August 2013 

NASA/TM2013-218032 

NESC-RP-09-00605 

 

Development of Autonomous Aerobraking 

Phase 2 

Daniel G. Murri/NESC 

Langley Research Center, Hampton, Virginia 
NASA STI Program . . . in Profile

Since its founding, NASA has been dedicated to the 

advancement of aeronautics and space science. The 

NASA scientific and technical information (STI) 

program plays a key part in helping NASA maintain 

this important role. 

The NASA STI program operates under the 

auspices of the Agency Chief Information Officer. 

It collects, organizes, provides for archiving, and 

disseminates NASA’s STI. The NASA STI 

program provides access to the NASA Aeronautics 

and Space Database and its public interface, the 

NASA Technical Report Server, thus providing one 

of the largest collections of aeronautical and space 

science STI in the world. Results are published in 

both non-NASA channels and by NASA in the 

NASA STI Report Series, which includes the 

following report types: 

x TECHNICAL PUBLICATION. Reports of 

completed research or a major significant phase 

of research that present the results of NASA 

Programs and include extensive data or 

theoretical analysis. Includes compilations of 

significant scientific and technical data and 

information deemed to be of continuing 

reference value. NASA counterpart of peerreviewed formal professional papers, but 

having less stringent limitations on manuscript 

length and extent of graphic presentations. 

x TECHNICAL MEMORANDUM. Scientific 

and technical findings that are preliminary or of 

specialized interest, e.g., quick release reports, 

working papers, and bibliographies that contain 

minimal annotation. Does not contain extensive 

analysis. 

x CONTRACTOR REPORT. Scientific and 

technical findings by NASA-sponsored 

contractors and grantees. 

x CONFERENCE PUBLICATION. 

Collected papers from scientific and 

technical conferences, symposia, seminars, 

or other meetings sponsored or cosponsored by NASA. 

x SPECIAL PUBLICATION. Scientific, 

technical, or historical information from 

NASA programs, projects, and missions, 

often concerned with subjects having 

substantial public interest. 

x TECHNICAL TRANSLATION. 

English-language translations of foreign 

scientific and technical material pertinent to 

NASA’s mission.

Specialized services also include organizing 

and publishing research results, distributing 

specialized research announcements and feeds, 

providing information desk and personal search 

support, and enabling data exchange services. 

For more information about the NASA STI 

program, see the following: 

x Access the NASA STI program home page 

at http://www.sti.nasa.gov

x E-mail your question to help@sti.nasa.gov

x Fax your question to the NASA STI 

Information Desk at 443-757-5803 

x Phone the NASA STI Information Desk at 

443-757-5802 

x Write to: 

 STI Information Desk 

 NASA Center for AeroSpace Information 

 7115 Standard Drive 

 Hanover, MD 21076-1320
National Aeronautics and 

Space Administration 

Langley Research Center 

Hampton, Virginia 23681-2199 

 

August 2013 

NASA/TM2013-218032 

NESC-RP-09-00605 

 

Development of Autonomous Aerobraking 

Phase 2 

Daniel G. Murri/NESC 

Langley Research Center, Hampton, Virginia 
Available from: 

NASA Center for AeroSpace Information 

7115 Standard Drive 

Hanover, MD 21076-1320

443-757-5802 

Acknowledgments 

The following were significant contributors to this effort: 

Jill L. Prince; Team Lead, LaRC 

Richard W. Powell; NESC Deputy Lead, Flight Mechanics TDT, AMA 

Alicia M. Dwyer Cianciola; LaRC Simulation, LaRC 

The use of trademarks or names of manufacturers in the report is for accurate reporting and does not 

constitute an official endorsement, either expressed or implied, of such products or manufacturers by the 

National Aeronautics and Space Administration.
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

1 of 97 

NESC Request No.: TI-09-00605 

Development of Autonomous Aerobraking 

(Phase 2) 

June 13, 2013

 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

2 of 97 

NESC Request No.: TI-09-00605 

Report Approval and Revision History 

NOTE: This document was approved at the June 13, 2013, NRB. This document was submitted 

to the NESC Director on July 30, 2013, for configuration control. 

