinder with a radius equal to the CM cross range, projected vertically from the center 

of the landing area (ignores the delta times due to actual reentry timeline) 

5. Toe of SM debris footprint ranges from 370 nm short of landing site in orbital plane with 

no raise maneuver to 1,300 nm with a maximum raise maneuver of 55 fps for L/D 0.35, 

and 230 nm and 1160 nm for L/D 0.30 

6. No attempt was made to assess the suitability of landing sites, capabilities of specific 

CM designs, compatibility of CM designs with landing site surfaces, or availability of 

more accurate wind forecasting tools. 

7. Analysis is focused on land-landings associated with returns from LEO since they are a 

precursor to lunar operations and must accommodate unique flight test and operational 

constraints. 

Availability is defined as the number of times that accessible opportunities can be utilized in the 

presence of operational constraints. The consideration of operational constraints can only reduce 

the number of potential landing opportunities, this study addresses those concerns. Key 

considerations included were: 

 Surface winds and gust conditions 

 Day/night restrictions 

 Ability to ensure SM debris footprints remain off shore 

A deterministic approach was used to determine landing site availability that included: 

• Collect and evaluate 2006 recorded and forecast wind data for EAFB. 

• Identify EAFB landing opportunities for various forecast criteria and evaluate outcomes 

for each using appropriate wind reports. 

• Consider potential correlation among EAFB, Carson Sink, NV, and the Utah Test Range 

wind data. 

• Consider potential flight demonstration build up plan. 

A statistical approach was also used to determine landing site availability that included an 

examination of the weather forecasts (primarily horizontal winds and gusts) versus the measured 

winds exceeding the CM’s horizontal wind landing capability. 

Additional concerns not addressed were: landing surface conditions, terminal area weather 

(snow, rain), and weather forecasting confidence. While these are broad assumptions, their use 

allows for identification of significant trends and sensitivities in key design drivers (i.e., orbital 
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mechanics and weather patterns and may lead to a more detailed evaluation). The complexity of 

this environment and large number of conditions that merit examination, led to the development 

of an analog evaluation tool, Spacecraft Landing Accessibility and Availability Model 

(SLAAM). SLAAM was created in multiple steps starting with using the simulated orbit data 

modeled in AGI’s STK, tracking the times when the orbiting satellite (ISS) would pass over the 

landing areas specified by the CM L/D and exporting this data into Microscoft® Excel ®. This 

data was then sorted and adjusted for local time, ascending or descending pass types, landing 

location, and SM disposal method. Weather data for EAFB was added in a separate database. 

Lastly, the tool mainly allows the user to sort the data based on their desired inputs. This model 

displays data in both daily and daily/hourly calendar format and geometrically accessible landing 

times while allowing for various user-imposed constraints. Those landing opportunities that fail 

the constraint from the display are removed from consideration. Figure 6.2-5 shows an example 

of the SLAAM user interface along with the sample output. 

Figure 6.2-5. SLAAM User Interface 

6.2.4 Task 4: CM Roll Control in Preparation for Landing Problem and Approach 

The NESC assessment team’s objective was to evaluate the requirements and capability of the 

CM roll control concept to orient the vehicle for an acceptable landing. The approach was to 

determine the previous Apollo missions roll control requirements, the current CM requirements 

for the roll control system, and then: 

• Compare Apollo and CEV CM requirements. 

• Determine what the requirements should be on the roll control system, and identify the 

simplest approach using a combination of modeling and simulation, design, and peer 

review. 

• Deployment altitude, descent times, and range of wind conditions including cross winds 

as function of altitude. 

• Analyze the RCS impulse performance and fuel allocation. 

• Identify accepted range of roll alignment requirement (with confidence of specified 

requirement) for landing. 
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• Estimate torque requirement for range of bearing diameters and compare with RCS 

torque output. 

• Size riser length and thickness requirement based on estimated peak deployment loads 

• Estimate oscillation frequencies, magnitude, and dampening of the CM/parachute system. 

6.2.5 Task 5: Investigate Parachute Release Times during Land-Landing Problem and 

Approach 

The NESC assessment team investigated parachute induced effects on CM stability during 

landing near or beyond the vehicle’s horizontal velocity capability. A concern of the CM 

landing is that structural accelerations will be large causing vehicle damage and/or crew injuries. 

Winds at the landing site could exceed the predicted values and thus exceed the CM capability 

(i.e., environment exceeds capability). Also, the CM accelerations could be large causing 

vehicle and/or crew injuries. The parachute effects are thought to have the potential to pull the 

CM over during conditions such as higher winds, or in some cases to stabilize the vehicle by 

preventing the motion after touchdown. 

The effect of releasing the parachutes at different times after touchdown was investigated in 

terms of CM accelerations and rollover in various horizontal wind conditions. A simplified 

parachute model was developed and coupled to a CM structural model. Simulations were 

performed with vertical and horizontal landing initial conditions and horizontal winds expected 

at the landing sites. 

6.2.5.1 Parachute Model 

The LS-DYNA® FEM originally created for the CM was extended to include the parachutes and 

parachute lines. While CM is expected to utilize three primary parachutes for landing, for the 

purposes of this study the effects of the parachutes and lines are simplified by combining the 

effects into a single parachute concept (Figure 6.2-6). The parachute is modeled as a lumped 

mass whose weight is equal to the total expected parachute weights. The connecting lines are 

modeled as a single line with elastic properties and pre-tensioning. The line is preloaded with an 

initial tension equal to the CM weight so that the coupled vehicle-parachute system is in 

equilibrium just before touchdown. The “cable” element available within LS-DYNA® is used to 

model the line so that the line transmits load when the line is in tension and carries no load when 

the line is slack. The weight of the line is distributed along the length and the mass of the 

parachute is concentrated at the top of the rigging. 
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Figure 6.2-6. Conceptual Automatic Parachute Separation on Landing to Prevent CM Tumbling 

(left) and Simplified LS-DYNA® Model (right) 

The parachute force is applied at the location of the parachute mass. The parachute force is 

computed from the simple drag equation: 

F = C x V2 

Where F is the parachute force, V is the velocity of the parachute mass, and C is the parachute 

coefficient. C may be determined from the relationship: 

C = 2

1

 x ρ x Cd x A 

Where ρ is the air density, A is the area of the parachutes, and Cd is the drag coefficient of the 

parachute. An alternative approach to computing the parachute force is to use the equilibrium 

conditions for a nominal landing to compute the coefficient, C. For a nominal landing, the 

parachute force must equal the CM weight so that: 

W = C x V2 

For this assessment, the nominal vertical velocity was 26 fps and the CM weight was 13,046 lbs 

leading to a parachute coefficient of 0.134. For the present assessment, this coefficient was used 

for both the vertical and horizontal directions. This was a reasonable approximation to the actual 

parachute forces since the parachutes tend to orient themselves so that they efficiently oppose the 

direction of the wind and vehicle motions. Currently, tests are being performed on CM 

parachute systems and as this data is processed a better understanding of the parachute forces 

will be defined. However, for the present study this simplified model of the parachute system 
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and forces was sufficient to explore the effects of the parachutes once the CM touches the 

ground. 

The parachute force was implemented into the LS-DYNA® simulation via a series of lookup 

tables for various wind conditions. During each time step in the transient simulation, the 

computed vertical and horizontal velocity at the location of the parachute mass was used with the 

lookup tables to extract a parachute force in each of the horizontal and vertical directions. These 

forces were then applied at the parachute mass location and used to compute the structural 

response for the step time step. When there was a horizontal wind present, the horizontal wind 

force was added to forces generated from the transient motion of the parachute mass. 

The vertical parachute force is defined as: 

Fv = 0.134 x Vv

2

 lbs 

And the horizontal force is defined as: 

Fh = 0.134 x (Vwind

2

 + Vh

2

) lbs 

Where Vv and Vh are the vertical and horizontal velocities of the finite element node at the 

location of the parachute mass. 

Figure 6.2-7 shows a comparison between the parachute forces computed using the identified 

formulation and the vertical parachute force obtained using the parameters supplied by LM in 

memorandum CEV-LRS-07-001. For descent velocities within those expected for the CM, there 

was close agreement between the LM and NESC assessment team’s predictions. At higher than 

expected velocities there is a divergence between the parachute forces since the LM formulation 

used a parachute drag coefficient that increases with descent velocity (thus increasing the drag 

force) while the team’s formulation employed a constant drag. 
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Figure 6.2-7. Comparison of Results of Parachute Forces from Two Different Computational 

Methods 

LS-DYNA® was used to perform the analysis of CM landings with parachute attachments. This 

program was selected because of its ability to simulate the complex transient dynamic behavior 

of the CM impacting a landing surface. The CM model consisted of a collection of structural 

parts as depicted in Figure 6.2-8. The main portion of the CM, which consists of the pressure 

vessel, associated structure, and internal components, was modeled as a rigid part having inertia 

properties equivalent to the LM 604 CM design. The total weight of the vehicle was 13,046 lbs 

which is lighter than the currently projected design, but was an adequate value for the purposes 

of the present study. 
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Figure 6.2-8. CM Structural Parts used in the Model 

Inside of the CM pressure vessel is the crew pallet. The pallet supports the seats and is 

supported by fifteen energy absorbing landing struts. A portion of the CM inertia properties are 

allocated to the pallet (modeled as a rigid part) to account for the crew and seat weights. While 

the pressure vessel and pallet are modeled as structurally rigid and non-energy absorbing, the 

pallet struts are modeled as energy absorbing since they provide the primary source of landing 

load attenuation. 

The parachute and rigging are attached to the CM FEM so that the parachute attachment point 

passes through the center-of-gravity with the vehicle oriented at the desired hang angle shown in 

Figure 6.2-9. The parachute is attached approximately 190 feet from the attachment point. The 

CM model shown includes a triangular shaped parachute used for visualization purposes only as 

the parachute is modeled as a concentrated mass. 

 
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Figure 6.2-9. Overview of the Complete Finite Element Model 

For the purpose of the present study the landing surface is assumed to be a relatively “soft” soil. 

The soft soil model has the effect of lessoning the resulting CM accelerations compensating for 

the fact that most of the vehicle structure is modeled as rigid. A coefficient of friction of 0.60 is 

used to model the contact friction between the CM and soil surface. 

6.2.5.2 Parachute System Elasticity 

Before proceeding with the assessment of the effect of parachute release times, it is important to 

measure the significance of the effect of the parachute system elasticity. For example, if a stiff 

rigging configuration responds significantly differently from a soft system, it is important to 

identify these differences and to either model the system as closely as possible to the actual “to 

be built system”, or to identify differences so that relevant information is available to subsequent 

designers. 

The overall parachute system elasticity is a combination of the elasticity contributed from the 

harness, riser, parachute suspension lines, and the air embedded inside the inflated parachute. To 
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assess the overall effect from these components they may be viewed as a single spring that 

connects the parachute to the CM. When the parachute pulls on the vehicle, the air inside the 

chute compresses and, along with the rigging lines and riser, provides system flexibility. While 

the parachute design may utilize materials such as nylon or Kevlar® which are readily 

characterized, the total system elasticity is more difficult to calculate since it is a combination of 

rigging material, geometry, and aerodynamics. In fact, the compression of the air inside the 

parachute may provide more elasticity than all of the rigging materials combined. 

To provide a reasonable assessment of the system elasticity, or to at least determine if elasticity 

considerations are even a significant issue, three different system stiffness’s were examined for 

their effect of the overall system response. The three stiffness’s are a relatively stiff system 

where the entire parachute/rigging elongates only 1 percent of its length when the CM weight is 

hung from the parachute (i.e. 13,046 lbs hung from a 190 foot parachute system stretches the 

system ~2 feet), a medium system (2 percent elongation), and a relatively soft system where the 

elongation is 5 percent. 

The CM rotation and acceleration resulting from a vertical and horizontal landing velocity of 26 

fps and 37 fps is shown in Figures 6.2-10 and 6.2-11, respectively, for the three levels of 

stiffness. The hang angle is 0 degrees. For all three stiffness’s the parachute is kept attached to 

the CM for the duration of simulation. As depicted in the figures, the CM rolls over regardless 

of the system stiffness, which is an expected response for a 0 degree hang angle and horizontal 

velocity. Furthermore, there is no significant different among the three levels of stiffness. The 

simulation resulting from the softest stiffness does rollover slightly less than the other two 

simulations, but the difference in rollover angle is not large. The accelerations at the location 

where the crew are seated are almost identical regardless of the stiffness level. The medium 

stiffness level was used for subsequent simulations since this level is thought to be the most 

reasonable characterization of the actual stiffness. 
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Figure 6.2-10. Effect of Rigging Stiffness on CM Rollover (0 degree hang, Parachute Remains 

Attached) 

Figure 6.2-11. Effect of Rigging Stiffness on Pallet Acceleration (0 degree hang, Parachute Remains 

Attached) 
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6.2.6 Task 6: Crew Protection System Enhancements Problem and Approach 

During the course of this assessment, tools were developed to: 

• Evaluate enhanced crew restraints and to develop restraint concepts that would enhance 

crew safety during high-g impacts or CM rollover. 

• Investigate the use of FEA techniques to model the biodynamic effects of high-g impacts 

with various seat and restraint systems. 

• Establish correlation between the FEA technique and standard methods (e.g., Brinkley 

Dynamic Response Method Model) for the case of standard restraints. 

• Employ FEA to explore one type of advanced crew restraint system. 

At the conclusion of this assessment, these tools were provided to the CEV Project for use in 

evaluating alternative crew seats. 

The effects of crew restraint options on the LM 604 seat strut and seat pallet concept designs 

using traditional “whole body” dynamic response models, and FEA models of (ATDs (e.g., crash 

test dummies) were investigating. The NESC team: 

• Investigated the basic crew seat configuration (legacy) versus race car seat. 

• Determined if FEA ATD models can be developed into design tool that will be 

predictive for advanced crew protection seats. 

• Performed FEA simulations of impacts in the Y- and Z-axis directions (most sensitive 

for crew injury). 

• Compared FEA to Brinkley Dynamic Response Method and biological data. 

• Developed and compare crew seat impact attenuation improvement concepts. 

• Found opportunity to add additional value that involved validating the FEA model 

with recent sled results run by the CEV Project and the Suit Project at Wright 

Patterson Air Force Base. 

7.0 Dat a Analysis 

The data analysis for the Phase I portion is presented in Section 7.1 and the Phase II data 

analyses are presented in Section 7.2. 

7.1 Phase I - Comparison of 604, TS-LRS001, and Apollo LRS Data 

Analysis 

The NESC team compared the configurations of the reference designs and developed an 

architecture comparison of these designs with the Apollo ELS. Table 7.1-1 lists the points of 

architecture comparison with the Apollo system. 
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Table 7.1-1. Areas of Comparison between the Apollo ELS and the LM and NASA CM LRSs 

The Apollo ELS Subsystem Manager was an integral part of the NESC team and identified the 

areas of design challenge in the Apollo ELS applicable to the LM and NASA LRS baselines. 

A vulnerability concern map that evaluated risk was prepared of sub-events/conditions mapped 

to the five major sequences outlined in paragraph 6.1. Concern colors for each of the reference 

designs were indicated for each sub-event. The 10 concerns, numbered C1 through C10, were 

associated with the sub-event or sub-events. They are: 

C1 (LM 604) - Main Chute Extraction 

C2 (LM 604) - Retrorocket Extraction 

C3 (LM 604) - Winds and Residual Horizontal Velocity 

C4 (LM 604 and TS-LRS-001) - Retrorocket Firing Chain 

C5 (LM 604 and TS-LRS-001) - Apex Cover Separation 

C6 (LM 604) - Retrorocket Plume Impingement 

C7 (LM 604 and TS-LRS-001) - Capsule Orientation Event 

C8 (LM 604) - Drogue Stabilization 
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 C9 (TS-LRS-001) - Impact Survivability 

 C10 (TS-LRS-001) - Heat shield Separation 

An additional level of categorization was also applied to the concerns: 

 1st Tier — Considered the most serious and challenging and has architectural 

ramifications for the design: C2, C3, and C4. 

 2nd Tier — Considered challenging, but fixable. Does not necessarily have cascading 

effects through the architecture: C1 and C5. 

 3rd Tier — Considered to have relatively straightforward fixes: C6, C7, C8, C9, and 

C10. 

Figure 7.1-1 illustrates the vulnerability concern map, the concern colors for each design, the 

numbered concerns associated with entry, descent, and landing sequences, and the 1st, 2nd, and 

3rd Tier concerns. 

 Figure 7.1-1. Vulnerability Concern Map 
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7.2 Phase II - Integrated LRS Risk Assessment Data Analysis 

This section presents the data analysis of the six tasks included in the Phase II assessment. 

7.2.1 Task 1: Data Analysis for Entry, Descent, and Landing System Quantitative Risk 

Assessment Data Analysis 

The purpose of the risk assessment was to identify the risk drivers of different configurations, 

and provide comparisons of the overall risk of the alternatives. This work was an extension of a 

previous NESC assessment that evaluated the risks of landing the CM on water versus land. 

This previous assessment did not analyze specific vehicles and assumed that key risk mitigating 

features were in place. The current assessment extended the previous analysis to evaluate 

specifics of the current CEV design concepts. The risk metric chosen for this analysis was 

probability of LOC (PLOC). 

7.2.1.1 CM Landing System Configurations Assessed 

Several candidate CM configurations were assessed and compared, which include: 

• LM 606 was the baseline configuration in the Spring of 2007. It is a primary landlander with airbags and horizontal retrorockets. It did not meet weight allowances, so 

variations were considered. 

• The ZBV was based on the LM 606, but had numerous parts removed (e.g., landing 

system and some redundant components) in order to meet the weight allowance. This 

was the starting point for alternative configurations assessed. It is also the starting 

point for the PDR point of departure (POD) vehicle design (currently scheduled to be 

approved in mid-November 2009). 

• Water-lander, which is essentially the ZBV with minor accommodations for waterlanding, including roll control. 

• Vertical retrorockets with no horizontal retrorockets on dry soil. 

• Vertical retrorockets with no horizontal retrorockets on wet soil. 

• Vertical and horizontal retrorocket land-lander (soil type does not significantly affect 

outcome), which is equivalent to the LM 606 configuration. 

• Airbags with no horizontal retrorockets on dry soil. 

• Airbags with no horizontal retrorockets on wet soil. 

• Airbags with horizontal retrorockets (soil type does not significantly affect outcome). 