Approved: Original Signature on File 7/30/13 

NESC Director Date 

Version Description of Revision Office of Primary 

Responsibility Effective Date 

1.0 Mr. Daniel Murri, 

NASA Technical 

Fellow for Flight 

Mechanics, LaRC 

6/13/13 

 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

3 of 97 

NESC Request No.: TI-09-00605 

Table of Contents 

Technical Assessment Report (Phase 2) 

1.0 Notification and Authorization ........................................................................................ 7

2.0 Signature Page ................................................................................................................... 8

3.0 Team List ........................................................................................................................... 9

4.0 Executive Summary ........................................................................................................ 10

5.0 Assessment Plan .............................................................................................................. 12

6.0 Problem Description and Proposed Solution ............................................................... 14

7.0 Data Analysis ................................................................................................................... 14

7.1 AADS Phase 2 Model Improvements ................................................................... 15

7.2 Improved Simulations ........................................................................................... 32

7.3 Monte Carlo Analysis ........................................................................................... 44

7.4 Processor Load Analysis/Benchmarking .............................................................. 51

7.5 Other Trades and Sensitivity Analysis .................................................................. 54

7.6 AADS Performance Assessment .......................................................................... 59

7.7 Open Phase 2 Investigation Topics ....................................................................... 89

8.0 Findings, Observations, and NESC Recommendations .............................................. 92

8.1 Findings................................................................................................................. 92

8.2 Observations ......................................................................................................... 93

8.3 NESC Recommendations...................................................................................... 93

9.0 Alternate Viewpoint ........................................................................................................ 94

10.0 Other Deliverables .......................................................................................................... 94

11.0 Lessons Learned .............................................................................................................. 94

12.0 NASA Standards and Specifications ............................................................................. 94

13.0 Definition of Terms ......................................................................................................... 94

14.0 Acronyms List ................................................................................................................. 95

15.0 References ........................................................................................................................ 97

List of Figures

Figure 7.1-1. AADS Phase 2 Block Diagram ............................................................................. 17 

Figure 7.1-2. Venus Mission Scenario RSE #1 Actual Temperatures versus Predicted 

Temperatures for Densities between 2 and 30 kg/km3

 ......................................... 23 

Figure 7.1-3. Venus Mission Scenario Actual Temperatures versus Predicted Temperatures 

for Densities between 30 and 60 kg/km3 .............................................................. 24 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

4 of 97 

NESC Request No.: TI-09-00605 

Figure 7.1-4. Venus Mission Scenario RSE #1 Maximum Solar Panel Temperature Residual 

versus Predicted Maximum Temperature for densities from 2 to 30 kg/km3 ....... 25 

Figure 7.1-5. Venus Mission Scenario RSE #2 Maximum Solar Panel Temperature Residual 

versus Predicted Maximum Temperature for densities from 30 to 60 kg/km3 ..... 25 

Figure 7.1-6. Venus Mission Scenario Model Fit Error Distribution ......................................... 26 

Figure 7.1-7. Venus Mission Scenario Model Representation Error Distribution ..................... 27 

Figure 7.1-8. Flowchart of the AADS N-sigma Bias Maneuver Logic ...................................... 29 

Figure 7.1-9. Scaling Flight-like Atmosphere for the ORT Simulation ..................................... 31 

Figure 7.2-1. Reference and Operational Mission Run-outs used to Establish Criteria for 

Phase 2 AADS Evaluation within an ORT ........................................................... 34 

Figure 7.2-2. SP Apses Timing Performance at Mars in AAHFS without Accelerometer 

Errors..................................................................................................................... 37 

Figure 7.2-3. SP Periapsis Altitude Performance at Mars in AAHFS without Accelerometer 

Errors..................................................................................................................... 38 

Figure 7.2-4. AADS Corridor Performance at Mars in AAHFS without Accelerometer 

Errors..................................................................................................................... 38 

Figure 7.2-5. SP Apses Timing Performance at Titan in AAHFS without Accelerometer 

Errors..................................................................................................................... 39 

Figure 7.2-6. SP Periapsis Altitude Performance at Titan in AAHFS without Accelerometer 

Errors..................................................................................................................... 40 

Figure 7.2-7. AADS Corridor Performance at Titan in AAHFS without Accelerometer Errors, 

Overlaid by POST2 Results from Phase 1 ............................................................ 40 

Figure 7.2-8. SP Apses Timing Performance at Venus in AAHFS without Accelerometer 

Errors..................................................................................................................... 41 

Figure 7.2-9. SP Periapsis Altitude Performance at Venus in AAHFS without Accelerometer 

Errors..................................................................................................................... 42 

Figure 7.2-10. AADS Corridor Performance at Venus in AAHFS without Accelerometer 

Errors..................................................................................................................... 42 

Figure 7.2-11. Block Diagram for the EML Version of AADS ................................................... 44 

Figure 7.3-1. AADS Monte Carlos Results: Aerobraking Duration .......................................... 46 

Figure 7.3-2. AADS Monte Carlos Results: Final LTST ........................................................... 47 

Figure 7.3-3. AADS Monte Carlo Results: Number of Maneuvers ........................................... 47 