In the ZBV, some multi-level redundancy was removed. For example, some avionics electronics 

was previously multiple fault-tolerant, but was reduced to single fault tolerance (e.g., three 

computers were reduced to two). The NESC team’s calculations determined the reduction in the 

number of computers made a negligible difference because of the used of a common cause 

failure rate of 10 percent of the independent failure rate, so the third unit typically reduces the 

failure rate by less than 1 percent. It was not clear if all manual backup modes were maintained 

in the ZBV, but the team assumed that they were. Similarly, ballistic mode and some trajectories 

were removed from the ZBV. The team did not analyze in detail the risk impact, but instead 

assumed that it would have a minimal impact (i.e., other design features would be used to 
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mitigate new risks that arose from the change). It was also assumed that there would be roll 

control. The importance of roll control for water-landing was not analyzed in detail, but it is 

crucial for land-landings. 

The airbag systems assessed were the Generation 1 airbags by Airborne Systems and the 

Generation 2 biased airbags by Boeing. Boeing’s biased airbags have different pressures in the 

front and rear bags to counteract the center-of-gravity offset, but was not optimized in order to 

maintain acceptable stability at all landing orientations (roll). Generation 2 airbags showed 

improvement over the Generation 1 airbags, thus only the Generation 2 airbags results are 

presented in this study. 

It should be noted that for a given primary land-lander CM configuration, the risk can vary 

widely based on the landing site and landing wind criteria. Section 7.2.3details the analysis that 

was done in selecting landing criteria and the influence of that on landing availability. In this 

section, the primary landing site was EAFB (with backup sites), and the wind forecast go-forlanding criteria was 18 fps. In this case tumbling is the number one entry, descent, and landing 

risk for land-landing configurations except for airbags with horizontal retrorockets. At the same 

time, for day landings with three sites, there would be a 10 percent chance of having to wait 200 

hours to land for an ISS return (i.e., landing availability is not likely to be reduced further in 

order to reduce risk). 

7.2.1.2 Mission Risk Description and Design Impacts 

The risk analysis was focused on the landing systems. The mission segments considered was the 

entry, descent, and landing for an ISS mission. This mission segment is shown in Figure 7.2-1, 

along with risk drivers, and key design features that differentiate the LM 606 design from the 

idealized land-landing design previously analyzed. The reentry risk drivers are: 

 Entry 

o Guidance Navigation and Control (GN&C) System Fault After Entry: This 

failure mode represents failure of the system to deliver the vehicle to the proper 

landing site. If the CM does not land on a relatively smooth, flat surface it may 

be subject to tumbling. This failure mode applies only to land-landing. 

o Thermal Protection System (TPS) Failure: TPS failure could cause catastrophic 

failure to the CM, or cause the structure to become overheated and weakened. 

The weakened structure could cause the vehicle to sink during a water-landing. 

o FBC Release Mechanism: Failure of this mechanism and drogue parachute could 

cause the FBC to interfere with deployment of the drogue parachutes. 

 Parachute Descent 

o Two Drogue Chute Deployment Failure: This failure prevents the initial CM 

deceleration and stabilization to allow for successful deployment of the main 

parachutes. 
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o Two Main Chute Deployment Failures: This failure prevents the CM from 

decelerating to safe limits. This failure could result from system faults, or a 

postulated situation where the drogues are released at a time when there are 

sufficient rates to cause the CM to be inverted when the main parachutes are 

deployed causing failure of the deployment. This concept is being evaluated by 

the CEV Project and is not expected to be in the final CM configuration. 

o Apex Cover Re-contact: Once the apex cover is clear of the CM, it is possible for 

re-contact if its drogue chute does not deploy. 

 Landing Risk Drivers 

o Land-Landing 

 No Heat Shield Separation: The land-landing systems are designed such 

that the heat shield must be jettisoned to allow for operation of the active 

attenuation systems (retrorockets/airbags). Failure of this system will 

cause the capability of the seat attenuation system to be exceeded, or 

make the CM susceptible to tumbling. The idealized vehicle in the 

previous study was assumed to be designed to have a possibility of 

survival. 

 Altimeter or Retro Failure (Retro Systems): This failure results in an 

early, late, or no actuation of the Retro System. Failure of this system 

will cause the capability of the seat attenuation system to be exceeded, or 

make the vehicle CM to tumbling. This type of system requires a narrow 

timing window for actuation (approximately 1 second), with no back-up. 

The idealized vehicle in the previous study was assumed to be designed to 

have a possibility of survival. 

 Airbag Failure to Inflate or Deflate Properly (Airbag Systems): A failure 

in this system will cause the capability of the seat attenuation system to be 

exceeded, or possibly make the CM tumble. Timing of the actuation of 

this system is not as critical as for retrorockets. The idealized vehicle in 

the previous study was assumed to be designed to have a possibility of 

survival. 

 Mispredicted Surface Winds: This failure mode is applicable to systems 

that do not have the capability to remain upright in high surface winds 

(either by design or through active systems such as horizontal 

retrorockets). These systems will typically require that the winds be 

placarded to ensure the capability of the system is not exceeded. This 

failure mode represents the possibility that the weather prediction meets 

the placard, but the actual winds at time of landing can cause the CM to 

tumble. The idealized vehicle in the previous study was assumed to be 

designed to have a possibility of survival. 
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o Water-Landing (Cabin Flooding/high leak rate) 

 Pressure Relief Valve/Vent Failure: This failure would cause the CM to 

take on water and begin t sink. Requiring crew emergency egress if the 

recovery is not prompt. 

 Structural Failure (due to TPS failure): This failure represents the 

possibility of a TPS failure causes overheating (but not destruction) of the 

CM such that it fails during water impact. This failure is assumed to be 

severe enough to cause immediate sinking. 

Another potential risk driver was the contribution of the landing system during a pad abort. 

There is the possibility of a pad abort occurring when there are sufficient winds to blow the 

vehicle back on land. 

Figure 7.2-1. Summary of Risk Drivers for the Nominal ISS Return Mission – Updated from the 

NESC CEV Water Versus Land-Landing Study (Before and After) 

7.2.1.3 Risk Quantification Updates 

Each of the risk drivers that had new information was re-quantified for the LOC end state. The 

overall risk matrix is shown in Appendix A. The results of this are summarized in qualitatively 

in Figure 7.2-2 and Table 7.2-1. Rationale for the quantification is provided in the following 

sub-sections. 
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Figure 7.2-2. Results of the Risk Update 
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Table 7.2-1. PLOC for Risk Driver (Tabular) 

 Vehicle Analyzed 

Updated Risk 

Driver 

WvL Land 

(Vertical, 

Horizontal 

Retrorockets

) 

LM 606 

with 

Airbags 

(dry soil) 

LM 606 

with 

Airbags and 

Horizontal 

retrorockets 

LM 606 

WaterLanding 

Only 

LM 606 

with 

Vertical and 

Horizontal 

retrorockets 

LM 606 with 

Vertical 

retrorockets 

Only (dry soil) 

Pad Abort 8E-5 1.2E-5 1.2E-5 1.2E-5 1.2E-5 1.2E-5 

TPS Failure 2.6E-5 2.7E-5 2.7E-5 2.7E-5 2.7E-5 2.7E-5 

GNC and Other 

Entry 

1.1E-5 1.5E-5 1.5E-5 1.5E-5 1.5E-5 1.5E-5 

Apex Cover 

Separation 

1E-4 2E-4 2E-4 2E-4 2E-4 2E-4 

Heat Shield 

Separation Fault 

0 8.8E-5 8.8E-5 0 8.8E-5 8.8E-5 

Parachute Failure 2.6E-4 2.6E-4 2.6E-4 2.6E-4 2.6E-4 2.6E-4 

Roll Control 

Failure 

1E-6 4.2E-5 4.2E-5 1.1E-5 4.2E-5 4.2E-5 

Retro System 

Failure 

0 0 0 0 2E-4 2E-4 

Airbag System 

Failure 

5E-6 (airbag 

alternative) 

1.6E-4 1.6E-4 0 0 0 

Mispredicted 

Winds 

0 1E-4 0 0 1E-3 1.5E-2 

Structural Failure 

due to Overheat 

(Water-Landing) 

4.4E-5 0 0 2.3E-5 0 0 
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Heat Shield Failure 

Several faults can lead to heat shield failure resulting in LOC. For ISS missions this is 

dominated by Micro Meteoroid Orbital Debris (MMOD) damage during extended stays at the 

ISS. This analysis assumed no covering of the heat shield shoulder region that is above the SM. 

The damage leads to increased turbulence during entry and catastrophic failure (LOC). 

Less severe MMOD damage can lead to weakening of the structure that can cause CM sinking. 

For land-landing, similar damage could lead to overheating the landing system, but this was 

assumed to be less likely. An additional, but less probable cause of sinking is structural failure 

due to insufficient margins. 

Heat shield failure during lunar return was also analyzed, but not included because most risks do 

not differ between lunar and ISS cases and ISS return is more frequent. Due to the high heat 

loads for lunar return. Heat shield failure leading to LOC is more likely with a probability of 

1.3E-4. Sinking due to heat shield overheating is also more likely at 2.6E-4. 

Apex Cover Separation 

Risk of apex cover/FBC separation failure increased from the Water versus Land-Landing 

Assessment, mostly due to new concerns about interference with the drogue chutes. 

Heat Shield Separation 

Heat shield separation faults do not apply to the water-lander because it retains the heat shield. 

The driving failure modes are early separation (which results in LOC in all cases) and no 

separation. In the Water versus Land-Landing Assessment, heat shield failure to separate was 

not considered to cause LOC due to the enhanced seat system, and early separation was not 

considered. In the current study, both cases result in LOC, although early separation was given a 

0.75 weighting due to non-fatal cases (i.e., early separation and cases where the aerodynamic 

loads keep the heat shield in place until parachute deployment). 

Roll Control Failure 

Roll control/RCS thruster alignment failure has the same probability of system failure for the 

Water versus Land-Landing investigation and this assessment. However, in the earlier study the 

system failure did not lead to as high of a PLOC due to the superior seat system. Roll control is 

not as critical for water-landing, but there is some risk due to the asymmetric CM structure 

design and the increased risk of sinking at off-nominal roll angles. 

Retro System Failure 

Vertical retrorocket system failure applies only to that particular system, and is a significant 

discriminator between CM configurations. In this study all retrorocket failures were presumed to 

result in LOC, whereas in the Water versus Land-Landing Assessment the seats mitigated the 

fault completely. 
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Airbag System Failure 

Airbag system failure applies only to that particular system, and is a significant discriminator. In 

this study airbag failures were presumed to lead to LOC, whereas in the Water versus LandLanding Assessment the seats mitigated all failure modes except for the pressure release valve 

failure, which was 95 percent mitigated. 

Mispredicted Winds at Landing Site 

The possibility of choosing to land based on acceptable wind forecasts at the time of the go/nogo decision, but encountering high winds upon landing was not considered during the Water 

versus Land-Landing Assessment. This was partly because the baseline design included 

horizontal retrorockets that would counteract the winds at the time of landing. However, during 

the current study it was found to be the dominant risk in several vehicle configurations. Details 

about the variability of the winds and landing criteria are provided as part of the separate wind 

task; only the effect on PLOC is discussed here. 

The CM’s horizontal velocity at impact is roughly equal to the surface wind speed. Waterlanding is not affected by horizontal winds. Airbag systems have been demonstrated to handle 

higher horizontal velocity without tumbling than vertical retrorocket systems. This was achieved 

by timing the pressure release of the outer bag and shaping the airbags to behave like skids. It 

was also found through simulation that the CM will tumble at a lower horizontal velocity on wet 

than on dry soil. In all cases it was assumed that tumbling results in LOC (unlike Soyuz or the 

Water versus Land-Landing Assessment vehicles with their improved seat attenuation). 

These effects led to the CM configuration with vertical retrorockets, but no horizontal 

retrorockets to have unacceptable PLOC of over 1 percent. Airbag systems without horizontal 

retrorockets are marginal (especially when the landing soil is wet and soft). While the risk could 

be reduced with increased landing constraints, those constraints would probably be unacceptable 

due to the resulting loss of landing opportunities (i.e., there could be many days with no landing 

opportunities). Vertical retrorocket systems with horizontal retrorockets have similar 

performance to airbag systems without retrorockets. In this study, it was found that winds are 

handled well by only the airbag system with horizontal retrorockets and the water-lander. 

Structural Failure Due to Overheat 

Structural failure can lead to sinking of a water-lander, which has a high probability of causing 

LOC in a CM with up to six crew members. Although early Apollo tests resulted in sinking 

under nominal conditions, it is assumed that for the CEV Project the main cause of sinking 

would be TPS failure to meet temperature requirements, leading to a weakened pressure vessel 

and high leak rate upon splashdown. 
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7.2.1.4 Risk Observations and Results 

Role and Benefit of Quantitative Risk Assessment to the Process 

The quantitative risk assessment process goes beyond a reliability and safety analysis and is 

based on an evaluation of the entire mission to identify scenarios that could lead to LOC. This 

analysis integrates hazards and equipment reliability with the physics of the scenarios. It may be 

difficult to validate the estimates generated by this process. However, the discipline of 

developing, justifying, and comparing estimates for internal consistency helps keep the project 

focused on appropriate risk drivers. Opening the analysis beyond the standard safety and 

reliability framework provides the opportunity to analyze situations that go beyond hardware 

failure (i.e., mispredicting winds), and considering the possibility of failures due to design 

problems and unknown unknowns that have occurred in the past (i.e., failure of the retrorocket 

system to actuate in the less than 1 second window). 

The risk assessment made it possible to link the CM design features to the mission. In this way it 

was possible to understand the benefit of crash protection in the context of the landing system 

failure modes. Furthermore the analysis helped expose the difference between a land-landing 

capability (on a smooth, flat, dry surface with low horizontal velocities), and land-landing 

survivability during a pad abort. This distinction helps to make visible the capabilities of the 

land-landing vehicle, and understand the benefit of crash protection for any system type. 

Qualitative Results 

It was recognized that the risk analysis is subject to large uncertainties. In order to provide a 

high level interpretation of the results, the risks were color coded to reflect the rough order of 

magnitude of their likelihood. The risks shown in Figure 7.2-1 are: Before (B) representing the 

risk of idealized landing systems (used in the earlier Water versus Land-Landing Assessment), 

and After (A) representing the LM606 configuration. 

Figure 7.2-1 shows an increase in failure from TPS and apex cover separation. These risks are 

currently under design so they may change. But if the apex cover separation requires a parachute 

system to mitigate the risk, then it may remain high along with the other parachute systems 

(yellow to red risk). 

The risks during descent have remained the same. The landing phase shows a risk increase for 

land-landing systems due to the lack of capability to protect the crew during a hard landing 

(resulting from system failure), or after tumbling (due to mispredicted winds). The mispredicted 

surface winds became red due to the lack of horizontal retrorockets that were assumed to prevent 

tumbling in high wind conditions. 
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Quantitative Results 

The total risk for each CM configuration was computed by adding the PLOC for each risk item. 

This approach is viable because the fault probabilities are relatively independent and low. 

Airbag and vertical retro systems without horizontal retrorockets were evaluated for both wet 

and dry soils due to the significantly different tumbling velocities for the soil types. The results 

are presented in Figures 7.2-3 and 7.2-4. 

Figure 7.2-3. Comparison of All Options (Uses Generation 2 Airbags) 

Mispredicted wind (landing with high horizontal velocity) was found to dominate risk for 

vertical retro systems without horizontal retrorockets. This risk was unacceptable and tended to 

saturate the bar chart used to compare options. After the vertical retro systems without 

horizontal retrorockets were removed, a new bar chart was generated (Figure 7.2-4) to compare 

the remaining options. 
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Figure 7.2-4. Comparisons of Lower Risk Options with Generation 2 Airbags (bottom) 

When using the Generation 2 airbags, failure rates within an order of magnitude (PLOC = 1.8 E3 to 9.4 E-4) indicated that they were relatively close considering the order of magnitude 

uncertainty attached to most of the constituent probabilities. Water-landers had the best risk 

(PLOC = 5.3 E-4) due to their insensitivity to wind, soft landing surface, and simplicity (no 

landing mechanisms or retrorockets). Airbags with horizontal retrorockets have the next best 

risk, and are similar to water-landers, but with some added risk for the landing system. 

Airbag systems without horizontal retrorockets on dry soil and vertical retro systems with 

horizontal retrorockets were found to have similar risk, with similar increases in risk over the 

airbag/horizontal retro system due to having similar (decreased) wind speeds for tumbling. 

However, airbag systems without horizontal retrorockets tumble at a lower wind speed than 

those options on wet soil, and so have higher risk. Because it is likely that most landing sites 

would have wet soil at the same time (during rainy winter months), this risk must be considered 

to be an integral part of the risk for the airbag/no horizontal retrorocket CM configuration. 

Overall, mispredicted winds (high surface winds) were found to be the dominant risk due to 

resultant vehicle tumbling. If this risk is mitigated (e.g., using horizontal retrorockets or 

improved seat systems), all of the options are found to have similar risk magnitude considering 

their uncertainties. 
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7.2.2 Task 2: Landing System Test and Verification (T&V) Data Analysis 

The data analysis effort is summarized in Appendix B. Figure 7.2-5 depicts the events and subevents that were examined. 

Figure 7.2-5. CM Event Sequence Diagram Used to Investigate the Completeness of the CEV 

Project’s T&V Approach 

7.2.3 Task 3: Landing Site Accessibility and Availability Data Analysis 

The land-landing site availability from LEO assessment addressed the interaction of CONUS 

landing accessibility and availability as discussed in the following sections. 

7.2.3.1 Landing Site Accessibility 

Land-landing site accessibility was based on numerous factors. Each of these factors is 

discussed, followed by an integrated graphical representation. A landing opportunity was 

assumed possible when the ISS ground track fell within a specified cylinder projected vertically 

from the center of the landing target. The cylinder radius was equal to the CM cross range 

defined by its anticipated L/D capability. 

CM Lift to Drag Ratio (L/D) 

For this analysis, two L/Ds were under consideration, 0.30 and 0.35. This analysis looked at 

each ratio; the cross range possibilities each presented and determined the sensitivity of L/D 

design. The cross range possibilities for the 0.30 and 0.35 L/D are 60 and 85 nm, respectively. 
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Using a one site landing location as a baseline, this analysis determined how many landing 

opportunities were presented for 180 contiguous days. The cross range capabilities meant that 

the ISS did not need to be directly over a landing site for a landing opportunity to occur, but 

rather the ground track must intersect the cylinder projected vertically from the landing site, thus 

allowing the CM to be guided to the landing site on that orbital track. 

Table 7.2-2 shows the SLAAM output for landing opportunities over a 14 day period, but the 

total passes and average pass per day show the amounts for the180 days calculated by the 

model. This model displays data in both daily and daily/hourly calendar format and all 

geometrically accessible landing times while allowing for various user imposed constraints. 

Those landing opportunities that fail the constraint from the display are removed from 

consideration. In addition, the SLAAM results displayed represent 180-day periods, starting on 

January 1, 2006. It is reasonable to allow for trends in the 180-day data to represent yearly data. 

The SLAAM display in the table uses following nomenclature: 

 ‘a’ – ascending pass 

 ‘d’ – descending pass 

 ‘OK’ – raise maneuver not necessary 

 ‘R’ – raise maneuver necessary 

Table 7.2-2. Comparison of L/D 0.30 and 0.35 for EAFB (Edwards) landing site, all options 

available (day and night), 180 days (first 14 days shown) 

L/D = 0.35 

Start Date 1/1/2006 

Number of Days 180 

Site Name Day 1 2 3 4 5 6 7 8 9 10 11 12 13 14 

Edwards Pass 1 dR aOK dR aOK aOK dR dR aOK

Edwards Pass 2 

 Day Total 0 1 1 1 1 0 1 0 0 0 1 0 1 1 

 # of Landings Total Passes 76 

 Ave Pass Per Day 0.4 

 

L/D = 0.30 

Start Date 1/1/2006 

Number of Days 180 

Site Name 1 2 3 4 5 6 7 8 9 10 11 12 13 14 

Edwards Pass 1 dR aOK dR dR 

Edwards Pass 2 

 Day Total 0 1 0 0 1 0 0 0 0 0 1 0 1 0 

 # of Landings Total Passes 56 

 Ave Pass Per Day 0.3 
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Decreasing the L/D by 0.05 (from 0.35 to 0.30) decreases the CM cross range capability from 85 

to 60 nm with a corresponding reduction in the amount of landing opportunities from 76 to 56, a 

loss of approximately 26 percent. Some sites are available twice per 24 hour period (Pass 1 and 

Pass 2) while on some days the site is not available due to orbital mechanics and the CM cross 

range capability (Figure 7.2-6). An additional note is that some land-landing sites require the 

raise maneuver, shown in Figure 7.2-7. The reentry trajectory targets a SM disposal area off the 

US western coast. After SM jettison, the CM must perform an RCS burn lasting up to 50 

seconds to extend the trajectory downrange to clear the intervening mountains and achieve the 

desired landing site. 

Figure 7.2-6. Problem geometry: illustration of cross range options, ascending and descending 

passes and 3 orbit ground trace 
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Figure 7.2-7. Concept of Operations: CM Disposal and Raise Maneuver 

Ascending or Descending Passes 

The latitude of the landing site greatly influences the ascending/descending pass time spread. 

The closer the landing site to the equator, the more evenly spaced the landing opportunities 

become and conversely, the further toward either pole, the shorter the time between ascending to 

descending passes. For most of the sites chosen for this study, the ascending/descending time 

spread is approximately 16/8 hours (i.e., 16 hours between descending to ascending passes, and 8 

hours between ascending to descending passes). While this aspect of orbital geometry may not 

directly influence landing accessibility, whether or not a pass is ascending or descending might, 

depending on what occurs in each of those scenarios. For example, there may be landing sites 

where all the ascending passes mean a particular SM deorbit procedure is required, therefore 

sorting on ascending or descending was deemed a necessary study tool. The model shows that 

choosing to sort for either ascending only or descending only passes results in a 50 percent 

reduction of landing opportunities. 

Landing Network 

As shown previously in Table 7.2-2, having one landing site has significant limitations. One 

relatively simple idea to increase landing accessibility is to have a network of landing sites. 

Table 7.2-3 displays the results for a representative network of three US western coast landing 

sites: Carson Sink, NV (Carson), EAFB (Edwards), and the Utah Test Range (Utah). Having 

two additional landing sites increases accessibility over one site by 241 percent. 
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Table 7.2-3. Landing accessibility for NASA’s landing site network, all landing options available, 

180 days. 

Start 

Date 1/1/2006 

Number 

of Days 180 

Site 

Name 1 2 3 4 5 6 7 8 9 10 11 12 13 14 

Carson Pass 1 aOK aOK aOK 

Carson Pass 2 dOK dOK dOK 

Edwards Pass 1 dR

aO

K dR aOK aOK dR dR aOK

Edwards Pass 2 

Utah Pass 1 aR dR dR dR aR aR aR dR

Utah Pass 2 

 

Day 

Total 1 1 2 1 2 2 2 3 0 3 1 1 1 2 

 

Total 

Passes 259 

 

Ave 

Pass Per 

Day 1.4 

7.2.3.2 Landing Site Availability 

As mentioned previously, landing site availability is the number of times that accessible landing 

opportunities can be utilized in the presence of operational constraints. The operational 

constraints studied were SM deorbit, day/night landing preferences, wind velocity threshold, gust 

conditions, and the use of water as a CONUS landing site backup is described in the following 

sections. 

Day/Night Constraints 

In certain circumstances and locations, landing either at night only or day only may be 

preferable. Table 7.-4 shows the differences for day and night only landings, with day and 

night-time each containing 12 hours. For January of 2006, day and night also coincide with 

descending and ascending passes respectively, which is not the case throughout the year. For 

example, in June 2006 day only passes for EAFB opportunities occur on both ascending and 

descending passes. 

Table 7.2-4. Landing site availability for Edwards with day landing only constraint 

Start Date 1/1/2006 

Number of Days 180 

Site Name 1 2 3 4 5 6 7 8 9 10 11 12 13 14 

Edwards Pass 1 dR dR dR dR 

Edwards Pass 2 

 Day Total 0 1 0 1 0 0 0 0 0 0 1 0 1 0 

 Total Passes 38 

 

Ave Pass Per 

Day 0.2 
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Surface Wind Constraints 

Past studies indicated that landing systems were sensitive to surface wind speeds. In this 

analysis, the objective was to study landing sensitivity to various wind speed thresholds, diurnal 

and seasonal effects, and site to site wind correlations 

The data for the threshold sensitivity, diurnal, seasonal studies were obtained from the EAFB 

staff meteorologist and included observation and forecasted data for each hour recorded and gust 

conditions. The data for the site correlation study came from Jet Propulsion Lab (JPL) and 

NOAA. 

Diurnal Study 

The diurnal study was completed by examining the hourly wind speed observation data for 

EAFB in 2006, sorting the observations into six 4-hour time blocks, and taking the average wind 

speed for each time block for the year. This study indicated that there are diurnal effects at 

EAFB and the 1600 – 1900 time block has the highest average wind speed. Final results are 

shown in Figure 7.2-8. It is worth noting that while diurnal effects were seen, there were many 

instances of wind variability at all times of day. This study has shown that avoiding certain 

times of day for landing due to diurnal effects is not an effective remedy to the wind situation. 

Figure 7.2-8. Diurnal wind study results, Edwards 2006 

Seasonal Wind Study 

This study examined the recorded data for seasonal variations. Data for 2006 in the 1600-1900 

time block was used. The year was broken into seasons by month and the average seasonal wind 

speeds were compared with the average yearly wind speed for the chosen time block. The 

5.8

4.7

5.35

7.7

10.4

7.6

0

2

4

6

8

10

12

000-300 400-700 800-1100 1200-1500 1600-1900 2000-2300

Time Block (hours)

Wind Speed (knots)

Average Wind
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seasonal study also indicated that while the average wind speeds tend to be benign, the data 

shows that averages can be misleading. Figure 7.2-9 shows the results for summer 2006. See 

Appendix B for all seasons. Because this analysis is only for the 1600-1900 time block, the 

annual average differs from the total annual average. 

Figure 7.2-9. Seasonal Study: Summer 2006, EAFB 

Wind Threshold Sensitivity Study 

The purpose of the wind threshold sensitivity study was to determine what happens to landing 

availability when winds above certain thresholds are eliminated. SLAAM has the capability to 

take wind threshold as a user input and sort accordingly. As the wind threshold increases, 

landing availability increases. Table 7.2-5 through 7.2-8 through show the results for 5, 10, 15 

and 20 knot thresholds at EAFB for 180 days. 

Summer

0

5

10

15

20

25

30

0 50 100 150 200 250 300 350

Observation Number

Wind Speed (knots)

Wind Speed (knots)

Average Speed (knots)

June, July, August 2006 

1600-1900 Average Wind Speed Summer = 12.80 (knots) 

1600-1900 Time Block 

Annual average = 10.4 knots
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Table 7.2-5. 5 knot wind threshold (8.4 fps) 

Start Date 1/1/2006 

Number of 

Days 180 

Site Name 1 2 3 4 5 6 7 8 9 10 11 12 13 14 

Edwards Pass 1 aOK dR aOK aOK dR dR aOK

Edwards Pass 2 

 Day Total 0 0 1 1 1 0 1 0 0 0 1 0 1 1 

 Total Passes 40 

 

Ave Pass Per 

Day 0.2 

Table 7.2-6. 10 knot wind threshold (16.8 fps) 

Start Date 1/1/2006 

Number of 

Days 180 

Site Name 1 2 3 4 5 6 7 8 9 10 11 12 13 14 

Edwards Pass 1 dR aOK dR aOK aOK dR dR aOK

Edwards Pass 2 

 Day Total 0 1 1 1 1 0 1 0 0 0 1 0 1 1 

 Total Passes 55 

 

Ave Pass Per 

Day 

0.

3 

Table 7.2-7. 15 knot wind threshold (25.3 fps) 

Start Date 1/1/2006 

Number of 

Days 180 

Site Name 1 2 3 4 5 6 7 8 9 10 11 12 13 14 

Edwards Pass 1 dR aOK dR aOK aOK dR dR aOK

Edwards Pass 2 

 Day Total 0 1 1 1 1 0 1 0 0 0 1 0 1 1 

 Total Passes 69 

 

Ave Pass Per 

Day 

0.

4 

Table 7.2-8. 20 knot wind threshold (33.7 fps) 

Start Date 1/1/2006 

Number of 

Days 180 

Site Name 1 2 3 4 5 6 7 8 9 10 11 12 13 14 

Edwards Pass 1 dR aOK dR aOK aOK dR dR aOK

Edwards Pass 2 

 Day Total 0 1 1 1 1 0 1 0 0 0 1 0 1 1 

 Total Passes 75 

 

Ave Pass Per 

Day 

0.

4 
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Figure 7.2-10 shows how the cumulative wind speed capture percentage compares with the wind 

speed at EAFB for 2006. The data shows that if a CM is designed to withstand a wind speed 

listed on the y-axis, one can note the percentage of observations, based on the year’s wind data 

that the vehicle could be expected to accommodate. 

Figure 7.2-10. Percent Capture versus Wind Speed (knots) 

Because wind gusts can cause CM oscillations, this analysis investigated what happens to 

landing availably with gust conditions. Using gust data for the same time period (EAFB 2006), 

SLAAM was used to determine landing availability based on whether there were or were not 

gust conditions only. Shown in Table 7.2-9 are the results for gust elimination at EAFB for 180 

days. 

Table 7.2-9. Gust elimination, No wind threshold, EAFB 2006 

Start Date 1/1/2006 

Number of 

Days 180 

Site Name 1 2 3 4 5 6 7 8 9 10 11 12 13 14 

Edwards Pass 1 dR aOK dR aOK aOK dR dR aOK

Edwards Pass 2 

 Day Total 0 1 1 1 1 0 1 0 0 0 1 0 1 1 

 Total Passes 63 

 

Ave Pass Per 

Day 

0.

4 

Wind threshold and gust elimination can have a significant effect on landing availability. Wind 

thresholds can decrease availability by up to 49 percent with a 5 knot threshold. Gust 
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elimination alone can decrease availability by 17 percent. Due to the variability of gusts and the 

unpredictable nature of gusts elimination, an operational constraint may be necessary. 

Forecast Versus Actual Wind 

The differences between actual and forecast wind conditions are worth noting. This assessment 

looked at both forecast and actual data for 2006 at EAFB. In all, there were over 8,500 hourly 

wind and gust observations with two daily forecasts. The actual wind speed exceeded the 

forecast approximately 1,200 times and there were approximately 200 occasions where there 

were non-forecast gust conditions or gusts that exceeded the forecast. This showed that forecasts 

tend to be more conservative, but actual wind and gust conditions frequently exceed the forecast. 

Due to the nature of gust and wind recording methods, there may be situations where gust 

conditions exist or various wind speeds occur, but are not recorded. Reported wind data is the 

latest two-minute average wind speed for the previous hour, and gust data is the highest twominute wind speed average for the last 10 minutes of the previous hour. This data reporting 

method shows how vital information can be non-recorded. 

Wind/Gust Offset 

Because of the nature of weather forecasting and the methods for determining wind and gusts 

speeds, a data-offset capability was introduced into the model. This offset allows the user to use 

the highest value of either wind and/or gust velocity, forecast or actual data for their analysis 

based on time. For example, if a user wanted to use the tool to determine landing opportunities 

with a wind threshold, but wanted to include information on the wind forecast from six hours 

before landing they would have that option. The examples in Tables 7.2-10 and 7.2-11 show the 

impact of using the offset option. 

Table 7.2-10. Offset Example, 20 knot threshold, no pre- or post-offset, forecast data 

Start Date 1/1/2006 

Number of Days 180 

Data Used LD_.35 

Site Name 1 2 3 4 5 6 7 8 9 10 11 12 13 14 

Edwards Pass 1 dR aOK dR aOK aOK dR dR aOK

Edwards Pass 2 

 Daily Availability 0 1 1 1 1 0 1 0 0 0 1 0 1 1 

 Total Passes 74 

 Ave Pass Per Day 0.4 
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Table 7.2-11. Offset Example, 20 knot threshold, 6-hour-pre- and 3-hour-post-offset, forecast data 

Start Date 1/1/2006 

Number of Days 180 

Data Used LD_.35 

Site Name 1 2 3 4 5 6 7 8 9 10 11 12 13 14 

Edwards Pass 1 dR aOK dR aOK aOK dR dR aOK

Edwards Pass 2 

 Daily Availability 0 1 1 1 1 0 1 0 0 0 1 0 1 1 

 Total Passes 68 

 Ave Pass Per Day 0.4 

Due to the offset uses for the highest wind value for the time period specified, the result is a 

more conservative prediction of availability. The offset works similarly for the gust criterion. 

For more information about the offset option, see Appendix B. 

Site-To-Site Correlation 

Based on top level statistical analysis, this study showed that there is some correlation between 

the three previously discussed land network sites: Carson Sink, NV; EAFB, and the Utah Test 

Range. The correlation coefficients for the three pairs are listed below: 

1. Wind speed correlation for Carson and Edwards = 0.5460 

2. Wind speed correlation for Carson and Utah = 0.5039 

3. Wind speed correlation for Edwards and Utah = 0.5071 

A correlation of 0.5 indicates a generally positive relationship between wind speeds at both 

locations. This means that the plot of wind speed at one location versus wind speed at the other 

location tend to fall along a line with positive slope. The closer the correlation coefficient is to 1, 

the closer the points in the plot will fall on a straight line. So, if the wind speed is high at 

Carson, a similarly high wind speed is expected at EAFB. This is not to say that a lower wind 

speed at EAFB is not possible, but the probability of a lower wind speed is smaller than a higher 

wind speed. Table 7.2-12 shows the 2006 site winds for Carson Sink, NV, EAFB, and the Utah 

Test Range, respectively. 
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Table 7.2-12. 2006 Winds, Reported in Knots, at Representative Landing Sites 

SM Deorbit Method Constraint 

In order to reduce the likelihood of SM debris hitting land, sorting for raise and non-raise 

maneuvers was incorporated into SLAAM. Using the SM debris footprint, Table 7.2-13 was 

developed to classify the ascending and descending passes for each site based on L/D and SM 

debris footprint. “Raise” means a raise maneuver is necessary for landing and “okay” means no 

raise maneuver is necessary. 

Table 7.2-13. Raise maneuver classification 

Site Name Approach L/D 0.30 L/D 0.35 

Carson Sink, NV 

Ascending Raise Okay 

Descending Raise Okay 

EAFB 

Ascending Okay Okay 

Descending Raise Raise 

Utah Test Range 

Ascending Raise Raise 

Descending Raise Raise 

Eliminating opportunities that require raise maneuvers reduces availability at EAFB by 51 

percent as shown by the loss landing opportunities shown in Table 7.2-14. 

 2006 Winds, Reported in Knots, at Representative Landing Sites

CAR EDW UTTC

6 Ave 7 Ave 6 Ave

33 Max 29 Max 38 Max

33 Percentle 100% 29 Percentle 100% 38 Percentle 100%

15 95% 17 95% 17 95%

12 90% 15 90% 14 90%

10 85% 14 85% 11 85%

9 80% 12 80% 9 80%

8 75% 11 75% 8 75%

7 70% 10 70% 8 70%

6 65% 9 65% 7 65%

6 60% 8 60% 6 60%

5 55% 7 55% 6 55%

5 50% 6 50% 5 50%
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Table 7.2-14. Availability at EAFB without Raise Maneuvers 

Start Date 1/1/2006 

Number of 

Days 180 

Site Name 1 2 3 4 5 6 7 8 9 10 11 12 13 14 

Edwards Pass 1 aOK aOK aOK aOK

Edwards Pass 2 

 Day Total 0 0 1 0 1 0 1 0 0 0 0 0 0 1 

 Total Passes 37 

 

Ave Pass Per 

Day 0.2 

This analysis also quantified what impact the SM deorbit method constraint had on the landing site 

network. Eliminating opportunities that require raise maneuvers for the land-landing network 

decreased the availability by 52 percent, shown in Table 7.2-15. 

Table 7.2-15. Availability at Carson Sink, NV. EAFB, and the Utah Test Range without Raise 

Maneuvers 

Statistical Analysis of Landing Site Data to Determine Site Availability 

Landing site availability was determined statistically by an examination of the weather forecasts 

(primarily horizontal winds and gusts) versus the measured winds exceeding the vehicle capability. 

Weather forecast data from EAFB was used as input to a Monte Carlo analysis. Figure 7.2-11 

shows a flow chart of the analysis. 

Start Date 1/1/2006

Number of Days 180

Data Used LD_.35

Site Name 1 2 3 4 5 6 7 8 9 10 11 12 13 14

Carson Pass 1 aOK aOK aOK

Carson Pass 2 dOK dOK dOK

Edw ards Pass 1 aOK aOK aOK aOK

Edw ards Pass 2

Utah Pass 1

Utah Pass 2

Daily Availability 0 0 1 01212020001

Total Passes 124

Ave Pass Per Day 0.7
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Figure 7.2-11. Monte Carlo Simulation Flow Chart for Statistical Weather Analysis 

The simulation began by setting the parameters such as the landing, critical, and catastrophic 

criteria. The criteria determined if the CM would attempt to land in the estimated wind 

conditions, and whether the vehicle landed nominally under the experienced wind conditions. If 

the forecasted wind speed is lower than the landing criteria, then the CM could attempt to land. 