Figure 7.3-4. AADS Monte Carlo Results: Total ABM 'V ...................................................... 48 

Figure 7.3-5. AADS Monte Carlo Heat Rate Distribution ......................................................... 49 

Figure 7.3-6. AADS Monte Carlo Results: Heat Rate Statistics ................................................ 50 

Figure 7.3-7. AADS Monte Carlo Results: Periapsis Timing Errors ......................................... 50 

Figure 7.4-1. AADS Low-Rate Integrator State Determination Scaled Time for Mars Case 

using 21x21 Central Body Harmonic Gravity Model ........................................... 53 

Figure 7.4-2. AADS Low-Rate Integrator Orbit Propagator Scaled Time for Mars Case using 

21x21 Central Body Harmonic Gravity Model .................................................... 54 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

5 of 97 

NESC Request No.: TI-09-00605 

Figure 7.5-1. Ground Update State Uncertainty Propagated over 7-days .................................. 55 

Figure 7.5-2. Corridor Performance for Phase 2 AADS with 7-day Ground Update and 

Accelerometer Errors ............................................................................................ 56 

Figure 7.5-3. Apses Timing Performance for Phase 2 AADS with 7-day Ground Update and 

Accelerometer Errors ............................................................................................ 57 

Figure 7.5-4. Periapsis Altitude Performance for Phase 2 AADS with 7-day Ground Update 

and Accelerometer Errors ..................................................................................... 57 

Figure 7.6-1. Profiles of Observed Density (blue), Model Density with A0 =1 (solid red) and 

the Model Density Scaled to Match area under Observed Density Profile 

(dashed red) for a sample AADS Mission Orbit 209 ............................................ 62 

Figure 7.6-2. A0 Multipliers Observed and used in the Simulations ......................................... 63 

Figure 7.6-3. ORT1 Daily Analysis ............................................................................................ 65 

Figure 7.6-4. ORT1 Weekly Analysis: Orbit Period .................................................................. 67 

Figure 7.6-5. ORT1 Weekly Analysis: LTST ............................................................................ 67 

Figure 7.6-6. ORT1 Heat Rate versus Days of Aerobraking ...................................................... 68 

Figure 7.6-7. ORT2 Heat Rate Corridor ..................................................................................... 70

Figure 7.6.8. ORT2 with the 3-Sigma High Heat Rate Predictions ........................................... 71 

Figure 7.6-9. ORT2a with the Density 3-Sigma Correction ....................................................... 72 

Figure 7.6-10. Data Flow Chart for the Two Types of ORTs; EOM ........................................... 73 

Figure 7.6-11. ORT Corridors ...................................................................................................... 75 

Figure 7.6-12. ORT Glide Slope Comparison to Reference ......................................................... 76 

Figure 7.6-13. LTST of the Ascending Node Comparison to Reference ..................................... 77 

Figure 7.6-14. ORT LTST versus Orbit Period Comparison to Reference .................................. 77 

Figure 7.6-15. ODY Aerobraking Maximum Heat Rate .............................................................. 83 

Figure 7.6-16. AADS ODY Aerobraking Mission Run-out Maximum Heat Rate. ..................... 84 

Figure 7.6-17. AADS ODY Narrow Corridor Simulation Peak Heat Rates ................................ 85 

Figure 7.6-18. Mission versus AADS Glide Slope Comparison .................................................. 86 

Figure 7.6-19. Mission Compared to AADS Two Replication Simulations ................................ 86 

Figure 7.6-20. MRO Aerobraking Peak Heat Rates ..................................................................... 87 

Figure 7.6-21. AADS MRO Aerobraking Mission Run-out Maximum Heat Rate ...................... 88 

Figure 7.6-22. MRO versus AADS Glide Slope Comparison ...................................................... 88 

Figure 7.6-23. MRO versus AADS Aerobraking Mission Parameter Comparison ..................... 89 

List of Tables 

Table 7.3-1. Monte Carlo Parameters ....................................................................................... 45 

Table 7.5-1. Impact of Gravity Model Errors on SP Accuracy and Timing for 7-day Ground 

Update at Mars when Compared Against Fixed 21 degree, 21 order Gravity 

Truth Model .......................................................................................................... 58 

Table 7.6-1. Aerobraking Operations Ground-Based Analysis Teams ..................................... 60 

Table 7.6-2. ORT Performance Metrics .................................................................................... 74 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

6 of 97 

NESC Request No.: TI-09-00605 

Table 7.6-3. Issues Related to ORT Simulation and their Impacts and Resolutions ................ 79 

Table 7.6-4. Comparison of ODY Flight and AADS Replicated Missions .............................. 83 

Table 7.6-5. Comparison of MRO Flight and AADS Replicated Mission ............................... 87 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

7 of 97 

NESC Request No.: TI-09-00605 

Technical Assessment Report (Phase 2) 

1.0 Notification and Authorization 

The NASA Engineering and Safety Center (NESC) received a request from Mr. Daniel Murri, 

NASA Technical Fellow for Flight Mechanics at Langley Research Center (LaRC) to develop an 

autonomous aerobraking (AA) capability. 