If the actual wind speed is higher than the critical criteria, then the CM could experience tipping, 

and if the actual wind speed is higher than the catastrophic criteria then the CM will tumble and 

LOC was assumed. 

Different values correspond to different CM configurations. The baseline landing criteria used 

for previous studies was 18 fps, according to the operational number presented by Marshall 

Space Flight Center (MSFC) in May. The baseline CM configuration used was a vehicle with 

airbags, no horizontal retrorockets, and landing on dry soil. The corresponding critical criteria 

was 38 fps, and the catastrophic criteria was 47 fps. 
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Another parameter was wind gust chance. The gust data used in the studies only gave the 

maximum gust speed experienced during a ten-minute period in an hour, and no information 

about gust length or frequency. Thus, if conditions were gusty, the gust chance parameter 

represented the probability that a gust would happen during the critical landing period. The 

baseline value used in the studies was 20 percent, which is an approximation based on the 

characterization of gusts as roughly sinusoidal. This 20 percent number meant that landing 

vehicle was estimated to experience gusty wind speeds rather than average wind speeds 20 

percent of the time. 

The user could also select the available landing sites. Cases with one site used EAFB, while 

cases for three sites used EAFB, Carson Sink, NV, and the Utah Test Range. The NESC team 

only had gust data for EAFB and was used as a reasonable proxy at low fidelity for the two other 

sites. The other sites provide additional landing opportunities based on orbital analysis. 

The user was able to select the forecast time. This is the length of time prior to touchdown that 

the go/no-go landing decision has to be made. Thus this describes how old the wind forecast 

data will be by the time the landing takes place. This value is typically a product of the 

operational setup, and the value of three hours was used as a baseline. This means that the 

landing decision will be made based on the forecast in effect three hours before the scheduled 

landing. 

After framing the problem, the simulation performed runs up to the specified amount, usually 

about 106 runs. For each run the simulation selects a random day to be the first day in which a 

landing is desired. The next landing opportunity is determined from a database of orbital 

accessibility times (assuming a CM L/D of 0.3). Then a set of wind data is chosen randomly 

from the wind speed database according to time of day and time of year of the landing 

opportunity. The data is binned by hour and month, but in order to increase the number of data 

points to choose from the simulation can select from data sets one hour earlier or later; and from 

data sets one month earlier or later. Thus for a landing at 13:00 in May, the simulation would 

select data from the hours of 12:00, 13:00, and 14:00, and from the months of April, May and 

June. Each set of wind data includes the forecasted maximum wind speed, the actual observed 

average wind speed from the end of the hour, and the maximum gust speed in that hour. The 

forecasted data, actual data, and the gust data come from the same hour. 

The forecasted wind speed is then compared to the landing criteria speed. If the forecasted speed 

is too high then the landing is delayed until the next landing opportunity. If the forecasted speed 

is acceptable, then the landing is attempted. To determine the outcome of the landing, the actual 

wind speed (or gust speed depending on the gust chance) is compared to the critical and 

catastrophic criteria. If those criteria are not exceeded then it is a nominal landing, and if they 

are exceeded, then the landing is critical or catastrophic. The data is recorded and the next run is 

completed. 
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After completing all runs, the probabilities of nominal, critical, and catastrophic outcomes can be 

examined. Also the probability of landing delay, and the total delay times for each run can be 

analyzed. 

Critical and catastrophic criteria were chosen for several CM configurations seen in Table 7.2-

16. These values are for a three hour forecast time (i.e., a forecast in effect three hours before 

the scheduled landing) and 20 percent gust chance (with 0 percent gust chance in parentheses). 

The landing criteria was kept constant for comparison. The tumbling capabilities affect the risk 

directly, with higher capability being better. 

Table 7.2-16. Results from the Monte Carlo Weather Analysis 

Air-Bag Dry Soil

(no horizontal 

retrorocket) 

Air-Bag Wet 

Soil (no 

horizontal 

retrorocket) 

Air-Bag 

(with horizontal 

retrorocket) 

Retro Dry Soil 

(no horizontal 

retrorocket) 

Retro Wet Soil 

(no horizontal 

retrorocket) 

Retrorocket 

(with 

horizontal 

retrorocket) 

Go for Landing 

Forecast Criteria, 

fps (kts) 

18 

(11) 

18 

(11) 

18 

(11) 

18 

(11) 

18 

(11) 

18 

(11) 

Actual Landing 

Tipping Criteria, 

fps (kts) 

38 

(22) 

30

(18)

54

(32)

22

(13)

14

(8.5)

38

(22)

Actual Landing 

Tumbling 

Criteria, fps (kts) 

47 

(27) 

38 

(22) 

61 

(36) 

30 

(18) 

22 

(13) 

47 

(27) 

Probability of 

Tipping, Percent 

0.3 

(0.2 - no gusts) 

1.1 

(1.1 - no gusts)

0.04 

(0 - no gusts) 

2.8 

(2.8 - no gusts)

13 

(13 - no gusts) 

0.3 

(0.2- no gusts)

Probability of 

LOC, Percent 

(Tumbling 

Criteria 

Exceeded) 

0.1 

(0.04 - no gusts) 

0.4 

(0.2 - no gusts)

0.01 

(0 - no gusts) 

1.5 

(1.3 - no gusts)

4.3 

(4.1 - no gusts) 

0.1 

(0.04- no 

gusts) 

The Monte Carlo was run for a range of landing criteria in order to examine the sensitivity of the 

landing delay chance and crew risk to that parameter. The assumptions were airbags, no 

horizontal retrorockets, dry soil (38 fps critical, 47 fps catastrophic), 20 percent gust chance, 

three possible landing sites, and a three hour forecast. The results are in Figures 7.2-12 and 

7.2-13. 

The baseline case is 18 fps for comparison. The PLOC is at about 1e-3 for the baseline, and goes 

above 1e-2 at about 50 fps landing criteria. The chance of abort is 47 percent at 18 fps. 

As the landing criteria value is increased, the CM is allowed to land under rougher conditions. 

This decreases the chance of landing delay while increasing the chance of crew injury. As the 

landing criteria value is decreased, the operational procedure becomes stricter. The chance of 

delayed landing goes up since it is less likely the conditions will be acceptable, but the risk to the 
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crew goes down since the winds would have to exceed their forecast to a greater degree to be 

dangerous. Thus this is a trade between landing availability and crew injury risk. 

To decrease risk without decreasing availability, it is necessary to improve forecasting to avoid 

being surprised by high winds or to increase the CM capability. The CM capabilities were 

varied to examine the results. The assumptions were landing criteria of 18 fps, 20 percent gust 

chance, and a 3 hour forecast. The results are in Figure 7.2-14. 

Figure 7.2-14 illustrates that increasing the CM capability decreases risk. When the tumbling 

criteria was increased, the CM sees more severe winds in order to roll and injure its occupants. 

When the tumbling critieria is decreased, the CM will be dangerous in a greater range of wind 

speed. The baseline tumbling criteria for an airbag system without horizontal retrorockets 

landing on dry soil is 47 fps. 

Figure 7.2-12. Landing delay risk - Chance of abort is 47 percent for the 18 fps baseline case 

Landing Abort Probability

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

0 20 40 60 80 100

Landing Criteria Wind Speed (fps)

Probability

Landing Abort
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Figure 7.2-13. Logarithmic plot of risk of LOC over criteria 

Figure 7.2-14. Logarithmic plot of risk of LOC over capability 

Another interesting output was the amount of delay before landing for each run. The time 

between the first available landing opportunity and the actual landing was recorded for every 

Off-nominal Landing Probability (Airbags, Dry)

0.0001

0.001

0.01

0.1

1

0 20 40 60 80 100

Landing Criteria Wind Speed (fps)

Probability

Critical

Catastrophic
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run, with some landing immediately and some being repeatedly delayed (these delays being 

independent). Due to orbital tracks, these attempts can be separated by days or weeks. This data 

can be visualized as a cumulative probability chart, where each point in the chart gives the 

probability that the CM would have landed with that much delay or more. The assumptions used 

are a CM with an L/D of 0.3 (which effects site accessibility), 20 percent gust chance, and three 

hour forecast. Some of the characteristics that were varied were time of day (any time or 

daytime only), number of sites (one or three), and landing criteria. The results are in 

Figures 7.2-15 through 7.2-19. 

For the case of 1 site, 18 fps, and any time landing, there is an approximate10 percent chance 

that the CM will not have landed after 10 days. If 3 sites are used, then the change of landing 

drops slightly more than an order of magnitude. For the case of 1 site, 18 fps, and daytime 

landing only, there is an approximate10 percent chance that the CM will not have landed after 25 

days. Allowing three landing sites changes with 90 percent confidence time to 10 days. For the 

strictest case (1 site, 13 fps, daytime only) it takes over 40 days of landing delay to achieve 90 

percent confidence that the CM would have landed. 

Ways to improve these values and lessen the landing delay are to increase the landing criteria, 

allow landing under a wider variety of conditions, or to increase the accessibility of landing sites 

by increasing the number or improving the CM landing range. 

Figure 7.2-15. Landing delay cumulative probability, 1 site, 18 fps 
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Figure 7.2-16. Landing delay cumulative probability, 3 sites, 18 fps 

Figure 7.2-17. Landing delay cumulative probability, 1 site, 18 fps, daytime 
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Figure 7.2-18. Landing delay cumulative probability, 3 sites, 18 fps, daytime 

Figure 7.2-19. Landing delay cumulative probability, 1 site, 13 fps, daytime 

Another Monte Carlo simulation was run to judge the effectiveness of the Generation 2 airbags. 

The results are summarized in Table 7.2-17. The PLOC for an airbag equipped CM on soil is 

near zero in this simulation when compared with Table 7.2-16. 
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Table 7.2-17. Monte Carlo Results Incorporating Increased Capability of Generation 2 Airbags 

Shown in Figure 7.2-20 is an update risk trade of landing opportunities versus risk. An analysis 

of the Generation 2 airbag system resulted in using 70 fps for the tumbling criterion on hard soil, 

and 50 fps for the tumbling on soft soil. The CM tipping criteria was set to 10 fps less than the 

tumbling criteria based on CEV Project provided data. The plots in Figure 7.2-20 were based on: 

Generation 2 airbags, hard soil, 20 percent gust, three sites, day landings only, and a three hour 

forecast. As a result, the 18 fps ‘Go for Landing’ criterion resulted in and average of two 

opportunities per week, and a 1E-3 PLOC. 

0.1%

(0.04%- no gusts)

4.3%

(4.1% - no gusts)

1.5%

(1.3% - no gusts)

0.1%

(0.02% - no gusts)

0%

(0% - no gusts)

Probability of

Tumbling

2.8%

(2.8% - no gusts)

30 FPS

(18 KTS)

22 FPS

(13 KTS)

18 FPS

(11 KTS)

Retro Dry Soil (no

horizontal

retrorocket)

0.3%

(0.2%- no gusts)

47 FPS

(27 KTS)

38 FPS

(22 KTS)

18 FPS

(11 KTS)

Retrorocket

(with horizontal

retrorocket)

13%

(13% - no gusts)

0.3%

(0.2% - no gusts)

0.01%

(0% - no gusts)

Probability of

Tipping

22 FPS

(13 KTS)

50 FPS

(30 KTS)

70 FPS

(41 KTS)

Actual Landing

Tumbling Criteria

14 FPS

(8.5 KTS)

40 FPS

(24 KTS)

60 FPS

(36 KTS)

Actual Landing

Tipping Criteria

18 FPS

(11 KTS)

18 FPS

(11 KTS)

18 FPS

(11 KTS)

Go for Landing

Forecast Criteria

Retro Wet Soil (no

horizontal

retrorocket)

Air-Bag Soft Soil

(no horizontal

retrorocket)

Air-Bag Hard Soil

(no horizontal

retrorocket)
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Figure 7.2-20. Updated Landing Opportunities versus Risk 

7.2.3.3 Water-Landing Possibilities 

The purpose of this portion of the analysis was to determine how water-landing availability could 

work with accessibility. In addition, this analysis was used to determine if and how water could 

be used as a backup to land-landings. 

Water Accessibility 

Water-landing accessibility is similarly based on geometry. However, choosing a location that 

maximized the opportunity for landing is easier in the water due to ship mobility and lack of SM 

debris constraints. Day/night constraints may still play a part. As an assumption to the water 
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accessibility study, a ship was placed strategically to support an ascending and descending 

landing opportunity. Also, it was assumed that this ship could travel at a speed of 15 knots. The 

ship’s motion, in combination with the CM cross range capability meant a large surface area 

would be considered for landing opportunities as shown in Figure 7.2-21. 

Figure 7.2-21. Water-landing illustration 

For example, a ship could be placed at the intersection of the magenta passes and if Opportunity 

1 and Opportunity 2 were missed, could be moved to the intersection of the purple passes with 

sufficient time to “catch” Opportunities 3 and 4. This pattern could be propagated reasonably for 

days and various landing opportunities. For situations where water-landings were limited to day 

or night only, the corresponding number of opportunities would be reduced by approximately 50 

percent. 

Water As Land-Landing Backup 

Due to the flexibility of water-landings and the ability to move the starting point, water-landings 

could readily make the transition from primary to backup landing. For example, in Figure 7.2-

22, the yellow arrow shows an opportunity to land at EAFB. If this opportunity was missed, the 

subsequent water passes could act as a feasible backup until another land opportunity occurred. 
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Figure 7.2-22. Water as back up to Edwards on 1/5 

To complete this analysis of investigating water as a backup to land passes, SLAAM was 

adjusted to output data in both hours and days to provide the information in a more readable 

format. Figure 7.2-23 shows landing opportunities in relation to each other over time. This is 

the same information in Table 7.2-17 

Figure 7.2-23. Landing accessibility for the landing site network, all landing options available, 180 

days (14 days shown for clarity), in hours and days 
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Figure 7.2-24 shows how the opportunities to land change with water as a backup to EAFB pass 

on January 1 and 11. Note how the water-landing site fills in the gaps left by EAFB passes. 

Figure 7.2-24. Water as back up to EAFB 

CONUS for 28.5 Degree Inclination Orbit 

The purpose of looking at the 28.5 degree inclination orbit was to determine if any CONUS 

landing opportunities existed. As illustrated in Figure 7.2-25, there are no CONUS land-landing 

sites available for 28.5 degree inclination orbits. Figure 7.2-25 also shows an example landing 

location in the Pacific Ocean, with a circle reporting the radius of where a ship could move in 24 

hours to support a landing opportunity. This study did verify that if an appropriately sized deltavelocity was applied to the CM for an inclination change, a CONUS based landing site could be 

reached. 
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Figure 7.2-25. Lunar inclination CONUS landing evaluation 

7.2.4 Task 4: CM Roll Control in Preparation for Landing Data Analysis 

While roll control was included in the original Apollo specification, it was ultimately not 

included as a capability for use during landing. However, the CM was required to land in any 

possible roll orientation. Roll control was referenced in the original Apollo Program 

requirements. Although the potential advantages of roll control were recognized, no attempts 

were made to implement roll control for landing. Apollo had requirements for both nominal and 

emergency crew acceleration limits and onset rates (Figure 7.2-26). When it was determined that 

these levels could not be met during a land-landing, the Apollo Program accepted the risk based 

on the low probability of a pad abort. 

The NESC team found that there should be sufficient capability in the RCS to orient the CM per 

the CEV Project requirements with low riser angles. The CEV Project had performed parachute 

riser / attachment torque tests to verify the capability. The CEV Project found torque resistance 

of baseline riser /attachment design needed to be reduced, and subsequently tested designs with 

sufficiently low torque resistance (with margin). The results of these tests are provided in Figure 

7.2-27. 

No land-landing sites 

available for lunar 

inclination orbit 
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Figure 7.2-26. Crew Acceleration Limits for the Apollo Command Module 
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Figure 7.2-27. CM Roll Torque Capability Under Main Parachutes 

The CEV Project tests show that the CM RCS has sufficient torque to meet the +30 degree roll 

requirement with current parachute harness / riser designs 

7.2.5 Task 5: Investigate Parachute Release Times During Land-Landing Data Analysis 

The effectiveness of releasing the parachute from the CM at different release times post 

touchdown is examined for a vertical and horizontal landing velocities of 26 and 37 fps, 

respectively. A 37 fps horizontal velocity was used since this velocity represents one of the 

more extreme conditions which is most likely to cause higher rollovers. The 15 degree hang 

angle is fairly optimal for CM stability in the presence of a horizontal landing velocity, while a 

flat 0 degree hang angle may be advantageous for a purely vertical landing. 

The transient rigging force is shown in Figure 7.2-28. The force starts at 13,046 pounds which is 

the CM weight and the initial tension in the rigging just before touchdown. Upon touchdown the 

rigging force quickly decreases as the ground stops the CM from further motion and the 

parachute continues to descend. At 0.17 seconds the rigging goes completely slack and remains 

slack until near 0.4 seconds where the tension begins to increase as the horizontal wind load 

drags the parachute beyond where the CM is contacting the ground. 
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Figure 7.2-28. Rigging Force without Chute Release at a 15 degree Hang Angle 

Figure 7.2-29 depicts the effect on CM roll of releasing the parachute at different times beyond 

initial ground contact. When the parachute is released at touchdown (chute quick release), the 

CM roll is a minimum compared to all of the other release times examined. When the release 

time is at 0.17 (when the rigging first becomes slack), 0.08, and 1.00 seconds, the CM roll is 

similar and is greater than when the parachute is released at 0 seconds. When the parachute is 

kept attached for the duration, the horizontal wind force on the parachute dominates and the CM 

tumbles. 
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Figure 7.2-29. Effect of Parachute Release Time on CM Roll - 15 degree Hang Angle, CM does not 

rollover when chute released before 1.0 seconds. 

Figure 7.2-30 shows the resultant acceleration at the location of the crew members for the 

different values of parachute release times. The maximum acceleration occurs below 0.20 

seconds. Beyond the peak acceleration, the acceleration profiles are similar for all release times 

except for the case where the parachute is not released. For this case the second peak in the 

acceleration profile near 0.50 seconds is larger than for the rest of the cases. This peak occurs 

after the CM “bounces” and incurs a second landing. The peak is larger for the parachute 

without release since the parachute pulls the CM higher off the ground than for the other cases 

where the parachute is released. 
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Figure 7.2-30. Effect of Parachute Release Time on Pallet Acceleration 15 degree Hang Angle 

Figure 7.2-31 shows a CM cross section for the 0 degree hang angle. For this angle the 

parachute attachment point is moved to directly above the center-of-gravity. 
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Figure 7.2-31. 0 Degrees Hang Angle (Rigging Attachment Located to Maintain Hang Angle) 

Figure 7.2-32 shows the transient rigging force for the case where the parachute is not released. 

Similarly to the 15 degree hang angle results, the force starts at 13,046 pounds and upon 

touchdown the rigging force quickly decreases then the tension increases as the horizontal wind 

load drags the parachute. For the 0 degree hang angle the parachute first becomes slack at 0.16 

seconds. 
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Figure 7.2-32. Rigging Force without Chute Release 0 degree Hang Angle 

Figure 7.2-33 depicts the effect on CM roll of releasing the parachute at different times beyond 

the time of initial ground contact. In general, the CM roll is similar regardless of the release 

time. When the parachute remains attached for the duration, the CM roll is slightly less than for 

the other cases for part of the response but converges to the same roll at the end of the 

simulation. All the cases converge to the same roll since the CM has rolled over and is resting 

on its top. Figure7.2-34 shows the resultant accelerations for the different release times and 

similar to the roll profiles the acceleration profiles are similar for each case with a peak 

acceleration near 0.20 seconds then a second spike in the acceleration when the CM rolls over 

and the top strikes the ground. 
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Figure 7.2-33. Effect of Parachute Release Time on CM Roll 0 degree Hang angle, Vehicle rolls over 

for all cases 

Figure 7.2-34. Effect of Parachute Release Time on Pallet Acceleration 0 degree Hang Angle 

Figure 7.2-35 shows the effect of horizontal landing velocity and parachute release time on CM 

rollover. Figure 7.2-36 are for a CM with a 0 and 15 degree hang angle, respectively. In the 

presence of a horizontal landing velocity the 15 degree hang angle is more stable. While the 
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vehicle with a 15 degree hang angle can accommodate a horizontal landing velocity of close to 

65 fps, the CM with a 0 degree hang angle rolls over when the horizontal velocity is 30 fps. 

Note that these rollover velocities are relative and additional study is needed to determine values 

that are more absolute. As depicted in the figures, parachute release time has minimal effect on 

CM rollover. For the 15 degree hang angle the release time has no effect on rollover so long as 

the parachute is released before 0.18 seconds. For the 0 degree hang angle the allowable 

horizontal velocity before rollover occurs is slightly higher when the parachute is released at 

touchdown. However, similarly to the 15 degree results release time does not significantly affect 

rollover. Release times beyond 0.18 seconds because the rigging goes back into tension the 

tension is a result of the horizontal wind always has a detrimental effect on rollover. Thus, the 

parachute should always be released before the rigging regains tension. 

A CM employing airbags for landing load attenuation has conflicting requirements with a 

vehicle configuration striving to minimize rollover potential. Airbags are most effective for a 

flat (0 degree hang angle) landing and are less effective when there is a horizontal landing 

velocity or the vehicle lands with a hang angle. However, there can be situation when there is a 

horizontal landing velocity and if the CM hang angle is 0 degrees to optimize the airbags 

effectiveness, the vehicle will be more prone to rollover than if the hang angle was 15 degrees. 

Figure 7.2-35. Effect of Horizontal Velocity on Rollover 0 degree Hang Angle, Vertical velocity = 

26 fps, No Roll 
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Figure 7.2-36. Effect of Horizontal Velocity on Rollover 15 degree Hang Angle, Vertical Velocity = 

26 fps, No Roll 

7.2.6 Task 6: Crew Protection System Enhancements Data Analysis 

The last phase of all manned space missions is the landing and recovery phase. The CM will 

descend to earth with a parachute and should be capable of a primary land-landing and waterlandings. Consequently, landing systems must be designed for multi-terrain impacts. Earth 

terrain is highly variable, with soil at proposed landing sites at EAFB typically very hard and soil 

at Carson Sink, NV often very soft. 

To keep the size of parachute systems reasonable and to optimize the weight of the landing 

system, the terminal velocity of most parachute recovery systems is approximately 25 to 40 fps 

(17-27 miles/hour). In addition to the terminal sink (vertical) velocity, the parachute and capsule 

will move horizontally with the air mass depending on the wind speed. The velocities of the CM 

at touchdown are typical of crash impact velocities of small aircraft and helicopters. Thus, these 

impact velocities will produce injuries without some type of mitigation for impacts onto land. 

Water-landings do not require additional mitigation for nominal landings since the water acts as 

a natural impact attenuator. The Apollo capsule was designed to impact the water with a hang 

angle of 27 degrees. Since the Apollo capsule entered the water pitched at an angle (thus giving 

it a wedge shape), the acceleration experienced was nominally 10 g or less for a 30 fps impact 

[ref. 1]. The Apollo spacecraft also had crushable ribs and a crew seat pallet with stroking 

energy absorbers to alleviate off-nominal impacts. However, either a land-landing in an abort 

condition or a flat 0-degree attitude water impact with one chute out was expected to result in 

occupant injuries. Since the probability of a pad abort with a land impact was considered to be 

low, the Apollo Program accepted those risks. 
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Under ideal nominal conditions, either a land-landing with retrorockets or airbags, or a waterlanding will result in tolerable accelerations to the crew if the seats and restraint systems are 

adequately designed. For off-nominal landing conditions, a stroking seat pallet similar to the 

Apollo design is planned for CEV Project CM should reduce the impact accelerations to limits 

that are tolerable and prevent crew injury. However, some mitigation for more severe offnominal impacts can be designed with little weight penalty if a systems approach is taken early 

in the design phase. 

With properly designed seats (i.e., with side supports and restraints), relatively high accelerations 

can be tolerated with no or minimal injury, especially for impacts with the crew lying on their 

backs in a seat oriented as shown in Figure 7.2-37. The local axis system has the +X-axis 

upward, the +Z-axis pointed toward the head, and the +Y-axis to the right. Note that the axis 

system typically used for a seat is a left-handed system. For accelerations along the +X-axis, 

which would occur for a flat impact, the acceleration in the spine-to-chest direction is commonly 

called “eyeballs in.” An acceleration along the –X-axis (chest-to-spine) is called “eyeballs out.” 

An acceleration from pelvis-to-head along the +Z-seat axis is called “eyeballs down,” and an 

acceleration pointing in the –Z-axis is “eyeballs up.” The seat coordinate system is a local 

system, and is a non-inertial axis system. Hence, any angular rotations with respect to an inertial 

system must be considered. 

Figure 7.2-37. Six Crew Configuration with local X, Y, and Z-Axis 

Human tolerance to impact has been studied continuously since the advent of the automobile and 

airplane. During the Apollo Program, over 288 impact tests were conducted for NASA on 79 

male volunteers using the Holloman Daisy Accelerator [ref. 2]. The acceleration onset rates 

ranged from 300 to 2500 g per second. In general, lower onset rates are better tolerated. These 

tests were designed to help understand the effects of water impact on the crew. 
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The US Air Force and Navy have conducted impact test programs in an effort to make ejection 

seats as safe as possible. Sled impact testing of volunteers has been conducted for low, noninjurious accelerations. Cadavers, animal surrogates, and ATD’s, such as the Hybrid II and III 

have been used in sled tests to help determine the threshold of moderate and severe injuries. A 

comprehensive compilation of many of these tests that provides summary charts with regions of 

uninjured, moderate, and severe injuries versus maximum continuous sled acceleration and 

duration was produced by Eiband at NASA Lewis [ref.3]. The testing performed in this work is 

relevant to orientations where the loading is directly up through the spine from the buttocks to 

head and is less relevant to CM landings where the nominal loading is in the eyeballs out 

direction. 

7.2.6.1 Mathematical Lumped Mass Injury Models 

In order to reduce injuries produced during ejection from military aircraft, relatively simple but 

effective mathematical representations of a seated human were developed. Ejection seats were 

first designed based on acceleration limits in the +Z-direction and acceleration onset rates. Early 

on, it was determined that accelerations were better tolerated by aviators if the ejection 

acceleration onset rate was as low as practical. As ejection seats evolved and became more 

complex (i.e., those with thrust vectoring rockets and encapsulated seats in the B-70 and F-111), 

peak acceleration and onset rate proved ineffective for evaluating their performance. New 

design tools became necessary. The impact tolerance charts produced by Eiband were useful, 

however their application to an arbitrary acceleration pulse that includes accelerations along all 

three axes (X, Y, and Z) is subjective. An example of an Eiband chart for accelerations along the 

spine (+Z, eyeballs-down) is shown in Figure 7.2-38. A modified chart [ref. 4] with ejection seat 

design limits added is illustrated in Figure 7.2-39. 
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Figure 7.2-38. +Z Acceleration Tolerance with Ejection Seat Data Omitted 

Figure 7.2-39. +Z-Human Tolerance Limits showing Ejection Seat Design Limits 
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There are three areas to consider on each chart. The Area of Voluntary Human Exposures is the 

bottom-shaded area. In this region, all exposures were uninjured. For example, a maximum 

acceleration of approximately 14 g could be tolerated (uninjured) for up to 0.04 seconds (40 

milliseconds). The other shaded area at the top of the chart is the Area of Severe Injury. For 

example, acceleration above 100 g for longer than 3 milliseconds produces severe injury. The 

area between the two darkened areas is the Area of Moderate Injury. 

7.2.6.2 The Dynamic Response Index (DRI) Model 

The most effective human model for early ejection seat design was the DRI described by Stech 

and Payne in 1969 [ref. 5]. This work was performed at what is now called the Wright-Patterson 

Armstrong Aerospace Medical Research Laboratory (AAMRL) and was based on pioneering 

work in human biodynamics by von Gierke [ref. 6]. The DRI model was verified by 

comparison with actual ejection seat data (see Figure 7.2-40) by Brinkley [ref. 7]. For many 

years the US Air Force used the lumped mass-spring DRI model and only considered the +Zaxis. The DRI model was developed to estimate the probability of spinal fractures for 

accelerations in the pelvis-to-head or +Z direction. In Figure 7.2-40, the abscissa should be 

marked DR+Z as only the +Z component is considered in the DRI. 

The DRI method was generalized by Brinkley to include the other local orthogonal axes (-Z, +X, 

-X, +Y, -Y) [ref. 8]. A brief description of the Dynamic Response Method (referred to as the 

Brinkley Dynamic Response Model) follows. 

Figure 7.2-40. DRI+Z Related to Probability of Spinal Injury for Ejection Seats 

(Note that cadaver data was over conservative.) 
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The equations of motion of the human body modeling as a lumped mass are three second-order 

differential equations. Each equation is familiar as simply a forced spring-mass harmonic 

oscillator with damping. The forcing function “A” in Eq. 1 is the measured acceleration along 

each of the three seat axes. As a simplifying condition, the three equations are considered to be 

uncoupled. The second-order differential equation for the Brinkley Dynamic Response model is 

given by Eq. 1. This equation was given in the Constellation Program (CxP) specification CxP 

70024 Human System Integration Requirements (HSIR) December 2006: 

Axxx nn 

2 2   EQ. (1)

where:

is the relative acceleration of the dynamic system mass with respect to the accelerometer 

location in the given axis system (X, Y, or Z).

is the relative velocity of the mass with respect to the accelerometer location

in the given axis system.

is the deflection of the mass with respect to the critical point in the given axis system. A 

positive value represents compression.

 is the damping coefficient ratio.

n is the undamped natural frequency of the dynamic system.

A is the component of the measured acceleration along the specified axis. 

 Since the seat axis is not an inertial frame, rotational acceleration must be considered.

The dynamic response for each axis is given by:

 

 EQ. 2

Where Dynamic Response (DR) (t) is the response of the dynamic system and g is the 

acceleration of gravity. Note that DR is dimensionless. The value “x” when used along the +Z 

axis is the compression of the spine. Note that DR(t) is essentially the normalized response of 

the occupant to the input acceleration in g’s. The maximum value of DR(t) for any axis should 

be less than the limiting value. Thus, DR(t) can be plotted versus the input acceleration time 

history. The maximum of the response DR(t) can be greater or less than the maximum of the 

input acceleration A(t) depending on whether the simple harmonic oscillator amplifies or 

attenuates the driving acceleration. The following values for n and  were used:

x

x

x

gxDR n / 2
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nx = 62.8

ny = 58.0

nz = 52.9

x = 0.2

y = 0.09

z = 0.224

Limiting values of DR levels have been set by NASA for the CEV Project and are listed in Table 

7.2-18. These values are also provided in CxP 70024 HSIR. The Very low (nominal) limits 

were derived by Brinkley [ref.9] from studies initiated in the 1980’s for NASA to accommodate 

a “deconditioned” crew member who had been in space for a considerable time. The Low (off 

nominal) DR limit is the same as provided in [ref. 8]. The limits also consider DR levels for 

moderate and high risk [ref. 8]. This data is shown in Table 7.2-19. Side restraints raise the 

dynamic response limits for the Y-direction as shown in Table 7.2-19. Note that side restraints 

are not mentioned in the CxP HSIR specification.

Table 7.2-18. Brinkley Dynamic Response Limits (NASA HSIR Specification for the CEV Project) 

DR level 

X 

(eyeballs out, in) 

Y 

(eyeballs right, left) 

Z 

(eyeballs up, down) 

DRx < 0 DRx > 0 DRy < 0 DRy > 0 DRz < 0 DRz > 0 

Very low 

(nominal) -22.4 31 -11.8 11.8 -11 13.1 

Low 

(off-nominal) -28 35 -14 14 -13.4 15.2 

Table 7.2-19. Dynamic Response Limits from Brinkley and Moser [Ref. 8] 

DR level 

X 

(eyeballs out, in) 

Y 

(eyeballs right, left) 

Z 

(eyeballs up, down) 

DRx < 0 DRx > 0 DRy < 0 DRy > 0 DRz < 0 DRz > 0 

Low (same as 

NASA spec) -28 35 -14 14 -13.4 15.2 

Moderate -35 40 -17 17 -16.5 18 

High Risk -46 46 -22 22 -20.4 22.4 
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Also note that if side restraints are used, the DR levels for Y increase to -15 and +15 (Low), -20 

and +20 (Moderate), and -30 and +30 (High Risk). Caution – These values are not listed in CxP 

70024.

The limiting value of the DRlim in each axis from Table 7.2-18 is used in Eq. 3 to assess Very 

Low or Low risk. The value of beta, β, in Eq. 3 must be less than one to pass the Brinkley 

Dynamic Response Method criteria for very low or low risk. Note that beta could lag the input 

acceleration in time since the resonant frequencies of the lumped mass system are on the order of 

10 Hz. Consequently, DR should be calculated until the maximum for each direction has been 

identified.

      2

lim

2

lim

2

lim 















 















 













 

z

z

y

y

x

x

DR

tDR

DR

tDR

DR

tDR  < 1 EQ. 3 

In the single +Z axis DRI model, the human body was modeled as a single lumped mass with a 

spring and damper representing the spine for Z-direction impacts. Thus DRz (t) is the nondimensionalized response of the lumped mass of the torso to a forced acceleration input along 

the Z-axis. In general for multi-axis accelerations, the linear second-order differential damped 

oscillator given by Eq. 1 is solved independently for each direction. The input acceleration for 

the model should ideally be the acceleration of the seat. For an ejection seat, the primary 

acceleration would be eyeballs down, or in the +Z direction. Note that unlike Apollo where the 

crew was laying on their back in the seat, in a fighter jet the seat would be oriented with the Zaxis pointing up. From operational data, the DRI for Z-direction from Eq. 2 was correlated to 

injury using the data shown in Figure 7.2-40. From the chart, most of the maximum DR values 

from operational data ranged from a low of 16 to a high of 20. Any DR greater than 22 or 23 

was correlated to a spinal injury rate above 50 percent. Note that the DR+z limit (DR+Z

lim ) in 

Equation 3, shown in Table 7.2-18 for off nominal in the +Z direction is 15.2, which corresponds 

to an injury probability of less than 1 percent. The DR+z nominal from Table 7.2-20 is 13.1. 

Although the Brinkley Dynamic Response Method model has had success in predicting injury, it 

is a simplified model of a complex dynamic system (i.e., the human body). Also, the method of 

combining the results of accelerations in the X, Y, and Z-directions is not verified as most test 

data were gathered for a single acceleration direction. Insight is obtained by examining the DR 

for each direction before combining the results to obtain Beta in Equation 3. Finally, note that 

the Brinkley Dynamic Response Method model is only used as a predictive tool for injury. It is 

not correlated with fatalities. The threshold for fatalities is difficult to predict. 
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Table 7.2-20. Vertical Velocity and Brinkley Dynamic Response Method Injury Landing Vertical 

Velocity (fps)/Brinkley DRI 

In summary, the US Air Force Research Laboratories developed human tolerance models that 

were grounded in previously developed models to assist in the design and evaluation of ejection 

seats. Both human volunteers and ATDs were used along with operational ejection seat data to 

validate these models. The Brinkley Dynamic Response Method is a generalization from +Z 

impacts to impacts in all orthogonal directions. The model has proved to be extremely useful 

and is included in the military specifications for aircraft crew escape systems. 

In contrast to the US Air Force, the US Army developed design guides requiring features such as 

crashworthy cockpits and seats to protect the crew in the event of a “survivable crash” [ref. 10]. 

The automobile industry was faced with Federal regulations to design and build safer cars. 

ATDs with more human impact fidelity such as the Hybrid III were developed as instrumented 

passengers for controlled crash tests. The National Highway Traffic Safety Administration 

(NHTSA) regulates automobile safety and also sponsors research to make automobiles safer. 

The Brinkley Dynamic Response Method is limited in that it uses a simplified mechanical model 

of the human body. In this case the entire body is modeled as a point mass located at the center 

of mass of the human body and attached with a damped spring oscillator with an approximately 

10 Hz resonance frequency. A schematic of the physical layout of the Brinkley model, showing 

the central mass located in the mid-thorax, is shown in Figure 7.2-41. 
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Figure 7.2-41. Central Mass Located in the Mid-Thorax (Brinkley Dynamic Response Method 

Model) 

The nature of the Brinkley Dynamic Response Method model and its representation of the 

human body as a single mass preclude the model from distinguishing relative movement of 

individual body parts. Instead, the human body parts are all modeled as moving together as a 

point mass. While this model works reasonably well for vertical Z-axis loads, such as those of 

an ejection seat, where the major injury location is in the lower spine and the human body moves 

as a unit, it does not properly physically model some other impacts, such as side impacts, where 

sections of the body like the head relative to the body center of mass. In addition to the Brinkley 

Dynamic Response Method model not having the capability to discern different types of injuries, 

the Brinkley model cannot provide insight into the effect of any advanced restraint system that 

provides additional restrictions or protection to the head or other parts of the body. 