The NESC received this request on December 15, 2009. An initial evaluation for all phases of 

this assessment was approved to proceed at the February 4, 2010, NESC Review Board (NRB). 

The final report for Phase 1 was approved by the NRB on December 15, 2011. The assessment 

plan for Phase 2 was approved by the NRB on December 15, 2011. 

The key stakeholders for this assessment (all phases) are the NESC (including the technical 

disciplines of Flight Mechanics; Aerosciences; Passive Thermal; Guidance, Navigation, and 

Control (GN&C); Software; Loads and Dynamics; Human Factors), and future NASA Programs 

and Projects that may utilize aerobraking. 
NASA Engineering and Safety Center

Technical Assessment Report

Document #:

NESC-RP09-00605

Version:

1.0

Title:

Autonomous Aerobraking (Phase 2)

Page #:

8 of 97

NESC Request No.: TI-09-00605

2.0 Signature Page

Submitted by: 

Mr. Daniel G. Murri Date 

Significant Contributors: 

Ms. Jill L. Prince Date Mr. Richard W. Powell Date 

Ms. Alicia M. Dwyer Cianciolo Date 

Signatories declare the findings and observations compiled 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. 

5FBN4JHOBUVSF1BHFPO'JMF
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

9 of 97 

NESC Request No.: TI-09-00605 

3.0 Team List

Name Discipline Organization 

Core Team 

Dan Murri NESC Lead, NASA Technical Fellow for Flight 

Mechanics LaRC

Richard Powell NESC Deputy Lead, Flight Mechanics TDT AMA 

Jill Prince Team Lead LaRC 

Hollis Ambrose APL Simulation APL 

Angela Bowes LaRC Simulation LaRC 

David Carrelli APL Simulation APL 

Alicia Cianciolo LaRC Simulation LaRC 

John Dec Thermal LaRC 

Allen Halsell Aerobraking Analysis/AeroNav JPL 

Chris Johansen MTSO Program Analyst LaRC 

Robert Maddock LaRC Simulation LaRC 

Dan O’Shaughnessy APL Simulation Lead APL 

Fazle Siddique APL Simulation APL 

Tom Strikwerda APL Project Manager APL 

Robert Tolson Atmospheric Modeler NCSU/NIA 

Joseph White LaRC Simulation AMA 

Tung-Han You Aerobraking Analysis/AeroNav JPL 

Consultants/Peer Reviewers 

Neil Dennehy NASA Technical Fellow for GN&C GSFC 

Starr Ginn Loads and Dynamics TDT DFRC 

Cynthia Null NASA Technical Fellow for Human Factors ARC 

Steve Rickman NASA Technical Fellow for Passive Thermal JSC 

Dave Schuster NASA Technical Fellow for Aerosciences LaRC 

Timothy Brady Systems Engineering Office JSC 

Steven Gentz NESC Chief Engineer MSFC 

Administrative Support 

Linda Burgess Planning and Control Analyst AMA 

Terri Derby Project Coordinator AMA 

Erin Moran Technical Writer AMA 
NASA Engineering and Safety Center

Technical Assessment Report

Document #:

NESC-RP09-00605

Version:

1.0

Title:

Autonomous Aerobraking (Phase 2)

Page #:

10 of 97

NESC Request No.: TI-09-00605

4.0 Executive Summary

Phase 1 of the Development of Autonomous Aerobraking (AA) Assessment investigated the 

technical capability of transferring the processes of aerobraking maneuver (ABM) decisionmaking (currently performed on the ground by an extensive workforce and communicated to the 

spacecraft via the deep space network'61) to an efficient flight software algorithm onboard

the spacecraft. To accomplish this, aerodynamic and thermal models for a representative

spacecraft were developed for the onboard algorithm known as AA development software 

(AADS) and a ground-based “truth” simulation developed for testing purposes. An autonomous

ephemeris, atmosphere estimator (AE), and maneuver estimator (ME) were also developed and 

incorporated into the AADS. Previous aerobraking mission experience [refs. 1 and 2] indicated 

that an increase in the error of the predicted time of periapsis passage requires frequent (daily) 

ephemeris updates from the ground using DSN. The Phase 1 simulation analysis of AADS 

demonstrated the algorithm could provide state estimates within the previously accepted 

250-second periapsis timing error for more than a week before requiring a ground update, thus 

eliminating the need for continuous DSN coverage. AADS was tested in simulation at three 

destinations, Mars, Venus, and Titan, and included atmospheric perturbations and sensor (inertial 

measurement unit (IMU)) measurement errors. 