The current state of the art does not allow the direct modeling and simulation of a human body 

subjected to injurious high acceleration loading or impact conditions. To address this limitation 

the automotive industry has attempted to generate a relationship between the physical behavior 

of the various ATDs and the human body. This has been attempted via a combination of 

cadaver, sub-injury level testing on human volunteers, animal experimentation, and data 

gathering from actual accidents. The degree to which the ATD simulates the actual behavior of 

the human body is referred to as its “bio-fidelity”. This is a difficult mechanical engineering 

problem and the ATD have gone through an evolution and characterization to improve and 

understand their bio-fidelity. 

The general mechanical difficulty in developing a single three-direction ATD that has adequate 

levels of bio-fidelity has proven to be challenging and has motivated the automotive industry to 

develop a series of ATDs specific to various types of automobile impact conditions. The three 

main types of ATD types of interest are: frontal collision full body motion, rear impact, and side 

impact. For frontal impacts the industry utilizes the Hybrid III ATD as the standard. This ATD 
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has been calibrated with years of test data and is considered a good indicator for whole body 

motion for frontal impact. The HYBRID III, however, is considered to be inadequate for rear 

impacts so the industry has developed a number of rear impact ATDs. The rear impact ATDs 

have more accurate head and neck mechanisms and are better able to predict neck and whiplash 

types of injuries. The Rear Impact Dummy (RID) series of ATDs include RID1, bio-RID, RID2 

and Bio-RIDII. The most important aspects of these ATDs are the additional attention to the 

mobility of the spine particularly in the neck region and also in the area near the hips. The third 

type of ATD is the Side Impact Dummy (SID). The SID’s are similar to the RID’s but with 

additional fidelity for spinal motion in the sideways, Y-axis, direction. To reduce reliance on 

crash testing the industry has developed FEMs of the various ATDs. 

In the present effort, the Hybrid III FEM was selected as the baseline. This is the most 

commonly used ATDs and was thought to be a good starting point for comparing the results 

from the ATDs to those from the Brinkley Dynamic Response Method model. It was also 

thought that the Hybrid III was a reasonable ATD for assessing alternate crew protection 

concepts. It is recognized that for situations such as side and rear impact, models such as the 

SID or RID are more suitable. 

In order to directly relate the results of the FEA simulations (which provide forces and moments 

in strategic locations throughout the model) and actual injury, the nature of specific relevant 

injury, and the forces or moments that produce it must be understood. Both the US Air Force 

and the Navy have active programs in the area of biomechanics and injury potential of ejection 

seats. Much of this work is summarized in a 1993 study entitled “Aircrew Ejection Injury 

Analysis and Trauma Assessment Criteria” published by the University of Michigan 

Transportation Research Institute [ref. 11]. This work was commissioned by the US Navy as 

part of an effort to better understand the use of ATDs in the evaluation of escape systems. This 

paper is a survey of the state of the art in understanding the effect of ejection seat g forces on the 

human body. In particular the data indicates that injuries that are experienced at these 

acceleration levels (+ or -Z axis only) are almost always related to injuries of one type or another 

to the spinal column. Serious injuries to soft tissue that are not directly related to impact are 

effectively nonexistent. This fact is important to the CM effort since it indicates that the focus, 

at least in the Z-axis direction, can be primarily on predicting and preventing spinal column 

injuries. 

7.2.6.3 Finite Element Anthropomorphic Test Device (FE ATD) Simulations 

The CM (version 604), Figure 7.2-42 was used to generate vehicle responses so that a 

comparison could be made between injury predictions from the Brinkley Dynamic Response 

Method model and predictions from the FE ATD. To generate the response, the CM was 

oriented in three primary human body directions; eyes in/out (Z-axis), sideways (Y-axis), and the 

spinal (Z-direction). The CM was given an initial velocity just before impact with the ground. 

These orientation exaggerate realistic landing scenarios, however, they provide extreme cases for 

comparing the Brinkley Dynamic Response Method model to the ATD results. 
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Figure 7.2-42. CM 604 Seat Attenuation Model 

LS-DYNA® was used to perform the analysis of the CM impacting ground and resulting 

acceleration profiles were computed and extracted at the location of the crew members’ center of 

mass. The extracted acceleration profiles were then input into the Brinkley Dynamic Response 

Method model and the level of injury was computed using the Brinkley injury criteria. The 

initial impact velocities were adjusted until they corresponding to high, medium and low injury 

risk were identified (refer back to Table 7.2-20). 

Once the impact velocities corresponding to the three levels of Brinkley Dynamic Response 

Method injury were determined and the acceleration profiles were generated (Figures 7.2-43 

through 7.2-54) a separate LS-DYNA® model was created that included only the FE ATD 

constrained in a seat with a five point harness. The acceleration profiles were then applied to the 

seat in this model, simulations were run, and a comparison was made between the Brinkley 

Dynamic Response Method injury criteria and the response extracted from the FE ATD. 

a) Seat Attenuation Model

(Global Coordinate System)

b) Seat Attenuation Model 

(Brinkley Local Coordinate System)

(Pressure Vessel not Shown)

Y

Z

X

Vh

Vv

Z

X

Y

Global

a) Seat Attenuation Model

(Global Coordinate System)

b) Seat Attenuation Model 

(Brinkley Local Coordinate System)

(Pressure Vessel not Shown)

Y

Z

X

Vh

Vv

Z

X

Y

Global
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Figure 7.2-43. Configuration for X Acceleration (eyes in/out DRI) 

Figure 7.2-44. Acceleration Profile for Low Eyes In/Out DRI (Vn=33.3 fps) 
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Figure 7.2-45. Acceleration Profile for Moderate Eyes In/Out DRI (Vn=41.67 fps) 

Figure 7.2-46. Acceleration Profile for High Eyes In/Out DRI (Vn=58.33 fps) 
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Figure 7.2-47. Configuration for Y Acceleration (sideways DRI) 

Figure 7.2-48. Acceleration Profile for Low Sideways DRI (Vn=16.66 fps) 
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Figure 7.2-49. Acceleration Profile for Moderate Sideways DRI (Vn= 25 fps) 

Figure 7.2-50. Acceleration Profile for High Sideways DRI (Vn= 33.3 fps) 
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Figure 7.2-51. Configuration for Z Acceleration (spine DRI) 

Figure 7.2-52. Acceleration Profile for Low Spine DRI (Vn= 25 fps) 
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Figure 7.2-53. Acceleration Profile for Moderate Spine DRI (Vn= 29.2 fps) 

Figure 7.2-54. Acceleration Profile for High Spine DRI (Vn= 33.3 fps) 

The FE ATD provides a greater amount of information concerning human body response 

including actual motions of the body such as limb flailing, head motion, and loads within the 

body (e.g., acceleration levels at specific locations in the body and forces on individual body 

parts). Figures 7.2-55 through 7.2-62 show details of the Hybrid III ATD and some of the 

locations in the model where data is extracted. Figures 7.2-55 through 7.2.6-57 show the 
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location of where accelerations were extracted in the head, pelvis, and thorax. Accelerations at 

these locations are important since considerable research has been performed and injury criteria 

have been developed that provide correlations between acceleration levels and likelihood of 

injury. Figures 7.2-58 and 7.2-59 depict the lower lumbar region of the FE ATD model and 

where these forces and moments are extracted, while Figures 7.2-60 through 7.2-62 depict the 

neck region and locations where neck forces and moments are extracted. Similarly to the ATD 

accelerations, industry standards have been established that minimize injuries by limiting 

allowable predicted forces and moments in the ATD. 

Figure 7.2-55. Location of Head Acceleration 

Figure 7.2-56. Location of Pelvis Acceleration 
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Figure 7.2-57. Location of Thorax Acceleration 

Figure 7.2-58. Location of Lower Lumbar Spine Force 
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Figure 7.2-59. Location of Lower Lumbar Spine Moment 

Figure 7.2-60. Location of Neck Joints 
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Figure 7.2-61. Location of Upper Neck Moment 

Figure 7.2-62. Location of Lower Neck Moment 

As previously mentioned in this report, it is important to note that many of the industry accepted 

standards have been developed for the automotive industry or ejection seats, and may not be 

directly applicable to the type of conditions and injuries expected for the CM crew members. It 

is also important to note that the FE ATD is an analytical model of the physical simulator, which 

is made up of non-human parts made from materials such as steel, aluminum, rubber, and plastic. 

The FE ATD is designed to predict the response of the physical simulator, which in turn is 

designed to mimic how a human might respond during an actual impact. 
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7.2.6.4 Z-Axis (Spinal) Acceleration Simulations and Results 

Loading in the Z-axis (spinal) direction was examined since this is the loading direction that the 

Brinkley Dynamic Response Method model was originally designed to address and where the 

model is most accurate. Since the results of the Brinkley model are anchored in actual data for 

loading in the spinal direction, it was useful to compare the Brinkley predictions to the results 

generated from the FE ATD and to industry accepted injury criteria. Figure 7.2-63 shows the FE 

ATD in a five-point harness. This harness was used since the five-point harness is the harness 

used for the development of the Brinkley model. The harness was composed of 2 inch nylon 

webbing. The harness was attached to a relatively rigid steel simplified seat and base 

accelerations were applied to the seat in the Z-axis direction (from the pelvis upwards towards 

the head). The commercially available Hybrid III FE ATD model and LS-DYNA® were used to 

generate the transient response. 

Figure 7.2-63. Hybrid III ATD Constrained with Five Point Harness 

Figure 7.2-64 and Table 7.2-21 provide representative output from the LS-DYNA® simulations. 

Figure 7.2.6-28 shows two intermediate time steps from the simulation. The first frame shows 

the maximum head flexion, which occurs just after ground impact, and the second frame shows 

maximum head extension which occurs during the time period where the CM rebounds off of the 

ground. Table 7.2.6-4 provides a comparison among the three Brinkley Dynamic Response 

Method injury levels and a collection of injury criteria [ref. 12]. Allowable limits for the injury 

criteria are included in the table and conditions where the limits are exceeded are identified by 

red. 
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Figure 7.2-64. Hybrid III ATD Response at Intermediate Time Steps 

Table 7.2-21. Hybrid III Response and Brinkley Dynamic Response Method Injury Levels Z-Axis 

(Spinal) Excitation 

The Brinkley Dynamic Response Method injury criteria are consistent for low levels of injury in 

that none of the injury criteria allowable are exceeded. For medium Brinkley injury the 

allowable thorax acceleration limit is exceeded, while the remainder of the injury criteria is 

within allowable limits. For high Brinkley injury the thorax acceleration is close to double the 

allowable limit and the neck moment extension allowable limit is exceeded. It is not surprising 

JT# 26-37 -resultant 6672 N 2804 N 4037 N 4545 N

JSR# 17 -

JSR# 16 -

JT# 2-5 - resultant

JSR# 15 -

JT# 38 -resultant

N# 7501 - resultant

N# 1787 -resultant

N# 1787 CSI

N# 1 HIC

Location

Neck Moment Flexion 310 NM 66.42 NM 78.6 NM 79.42 NM

(-)

Pelvis G’s 32.7 G’s 49.42 41.9 G’s

6672 N 4080 N 4500 N 5500 N

Lumbar Force

Brinkley High 

Injury

Brinkley Medium 

Injury

Brinkley Low 

Injury

Criteria Allowable 1

Head Injury Criteria 700 3.4 4.57 59.54

(HIC)

Chest Severity Index 700 31.6 43 64.49

(CSI)

Neck Moment 135 NM 108.6 NM 130 NM 184 NM

Extension (+)

Neck Force 6806 N 887 N 698 N 1277 N

Lumbar Moment 131.4 NM 143 NM 124.3 NM

Thorax G’s 60 G’s 46.17 G’s 75.27 G’s 127.7 G’s

JT# 26-37 -resultant 6672 N 2804 N 4037 N 4545 N

JSR# 17 -

JSR# 16 -

JT# 2-5 - resultant

JSR# 15 -

JT# 38 -resultant

N# 7501 - resultant

N# 1787 -resultant

N# 1787 CSI

N# 1 HIC

Location

Neck Moment Flexion 310 NM 66.42 NM 78.6 NM 79.42 NM

(-)

Pelvis G’s 32.7 G’s 49.42 41.9 G’s

6672 N 4080 N 4500 N 5500 N

Lumbar Force

Brinkley High 

Injury

Brinkley Medium 

Injury

Brinkley Low 

Injury

Criteria Allowable 1

Head Injury Criteria 700 3.4 4.57 59.54

(HIC)

Chest Severity Index 700 31.6 43 64.49

(CSI)

Neck Moment 135 NM 108.6 NM 130 NM 184 NM

Extension (+)

Neck Force 6806 N 887 N 698 N 1277 N

Lumbar Moment 131.4 NM 143 NM 124.3 NM

Thorax G’s 60 G’s 46.17 G’s 75.27 G’s 127.7 G’s
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that the allowable neck moment is exceeded considering the head is not constrained and is 

predicted to undergo considerable motion during impact. It is interesting that the lumbar force is 

within allowable limits for all three Brinkley Dynamic Response Method injury levels 

considering that for this direction of loading spinal cord injuries, particularly in the lumber 

region, are prevalent. 

Many of the automotive injury criteria were developed for frontal impacts where the human body 

rapidly decelerates and body parts such as the head impacts some part of the automobile interior. 

Similarly, aircraft ejection seat criteria are designed for an upright seated pilot being ejected 

vertically through the aircraft canopy where the primary loading is up through the spine. For 

CM nominal landings, the crew members will be seated on their backs and the primary loading 

will be in a direction from the crew member's back towards the chest. Further studies are 

necessary before existing industry accepted injury criteria be used to assess CEV Project crew 

member injury. 

7.2.6.5 Y-Axis (Sideways) Acceleration Simulations and Results

Results of FEA simulations for the standard five-point harness were compared with simulations 

that contained a variety of lateral supports inspired by a modern race car seat. Modern race car 

seats provide excellent driver protection during impacts that might normally be fatal with a more 

traditional seat design. These seats provide a higher level of protection through the use of lateral 

supports, and head and neck restraint systems, and improved harnesses. A typical modern race 

seat is shown in Figure 7.2-65 alongside the seats from the Apollo Program. The differences in 

lateral support between the two seat designs are evident. Much of the improvements that have 

been realized in race car seats are the results of trial and error combined with results obtained 

during actual operations and accidents. CM seat designs have limited opportunities for 

development tests and therefore must take advantage of improvements in restraint or seat designs 

and subscale test using FE ATDs. 
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Figure 7.2-65. Modern Race Car Seat with Lateral Supports (left) and Apollo CM seats (right) 

The use of lateral supports of the type found in race car seats may not be directly applicable to 

the CEV Project since these seats may be impractical for crew ingress and egress, and for 

stowage during flight. In order for lateral restraints (or any other restraint or crew protection 

system) to be used in space flight applications the system must provide for a practical means for 

ingress, and rapid egress, and stowage when the system is not in use. In addition the restraint 

system must be flexible in its ability to accommodate crew members of different sizes and 

weights. 

In a standard seat such as that used in Apollo CM there is no lateral support other than that 

provided by the harness. There was a slight depression that the crew member sits, but the dish 

shape provided no real restraint during high g Y-axis accelerations. In the case of CEV Project 

CM there is the additional complication of having two distinct decks of seating, with two crew 

members placed in a second row. If permanent lateral restraints are added to the seats, they will 

most likely make it difficult or impossible to get in and out of the seats. 

One system that shows particular promise is a series of side airbags mounted on panels that are 

linked together and can be quickly extended and retracted by the crew for ingress and egress, see 

Figure 7.2-66. The illustration on the left shows the airbag pallets in the extended position and 

the illustration on the right shows the pallets in the stowed position. In this concept there is a 

retractable arrangement of airbags on each side of the individual seats. The airbags are mounted 

so that they pivot about one corner, and there is a mechanical linkage that connects them so that 

they move together as a group. In the up position they are held by a spring clip and by manually 

pushing any one of the airbags down the entire group can be quickly stowed. The airbag position 

can be changed for taller or shorter crew members by using an adjustable length linkage between 

airbag pallets. By adjusting the length of the linkage (Figure 7.2-67) the erect position of the 
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paddle can be pivoted up and down for different crew member sizes. This would allow the 

seating system to be adjusted in flight or on the ground to accommodate changing crew sizes. 

Figure 7.2-66. Airbag Concept shown in a CEV Six-Seat Configuration 

Figure 7.2-67. Airbag Sketch of Adjustable Length Linkage 

Simulations for impacts in the Y-axis direction were conducted using the Hybrid III FE model, 

harness, and seat as was used for the Z-axis study in the previous section. The purpose of 

performing simulations in the sideways direction was to evaluate the effect of various 

combinations of side impact constraint devices on crew member injury, and to assess the general 

effectiveness of the FE ATD for performing this type of study. The Hybrid III FE ATD is not 

ideally suited for simulating side impacts; however, since this model was already generated for 

the work performed in the previous section it was expedient to re-use the model for side impacts. 

The results presented in this section may not be as accurate as if they had been computed with a 

better-suited ATD such as the SID. However, for performing a first order comparison of the 

effect of side constraints, this model was adequate. 
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Figure 7.2-68 shows the four of the five configurations that were used to assess the effectiveness 

of side impact constraints. The baseline configuration was an unconstrained crew member in the 

five-point harness. The second configuration was comprised of thin pads with a three-inch gap 

between the crew member’s head and shoulders. The third configuration (not shown in the 

figure) was thicker pads with a one-inch gap between the crew member’s head and shoulder. 

The fourth configuration was comprised of thick pads in direct contact with the crew members 

head and shoulders, and the fifth configuration was identical to the fourth with hand and feet 

constraints added to retain the crew member’s arms and legs from significant flaying. For this 

preliminary study the pads properties and thicknesses were not optimized and the results are only 

indicative of overall trends. All of the simulations performed in this section were done using the 

input acceleration profile that generated a high Brinkley Dynamic Response Method injury in the 

sideways direction. 

 Figure 7.2-68. Comparison Among Padded and Constrained Crew Member for Sideways Loading 

Figure 7.2-69 and Table 7.2-22 show the results for the first configuration. Figure 7.2-69 shows 

the ATD response at three intermediate time steps for the model without any side padding. 

Large excursions of the head, arms, and legs are shown during both the initial impact and the 

rebound. The results in Table 7.2-22 agree with the Brinkley Dynamic Response Method 

prediction of high injury considering that four of the seven injury criteria are exceeded including 

the Chest Severity Index (CSI), the thorax maximum allowable g level, and the neck flexion and 

extension. It is important to note that the neck injury criteria are intended for frontal impacts and 

have been loosely applied to the present side impact application. For side impact there are no 

clearly defined neck moment criteria available from industry or defined in CxP specifications. 
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However, the NESC team felt that the industry frontal impact neck criteria could be reasonably 

applied, as an approximation, to side impacts. Additional study of neck injury criteria is required 

before firm criteria can be established by the CxP. 

Figure 7.2-69. Hybrid III ATD Response at Intermediate Time Steps for Side Impact (No Side 

Padding) 

Figure 7.2-70 and Table 7.2-22 show the results for the second configuration. This type of 

support is similar, in concept, to what is currently used for race car seat side supports, and has 

been proven highly successful in reducing injury during relatively violent high-speed race car 

crashes and was thought to have similar benefits for CM crew members also undergoing side 

impact loadings. Figure 7.2-70 shows the ATD response at three intermediate time steps for the 

model with thin side padding. The large excursions of the head, arms, and legs that were seen 

for the configuration without side padding is eliminated with the thin pad configuration. 
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Table 7.2-22. Comparison Among ATD Configurations Using Input Acceleration Profile that 

Produces High Brinkley Dynamic Response Method Injury in Lateral (Sideways) Direction 

The results in Table 7.2-22 show that the thin side pads help to reduce the neck moments, which 

was expected since the side pads prevent the head from bending over. However, the Head Injury 

Criteria (HIC), CSI, thorax, and pelvis acceleration criteria all exceed their allowable limits and 

are considerably larger than for the configuration without any pads. These results were also 

expected since the crew member head and chest impact into the relatively thin and stiff pads 

leading to short durations of large accelerations. 

Figure 7.2-70. Hybrid III ATD Response at Intermediate Time Steps for Side Impact (Thin Side 

Padding) 
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The configuration with the thicker pads and one-inch gap between the pads and head and chest 

provides a higher level of injury protection than the previous two cases. For this configuration, 

only the CSI and the thorax maximum allowable acceleration were exceeded. The thicker pads 

provide so much reduction in head acceleration that the head injury criterion is reduced by an 

order of magnitude over the configuration with thin pads. For the thick pad configuration where 

the pads are in contact with the head and shoulders, the head injury criteria and the CSI are 

further reduced. However; the pelvis acceleration is increased and is beyond the allowable limit. 

For the thick pads and constrained arm and leg configuration, all of the observed injury criteria 

are reduced compared to the other configurations except for the CSI and the thorax acceleration. 

Both of these criteria exceed their allowable limits. 

 

In general, introduction of the side supports eliminates the large head excursion and neck 

hyperextension. Neck hyperextensions are the main cause of severe injury for the case without 

lateral support, so it appears promising that lateral support will improve survivability for Y-axis 

direction impact. However, the body must still mitigate the crash, so while neck forces are 

reduced due to reduced head and neck extension, other body forces and moments may show 

increases. This is supported by the fact that padding had the effect of reducing the neck moment 

while increasing the CSI. The key design goal is to keep any individual force, moment or 

acceleration from exceeding the allowable injury level, which will require tradeoffs in the 

design. With further design improvements it may be possible to reduce the chest and thorax 

loadings to within allowable limits; however, for the designs that were considered, both the 

Brinkley Dynamic Response Method model and the FE ATD predicted high levels of injury. 

A significant difference between the Brinkley Dynamic Response Method model and the FE 

ATD is that the Brinkley model cannot directly determine the effects of adding constraints such 

as side padding while the FE ATD can simulate somewhat realistically the distributed motion of 

the body and predict accelerations and loads at discrete body locations. Furthermore, the 

Brinkley model only predicts the level of injury, not the injury location or type. The FE ATD 

provides a major advantage over the Brinkley model in that the ATD may be used to assess the 

effects of different seat designs, constraint systems, helmets, and crew protection systems. 

7.2.6.6 X-Axis (chest in/out) Acceleration Simulations and Results 

A comparison between the Brinkley Dynamic Response Method model and the FE ATD was 

made in the X-axis direction using an applied acceleration profile obtained from a maximum 

drag abort situation. This situation occurs when an emergency occurs during ascent and the CM 

must be separated from the launch vehicle using the Launch Abort System (LAS). When the 

LAS is activated, the crew members are pushed into their seat backs. Once the LAS rocket’s 

fuel is depleted the aerodynamic drag reverses the acceleration on the crew members and they 

are pulled out of their seats in the opposite direction. The Hybrid III FE ATD is probably most 

suited for loadings in the X-axis direction since this is the direction where the ATD was 

originally designed for automotive frontal impacts. The acceleration profile for the maximum 
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drag abort and the corresponding FE ATD response at intermediate time steps is shown in 

Figures 7.2-71 and 7.2-72, respectively. 

As expected, during LAS rocket firing the crew member is pushed into their seat and there is 

minimal motion of the head, arms, or legs. In the later stage when aerodynamic drag is dominant 

and the crew member is pulled out of his seat, both the head and arms and legs are extended 

leading to the potential for larger neck forces and head accelerations. Note, this particular model 

contained overly stiff elbow joints which explains the limited extension of the arms. 

Additionally, flailing of the arms and legs in the close confines of CM could be an issue since the 

arms and legs may impact surrounding structures or strike other crew members. To compensate 

for flailing, hand and foot restraints will be an important design consideration. The large head 

movement could also be an issue and some type of head/helmet constraint may be required. 

Figure 7.2-71. Acceleration Profile for Maximum Drag Abort 
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Figure 7.2-72. Hybrid III ATD Response at Intermediate Time Steps for Maximum Drag Abort 

Figure 7.2-73 shows the maximum drag abort acceleration that was applied to the crew member 

seat along with the resulting Brinkley Dynamic Response Method and the FE ATD response in 

the X-axis direction. Both the Brinkley and the ATD response follow, reasonably close, both the 

input acceleration shape and magnitude. The Brinkley response closely tracks the input 

acceleration profile during the rocket-firing phase then oscillates about the input acceleration 

during the aerodynamic drag phase. The Brinkley model employs an approximately 10 Hz 

oscillator, so it is expected that the response after the load reversal would oscillate and that the 

oscillations would be near 10 Hz. The maximum Brinkley dynamic response is 20 g, which is 

within the Brinkley low risk of injury range for X-axis (chest in/out) accelerations. The FE ATD 

response is within the same general range as the Brinkley response except for numerous short 

duration acceleration spikes through the transient response. The ATD response has been filtered 

to eliminate response above 180 Hz and further reductions, or elimination, of at least some of 

these acceleration spikes could be minimized by adding more damping to the ATD model, and 

model refinement through the inclusion of a higher fidelity seat model and seat padding. Even 

with these outlying acceleration peaks, the computed CSI is below the allowable limit (212 < 

700) leading to a consistent result with the Brinkley Dynamic Response Method prediction of a 

low probability of injury. 
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Figure 7.2-73. Acceleration Profile for Maximum Drag Abort, Brinkley Dynamic Response Method 

and FE ATD Chest Acceleration 

7.2.6.7 Applicability of FE ATDs

The FE ATD used in conjunction with the Brinkley Dynamic Response Method provides a 

useful set of tools for predicting potential crew injuries during CM landings. In addition these 

tools enable the design and evaluation of alternate crew protection systems. Specifically: 

 The Brinkley Dynamic Response Method reasonably ensures an acceptable environment 

for crew members in a five point harness. However, it has limited ability to provide 

insight into additional or modified crew restraint and protection systems. The Brinkley 

model has been modified to include the general effect of sideways crew member support. 

However, the Brinkley Model does not distinguish between types of support or their 

differences in effectiveness in preventing injury. 

 FE analysis and ATD tests are capable of providing insight into alternate crew member 

injury protection systems. The FEA model and the physical ATDs can be used to assess 

the effects of variations of restraint and support in a comparative manner and if validated 

with human response data, can provide quantitative assessments of crew injuries. 

 Additional evaluation of industry standard injury criteria should be initiated to ascertain 

the applicability of these criteria to crew member protection. 

 Practical methods for implementing additional crew protection appears viable and should 

be further developed. 
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8.0 Findings and Recommendations 

The Findings and NESC Recommendations first for the Phase I and Phase II portions of this 

study are presented. 

8.1 Phase I - Comparison of 604, TS-LRS001, and Apollo LRS 

8.1.1 Findings / Observations for Phase I - Comparison of 604, TS-LRS001, and Apollo 

LRS 

Note: This material was extracted from Appendix A, which represents the only documentation of 

Phase I. Pertinent sections of this appendix were brought forward into the body of this report. 

F-1. C2 (LM 604) - Retrorocket Extraction - Design susceptible to retrorocket/capsule 

contact and damage during extraction, especially under high body-rate conditions: 

 Risk of damage to radar antennas, retrorocket body/nozzles, and other components 

mounted near the CM apex. 

 Risk of severing pyro firing lines, radar avionics to antenna RF transmission lines, 

and parachute risers. 

 Extraction failure would result in failure of main chute to deploy properly due to 

series connection between the two subsystems. 

 Mars lander heritage benefits toward concept feasibility not applicable due to 

utilization of rate-limiting devices to lower the lander from confluence rockets after 

chute inflation. 

F-2. C3 (LM 604) - C3 - Winds and Residual Horizontal Velocity - Insufficient data has 

been produced to demonstrate that not including horizontal retrorockets in the design is 

viable. 

 Operational limit of 31 fps is too close to CM limits for crew safety and is relatively 

low; DSNE requirement > 45 fps. 

 Consequences of CM rollover are not well understood and could be potentially 

severe. 

Without horizontal velocity abatement, crew safety is critically dependent on the 

accuracy of a 3-hour weather forecast. 

 If weather forecast was 99.9 percent accurate, still non-trivial risk for possible 

rollover and threat to crew safety. 

Use of horizontal retrorockets on the TS-LRS001 is not risk free. 

 Mars lander experience is that performance dispersions and failure modes with 

horizontal retrorockets have potential to compound the problem. 
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F3: C4 (LM 604 and TS-LRS-001) - Retrorocket Firing Chain – Applicable to both 

baselines. Dispersions on retrorocket burn completion conditions have potential to cause 

load limits to be exceeded. 

 Difficult to control performance using high-impulse solids to deliver the CM to nearzero velocity near the ground. 

 Small knowledge errors in parameters such as thrust profile, altitude/velocity state, 

and attitude/rate state may lead to late firing, including impact prior to ignition. 

Retrorocket firing time critically dependent on accuracy of altitude and velocity sensors. 

The LM 604 radar antenna location on the retrorocket pack may result in degraded 

accuracy with potential to degrade ignition time calculations. 

F-4. C1 (LM 604) - Main Chute Extraction - Use of a single drogue to extract all main 

chutes in the current design involves rigging the mains together in a way where hang-up 

of a single chute has the potential to fail the extraction of all the others resulting in LOC. 

F-5. C5 (LM 604 and TS-LRS-001) – Apex/FBC Separation - Post separation impact with 

CM has potential to cause severe damage: 

 Current design is based on Apollo heritage of separation thrusters and pilot parachute. 

 The Apollo Program was only able to provide limited satisfaction in meeting the 

requirements of no re-contact, even after extensive work. 

F-6. C6 (LM 604) - Retrorocket Plume Impingement - Exhaust plume may be swept into 

riser lines or across the top of the capsule potentially resulting in safety related damage. 

F-7. C7 (LM 604 and TS-LRS-001) CM Orientation Event - Dispersions in ability to orient 

the CM correctly for touchdown may result in unacceptable crew loads: 

 Unclear if dynamics due to issues such as swivel friction, parachute dynamics, hang 

angle, and available thruster torque are acceptable. 

 Not yet demonstrated acceptable Y-axis (cross spine) accelerations for all crew 

members. 

F-8. C8 (LM 604) - Drogue Stabilization - During a pad abort scenario, timeline may be 

insufficient to detect a failed drogue to allow deployment of the back-up drogue: 

 Residual attitude and rates, potentially high due to pad abort, may hamper 

deployment of main chutes. 

 Detection scheme for failed drogue is unclear. 

F-9. C9 (TS-LRS-001) - Impact Survivability – The TS-LRS001 baseline could sustain 

impact or puncture from ground terrain that may damage critical aft bay components and 

pose safety risk and/or egress hazard. 
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 Tanks and unfired retrorockets may be vulnerable to puncture due to jettison of heat 

shield which would have served as a protective barrier 

The LM604 retrorocket pack may re-contact the capsule and pose safety risk 

 Retrorockets (used or unspent) may damage up-righting devices, location aids, or 

egress hatch. 

F-10. C10 (TS-LRS-001) – Heat Shield Separation - Potential for heat shield contact and 

damage of retrorocket nozzles on pad abort scenarios: 

 Mars lander heritage heat shield separation technique had low clearance margins at 

high attitude rates. 

Radar may track heat shield for up to 1,300 feet after separation and may prevent timely 

lockup on ground for an on pad abort scenario. 

8.1.2 NESC Recommendations for Phase I - Comparison of 604, TS-LRS001, and Apollo 

LRS 

R-1. C2 (LM 604) - Retrorocket Extraction – If the retrorocket system is to remain as part of 

the Project baseline, then utilize ejection mechanisms or other deployment aids; increase 

volume available to this system in the forward bay in order to accommodate devices. 

(F-1) 

 Consider strong design bias toward storing (and therefore deploying) retrorocket in an 

upright position in the forward bay in order to minimize deployment complexity. 

 Consider placement of retrorockets in location that does not require dynamic 

deployment. 

However, the NESC team strongly recommends that the Apollo parachute architecture be 

used since it is better understood and characterized based on its heritage. 

R-2. C3 (LM 604, TS-LRS-001) - Winds and Residual Horizontal Velocity - Conduct forecast 

reliability study. (F-2)

 Wind speeds restrictions may need to be reduced below 31 fps to reduce the risk of an 

incorrect forecast. 

 Evaluate whether the CxP paradigm of having at least (1 of 6) sites available still 

holds with more stringent wind forecast restrictions. 

 Implement LRS architecture that restricts landings to low winds, but design CM to be 

more tolerant to winds. 

 Decision not to use retrorockets for pad abort is inconsistent with NESC Land versus 

Water Assessment recommendation to preserve land-landing and ensure crew safety. 

 Perform thorough analysis of the LM 604 retrorocket firing dynamics to understand 

potential impact of further aggravating horizontal velocity situation. Conduct study 
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of the TR-LRS001 regarding performance dispersions on horizontal retrorockets, 

consider limiting use to high winds.

R-3. C4 (LM 604 and TS-LRS-001) - Retrorocket Firing Chain - Perform error budget 

analysis, and chose ignition altitude and proper retrorocket sizing to ensure proper 

touchdown conditions across range of dispersions. (F-3)

 Verify that CM impact attenuation system capable of absorbing dispersions in an 

acceptable manner. 

 Consider sustainer retrorocket burn, thrust tailoring, or thrust termination techniques. 

 Develop requirements for radar (or equivalent) as soon as possible. 

 Ensure proper field of view for sensor elements. 

 Consider use of retrorockets for pad abort to avoid crew injury for contingency landlanding scenarios. 

R-4. C1 (LM 604) - Main Chute Extraction Recommendation - Consider independent 

extraction for each of the main canopies. (F-4)

R-5. C5 (LM 604 and TS-LRS-001) - Apex Cover Separation - Design separation scheme to 

ensure that cover will clear capsule without contact, and then be laterally translated out of 

wake flow. (F-5)

 If a pilot chute is required, consider a time delay method in mortar firing to allow it to 

clear the docking tunnel. 

 Conduct study to determine possible safety risks to ground crew and civilian 

population.

R-6. C6 (LM 604) - Retrorocket Plume Impingement - Conduct exhaustive plume analysis, 

including acoustics, and determine adequacy of separation distance between capsule and 

retrorockets. (F-6)

 Verify adequacy of riser attachment locations relative to the retrorocket nozzles 

 Consider adding requisite retrorocket plume protective devices for riser lines and 

exposed components on forward deck. 

R-7. C7 (LM 604 and TS-LRS-001) CM Orientation Event - Accelerate modeling efforts and 

testing to define system design requirements in this area. (F-7)

R-8. C8 (LM 604) - Drogue Stabilization - Revisit Apollo Program heritage data. (F-8)

 Dual drogue deployment is not materially more susceptible to deployment 

entanglement than single drogue system. 

 Consider parallel drogue deployment at divergent angles consistent with the heritage 

Apollo parachute subsystem design.
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R-9. C9 (TS-LRS-001) - Impact Survivability - Consider protective structural elements to 

mitigate risk of impact to aft-bay or consequences of retaining heat shield. (F-9)

 Consider design of protective elements that allow for “high” horizontal velocity, even 

with use of horizontal retrorockets. 

 Consider method to safe unfired pyros at time of touchdown.

R-10. C10 (TS-LRS-001) - Heat shield Separation - Increase energy in separation devices to be 

robust to high attitude rates. (F-10)

 Add protection bumpers or other guards near retrorocket nozzle locations and 

consider heat shield lockup scenario when selecting altimetry architecture. 

8.2 Phase II - Integrated LRS Risk Assessment 

The Findings, Observations, and NESC Recommendations for each of the six Phase II tasks are 

presented. 

8.2.1 Task 1: Entry, Descent, and Landing System Quantitative Risk Assessment 

Findings, Observations, and NESC Recommendations 

8.2.1.1 Task 1 Findings and Observations 

F-1. The 606 design uses parachutes to make sure that the FBC is out of the way for drogue 

parachute deployment. Parachute system reliability is a risk driver and having a 

requirement for such a system significantly increases the risk to the CM. 

F-2. Task 3 of this assessment presented the predictability of horizontal winds at the landing 

site poses some risk in that the CMs horizontal landing speed could be exceeded due to 

mispredicted winds, potentially contributing to an increased PLOC. 

F-3. The landing systems may experience reduced capability due to failures. 

8.2.1.2 Task 1 NESC Recommendations 

R-1. Improved FBC Release (Apex cover) - Ensure the FBC will not interfere with drogue 

chute deployment with adequate reliability (i.e., without parachutes) and prevents CM recontact (including main parachutes). (F-1) 

R-2. Increased Horizontal Velocity Attenuation - Protect against unpredicted high horizontal 

winds / gusts. (F-2) 

 Horizontal velocity attenuation or increased CM capability will reduce the main risk 

driver for land landing systems, and provide margins that could be used to increase 

stability of the CM during landing. 

R-3. Improved Occupant Protection – Improve crew protection as a functional redundancy for 

landing system faults (e.g., heat shield separation, retrorockets, airbags, etc.). (F-3) 
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 Risk analysis has shown that active landing attenuation systems are susceptible to 

failure. Improved protection for the crew provides additional safeguards for 

unforeseen failure modes that may occur.

8.2.2 Task 2: Landing System T&V Data Analysis 

8.2.2.1 Task 2 Findings 

F-4. Logical evolution of early test plans in place for parachute and airbag subsystems will 

likely result in a robust T&V strategy for design certification of those individual 

subsystems: 

 Current focus is on early developmental testing to understand design space, not on 

long-term T&V planning. 

 Early developmental testing strategy appears to be robust and well conceived. 

F-5. Further investigation is warranted in the area of integrated LRS evaluation: 

 Unclear to as to which Orion Project element (either NASA or LM) has been charged 

with evaluating end-to-end LRS performance, including interactions between 

subsystems and environments. 

 Current plans to evaluate end-to-end performance of LRS are dependent on the 

utilization of a few major flight test opportunities (e.g., PA-1, AA-1, etc.). 

o Although these flight tests are valuable, they will not provide enough 

opportunities to exercise integrated LRS system under expected range of dynamic 

conditions. 

o Unclear if Orion Project has plans to add additional flight-like test opportunities 

dedicated toward evaluation of integrated, end-to-end LRS performance. 

8.2.2.2 Task 2 NESC Recommendations 

R-4. Explicitly add a test program to exercise end-to-end performance of LRS with emphasis 

on interactions between subsystems: (F-4 and F-5)

 Utilize air drops of LRS and CM hardware with flight realistic geometry and 

performance characteristics. 

 Executed across a wide range of dynamic conditions bounding expected attitudes, 

rates, and winds for nominal and abort landings to demonstrate system design 

margins. 

 Coupled to an integrated performance modeling task to extend the number of cases 

that can be evaluated. 

R-5. Formulate a working group charged with the responsibility for verification of end-to-end 

LRS performance with emphasis on subsystem-to-subsystem and subsystem-to-CM 

interactions. (F-4 and F-5)
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R-6. Adopt a strategy of developing specific LRS tests based on evaluation and verification 

needs as opposed to forcing the verification plan to conform to already existing “all 

encompassing” tests. (F-4 and F-5) 

R-7. T&V Scenario for FBC Release: (F-4 and F-5)

 Ensure that FBC can separate without striking and damaging forward bay 

components. 

 Ensure that FBC does not come back and strike capsule at a later time. 

 Ensure that FBC release does not induce adverse attitude rates on capsule that may 

affect downstream events such as parachute deployment. 

R-8. T&V Scenario for Parachute Extraction and Deployment: (F-4 and F-5)

 Ensure that capsule attitude, rates, physical configuration, and operating 

characteristics do not interfere in any way with chute extraction and deployment 

 Ensure that drogue parachute can stabilize capsule. 

 Ensure that capsule and components on upper deck can still function after exposure to 

parachute opening loads. 

 Ensure that parachute deployment chain of events does not result in adverse attitude 

or rates affecting downstream events. 

 Ensure that each component of the LRS satisfies its design requirements. 

R-9. T&V Scenario for heat shield Jettison (if required): (F-4 and F-5)

 Ensure that heat shield can separate without striking and damaging aft bay 

components. 

 Ensure that heat shield does not come back and strike the capsule at a later time. 

 Ensure that the jettisoned heat shield does not interfere with downstream events such 

as radar acquisition of the ground. 

R-10. T&V Scenario for Parachute Descent: (F-4 and F-5)

 Ensure that capsule attitude and dynamics under parachute does not result in adverse 

conditions for landing. 

 Ensure that RCS thruster firings are effective in orienting capsule 

 Ensure that RCS thruster firings and/or venting of residual gasses does not damage or 

degrade the parachute or airbag materials. 

R-11. T&V Scenario for Airbag Inflation (if required): (F-4 and F-5)

 Ensure that airbags can be extracted and inflated without damaging components in the 

aft bay 
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 Ensure that the inflated airbags do not interfere with other component functions in the 

aft bay and vice versa. 

R-12. T&V Scenario for Landing and Impact (for water or land): (F-4 and F-5)

 Ensure that impact loads are handled properly by entire end-to-end attenuation chain 

(e.g., airbags, structure, crew couch, and suit).

 Ensure that impact does not compromise structural integrity. 

 Ensure that parachute dynamics during landing and impact process do not result in 

unexpected and adverse consequences. 

 Determine through test and analysis system performance in off- nominal and failure 

case conditions. 

8.2.3 Task 3: Landing Site Accessibility and Availability Data Analysis Findings, 

Observations, and NESC Recommendations 

8.2.3.1 Findings 

F-6. The CxP’s land-landing site network provides high degree of physical access. 

F-7. 14 percent reduction in L/D results in 25 percent to 30 percent loss of individual site 

access. 

F-8. Raise maneuvers enable a significant portion of network access: 

 50 percent for L/D of 0.35, 85 percent for L/D of 0.30. 

 Raise required for docking module disposal - otherwise, SM deorbit would be 

preferred. 

F-9. Night restrictions will reduce access and create periodic network gaps. 

F-10. Separation between landing opportunities for ISS missions is generally measured in days: 

 Requires treating each access as a unique opportunity. 

 ISS support missions unlikely to undock without high probability of deorbit. 

F-11. No obvious CONUS land sites to support return from 28.5 degree orbits assuming no 

delta-v available from SM. 

F-12. The SLAAM tool allows the weather (primarily horizontal winds and gusts) to be 

included to show effects on landing opportunities. 

8.2.3.2 NESC Recommendations 

R-13. Implement forecasting tools and procedures at each candidate land-landing to ensure 

capability is understood when operations are required. (F-6 through F-12)
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R-14. Consider implementing additional crew protection in anticipation of misforecast of 

landing site winds / gusts. (F-6 through F-12)

R-15. Evaluate supplemental contingency crew protection as a means of avoiding operationally 

over constraining wind placards in the near term. (F-7, F-8)

R-16. Develop a test program that can accommodate ISS support schedules and be prepared to 

capitalize on performance envelope expansion opportunities as they emerge. (F-7, F-11)

R-17. Select water based landing location that takes advantage of ascending and descending 

pass opportunities. (F-7, F-11) 

8.2.4 Task 4: CM Roll Control in Preparation for Landing Data Analysis Finding and 

NESC Recommendation 

8.2.4.1 Task 4 Finding

F-13. Program tests show that the CEV Project CM RCS has sufficient torque to meet the +30 

degree roll requirement with current parachute harness / riser designs. 

8.2.4.2 Task 4 NESC Recommendation 

R-18. Endorsed the use of roll control: (F-13)

 Some form of roll control is useful to orient the vehicle Z plane with the direction of 

travel to maximize crew safety 

 Limited cabin volume and minimal human impact acceleration tolerance in the Y axis 

(lateral impacts) 

 During water-landings the CM can be oriented to minimize crew impact 

accelerations, which supports meeting reduced impact acceleration requirements for 

de-conditioned crew. 

8.2.5 Task 5: Parachute Release Times during Land-Landing Findings and NESC 

Recommendation 

8.2.5.1 Findings 

F-14. For landing conditions where there is a horizontal wind, retaining the parachutes has a 

detrimental effect on CM stability since the drag force on the parachutes can pull the 

vehicle over. 

F-15. The effect of rigging and parachute flexibility has minimal effect on the acceleration and 

roll response so that the trends reported should be applicable to most parachute system 

designs. 

8.2.5.2 Task 5 NESC Recommendation 

R-19. Require an automated parachute release system. (F-14 and F-15)
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 An automated system is required since the release may be required to occur within 

0.50 seconds of touchdown, which is not enough time for a crew operated manual 

release. 

 An automated release system would be a critical function which must ensure no 

unplanned release. 

8.2.6 Task 6: Crew Protection System Enhancements Data Analysis Findings and NESC 

Recommendations 

8.2.6.1 Findings 

F-16. The FEA modeling approach shows promise as a tool for seat designers. 

F-17. In addition to developing tools, it is important that a practical mean exist for 

implementing these design solutions. 

F-18. The Brinkley Dynamic Response Method criteria define an acceptable environment for 

crews restrained in a specific way. It does not provide insight into additional or modified 

crew restraint systems. 

8.2.6.2 Recommendations 

R-20. Development and adoption of contemporary design tools and techniques is necessary. 

(F-16) 

 The Brinkley Dynamic Response Method criteria define an acceptable environment 

for crews restrained in a specific way. It does not provide insight into additional or 

modified crew restraint systems.

 Use of FEA and ATD tests allow designers to evaluate the effectiveness of alternate 

restraint systems.

R-21. Assuming that the potential for unplanned landings on land cannot be eliminated for a 

CM configured for water penetration or other compromised landing configurations, 

additional emergency crew protection is highly desirable. (F-17)

R-23. Develop practical methods for implementing additional crew protection appear realistic 

and should be developed (e.g., lateral support). (F-17)

9.0 Alternat e Viewpoints 

There were no alternate viewpoints. 

10.0 Other Deliverables 

There were no other deliverables. 

11.0 Lessons Learned 

There were no lessons learned. 
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12.0 Definition of Terms 

Corrective Actions Changes to design processes, work instructions, workmanship practices, 

training, inspections, tests, procedures, specifications, drawings, tools, 

equipment, facilities, resources, or material that result in preventing, 

minimizing, or limiting the potential for recurrence of a problem. 

Finding A conclusion based on facts established by the investigating authority. 

Lessons Learned Knowledge or understanding gained by experience. The experience may 

be positive, as in a successful test or mission, or negative, as in a mishap 

or failure. A lesson must be significant in that it has real or assumed 

impact on operations; valid in that it is factually and technically correct; 

and applicable in that it identifies a specific design, process, or decision 

that reduces or limits the potential for failures and mishaps, or reinforces a 

positive result. 

Observation A factor, event, or circumstance identified during the assessment that did 

not contribute to the problem, but if left uncorrected has the potential to 

cause a mishap, injury, or increase the severity should a mishap occur. 

Alternatively, an observation could be a positive acknowledgement of a 

Center/Program/Project/Organization’s operational structure, tools, and/or 

support provided. 

Problem The subject of the independent technical assessment. 

Proximate Cause The event(s) that occurred, including any condition(s) that existed 

immediately before the undesired outcome, directly resulted in its 

occurrence and, if eliminated or modified, would have prevented the 

undesired outcome. 

Recommendation An action identified by the NESC to correct a root cause or deficiency 

identified during the investigation. The recommendations may be used by 

the responsible Center/Program/Project/Organization in the preparation of 

a corrective action plan. 

Root Cause One of multiple factors (events, conditions, or organizational factors) that 

contributed to or created the proximate cause and subsequent undesired 

outcome and, if eliminated or modified, would have prevented the 

undesired outcome. Typically, multiple root causes contribute to an 

undesired outcome. 
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13.0 Acr onyms List 

AAMRL Armstrong Aerospace Medical Research Laboratory 

AFB Air Force Base 

ATD Anthropomorphic Test Devices 

CEV Crew Exploration Vehicle 

CEVPO CEV Project Office 

CM Crew Module 

CSI Chest Severity Index 

CxP Constellation Program 

DR Dynamic Response 

DRI Dynamic Response Index 

FBC Forward Bay Cover 

FEA Finite Element Analysis 

GEM Government Equipment and Materials 

GN&C Guidance Navigation and Control 

HSIR Human System Integration Requirements 

JPL Jet Propulsion Lab 

L/D Lift to Drag Ratio 

LAS Launch Abort System 

LEO Low Earth Orbit 

LM Lockheed Martin 

LOC Loss of Crew 

LRS Landing and Recovery System 

MMOD Micro Meteoroid Orbital Debris 

NESC NASA Engineering and Safety Center 

NM Nautical Miles 

NOAA National Oceanic and Atmospheric Administration 

NRB NESC Review Board 

PDR Preliminary Design Review 

PLOC Probability of Loss of Crew 

RAC-3 Requirements Analysis Cycle-3 

RCS Reaction Control System 

RID Rear Impact Dummy 

RTE Return to Earth 

SID Side Impact Dummy 

SLAAM Spacecraft Landing Accessibility and Availability Model 

SM Service Module 

STK Satellite Toolkit 

T&V Test and Verification 

TIM Technical Interchange Meeting 

ZBV Zero Based Vehicle 
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14.0 Ref erences 

1. Jones, R. H., “Landing Impact Attenuation for Non-Surface-Planning Landers,” NASA 

SP-8046, April 1970. 

2. Brown, William K., et al, “Human Response to Predicted Apollo Landing Impacts in 

Selected Body Orientations,” Aerospace Medicine, pp. 394-398, April 1966. 

3. Eiband, A. Martin, “Human Tolerance to Rapidly Applied Accelerations: A Summary of 

the Literature,” NASA Memorandum 5-19-59E, June 1959. 

4. Parker, James F. and West, Vita R., “Bioastronautics Data Book,” NASA SP-3006, 1973 

5. Stech, E. L. and Payne, P. R., “Dynamic Models of the Human body, AAMRL-TR-66-

157, Aerospace Medical Research Laboratory, Wright-Patterson Air Force Base, Ohio, 

1969. 

6. von Gierke, H. E., “Response of the Body to Mechanical Forces – An Overview,” 

Lectures in Aerospace Medicine, 6th Series, School of Aerospace Medicine, Brooks Air 

Force Base, Texas, pp.325-344, 1967. 

7. Brinkley, J. W., and Shaffer, J. T., ”Dynamic Simulation Techniques for the Design of 

Escape Systems: Current Applications and Future Air Force Requirements,” Aerospace 

Medical Research Laboratory Report AMRL TR 71-292, Wright Patterson Air Force 

Base, Ohio, December 1971. 

8. Brinkley, J. W., and Mosher, S. E., “Development of Acceleration Exposure Limits for 

Advanced Escape Systems,” AGARD CP-472, April 1989 

9. Personal communications with James W. Brinkley, 2007. 

10. Coltman, J. W. et. al., “Aircraft Crash Survival Design Guide – Volume II – Aircraft 

Design Crash Impact Conditions and Human Tolerance,” USAAVSCOM TR 89-D-228, 

December 1989. 

11. Bowman, B.M., “Aircrew Ejection Injury Analysis and Trauma Assessment Criteria”, 

University of Michigan Transportation Research Institute, UMTRI-93-8, March 1993. 

12. Kai-Uwe Schmitt, Peter F. Niederer, Felix Walz: Trauma Biomechanics 

Springer Verlag, Berlin Heidelberg New York, 2004 

13. Peat, Christopher. “ISS-Orbit Data,” Heavens Above Retrieved 22 May 2007 

http://www.heavens-above.com/orbitdisplay.asp?satid=25544 

14. Bairstow, Brian (2007) “Wind Data for 8 Sites”, 24 July 2007. Email. 

 
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Volume II Appendices 

Appendix A. Crew Exploration Vehicle Integrated Landing and Recovery Evaluation 

Appendix B. Task 2 T&V Assessment 

Appendix C. Task 1 Integrated Entry, Descent and Landing Risk Matrix 
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Appendix A. Crew Exploration Vehicle Integrated Landing and 

Recovery Evaluation 
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Appendix B. Task 2 T&V Assessment 
NASA Engineering and Safety Center 

Technical Assessment Report 

Document #: 

NESC-RP06-060 

Version: 

1.0 

Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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NESC-RP06-060 

Version: 

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Title: 

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Version: 

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Technical Assessment Report 

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Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Version: 

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Title: 

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Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

Document #: 

NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

1.0 

Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Technical Assessment Report 

Document #: 

NESC-RP06-060 

Version: 

1.0 

Title: 

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Technical Assessment Report 

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NESC-RP06-060 

Version: 

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Title: 

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Appendix C. Task 1 Integrated Entry, Descent and Landing Ri

(Insert 11x17 paper) 
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NESC Request No.: 06-060-E 
REPORT DOCUMENTATION PAGE Form Approved

OMB No. 0704-0188

2. REPORT TYPE 

Technical Memorandum

 4. TITLE AND SUBTITLE

Constellation Program (CxP) Crew Exploration Vehicle (CEV) Project 

Integrated Landing System

5a. CONTRACT NUMBER

 6. AUTHOR(S)

Baker, John D.; Yuchnovicz, Daniel E.; Eisenman, David J.; Peer, Scott G.; 

Fasanella, Edward L.; Lawrence, Charles

 7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

 9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Washington, DC 20546-0001

 8. PERFORMING ORGANIZATION

 REPORT NUMBER

L-19798 NESC-RP-06-060

10. SPONSOR/MONITOR'S ACRONYM(S)

NASA

13. SUPPLEMENTARY NOTES

12. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified - Unlimited

Subject Category 16 - Space Transportation and Safety

Availability: NASA CASI (443) 757-5802

19a. NAME OF RESPONSIBLE PERSON

STI Help Desk (email: help@sti.nasa.gov)

14. ABSTRACT

Crew Exploration Vehicle (CEV) Chief Engineer requested a risk comparison of the Integrated Landing System design 

developed by NASA and the design developed by Contractor- referred to as the LM 604 baseline. Based on the results of this 

risk comparison, the CEV Chief engineer requested that the NESC evaluate identified risks and develop strategies for their 

reduction or mitigation. The assessment progressed in two phases. A brief Phase I analysis was performed by the Water versus 

Land-Landing Team to compare the CEV Integrated Landing System proposed by the Contractor against the NASA 

TS-LRS001 baseline with respect to risk. A phase II effort examined the areas of critical importance to the overall landing 

risk, evaluating risk to the crew and to the CEV Crew Module (CM) during a nominal land-landing. The findings of the 

assessment are contained in this report.

15. SUBJECT TERMS

Constellation Program; Crew Exploration Vehicle; Crew Module; Loss of Crew; Impact Dummy

18. NUMBER

 OF 

 PAGES

319

19b. TELEPHONE NUMBER (Include area code)

(443) 757-5802

a. REPORT

U

c. THIS PAGE

U

b. ABSTRACT

U

17. LIMITATION OF 

 ABSTRACT

UU

Prescribed by ANSI Std. Z39.18

Standard Form 298 (Rev. 8-98)

3. DATES COVERED (From - To)

October 2006 - October 2009

5b. GRANT NUMBER

5c. PROGRAM ELEMENT NUMBER

5d. PROJECT NUMBER

5e. TASK NUMBER

5f. WORK UNIT NUMBER

869021.05.07.07.06

11. SPONSOR/MONITOR'S REPORT

 NUMBER(S)

NASA/TM-2009-216165

16. SECURITY CLASSIFICATION OF:

The public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources, 

gathering and maintaining the data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this 

collection of information, including suggestions for reducing this burden, to Department of Defense, Washington Headquarters Services, Directorate for Information Operations and 

Reports (0704-0188), 1215 Jefferson Davis Highway, Suite 1204, Arlington, VA 22202-4302. Respondents should be aware that notwithstanding any other provision of law, no person 

shall be subject to any penalty for failing to comply with a collection of information if it does not display a currently valid OMB control number.

PLEASE DO NOT RETURN YOUR FORM TO THE ABOVE ADDRESS.

1. REPORT DATE (DD-MM-YYYY)

01- 12 - 2009