This document describes Phase 2 of this study, which was a 12-month effort to improve and 

rigorously test the AADS developed in Phase 1. Model improvements included rewriting the 

Phase 1 ephemeris estimator (EE) model, which is called the state propagator (SP) in Phase 2 to

more easily accommodate flight software structures and leverage existing flight software 

architectures. The AE was modified to include the prediction of atmospheric density 

uncertainty, which was incorporated into the maneuver decision logic. The AADS simulation 

environment, the Program to Optimize Simulated Trajectories (POST2), was upgraded,

decreasing run time and increasing simulation capability. Dispersion analysis, using statistical 

Monte Carlo approaches and error uncertainties, was performed to more accurately measure the 

operability of using AA in a perturbed environment. Additionally, initial assessments of the 

spacecraft dynamic dispersions were evaluated using a 6 degree-of-freedom (6-DOF) simulation. 

Trade studies were assessed with the SP and introduced the option of using AeroNAV, an 

onboard orbit determination (OD) tool that incorporates functionality not currently modeled in 

AADS, such as fault tolerances, orbit safing approaches, and sequence timing information. 

Studies also provided first order analysis of the onboard computational resources required by the 

software. Finally, the improved AADS performance was assessed by its ability to reproduce the 

Mars Odyssey (ODY) and Reconnaissance Orbiter (MRO) aerobraking missions and its 

performance in an ODY-like aerobraking mission operational readiness test (ORT). 

The extensive Phase 2 testing verified that the AADS algorithms were accurate, robust, and 

ready to be considered for application to future missions that utilize aerobraking and has 

extended the technology readiness level (TRL) to between five and six. Additional development 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

11 of 97 

NESC Request No.: TI-09-00605 

of AA, including studying the implications of operating the AADS in a real-time environment, 

and then installing the software onto a spacecraft using AADS during flight in a shadow-mode, 

where onboard-determined commands would be validated against aerobraking decisions made by 

the ground staff would further the TRL to eight. 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

12 of 97 

NESC Request No.: TI-09-00605 

5.0 Assessment Plan

The complete development of the AA capability was planned for four phases. Phase 1, 

summarized in reference 3, included the development of the AADS and evaluated its 

performance in a simulated setting at Mars, Venus, and Titan using an MRO like spacecraft. 

Individual components of the AADS, namely the AE, EE, and ME, were developed and tested in 

both a performance simulation (e.g., POST2) and a high-fidelity flight-like simulator based on 

the MErcury Surface, Space ENvironment, GEochemistry and Ranging (MESSENGER) 

spacecraft simulation. Details of this simulation environment and its applicability to aerobraking 

are provided in the Phase 1 final report, Section 7.5.2.1. The objective of Phase 1 was to 

demonstrate the technical feasibility of the software, and the simulation analysis showed that it is 

feasible for AADS to provide AA control of a spacecraft with ephemeris updates no more than 

once per week at all three sampled destinations potentially reducing cost incurred by both DSN 

coverage and ground support. 

This report summarizes the activities of Phase 2, which explored more sophisticated schemes for 

the ephemeris and atmosphere estimation as a means of improving AADS robustness and 

performance. Improved modeling and error checking were implemented and additional offnominal stress testing of AADS was performed. 

The primary tasks planned for Phase 2 are described below. All primary tasks were 

accomplished. 

x Improve AE and EE 

o Desensitize the AE to EE timing errors and investigate modeling methods 

independent of altitude. Incorporate 3-sigma estimates of atmospheric density in 

maneuver estimator logic. 

o Reduce errors in the EE by exploring new integration methods, step sizes, and realtime periapsis timing correction. 

x Incorporate operational elements in AADS 

o Streamline code execution. 

o Add error handling and fault detection and correction. 

o Accommodate safe-mode events (safety triggers) and off-nominal scenarios, such as 

pop-up maneuvers and typical aerobraking spacecraft contingencies. 

o Assess the feasibility of incorporating collision avoidance in AADS. 

o Add flexibility to the AADS maneuver logic and implement versatile control 

strategies that can meet project specific requirements. 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

13 of 97 

NESC Request No.: TI-09-00605 

x Improve POST2 and AA High Fidelity Simulation (AAHFS) “truth” models and simulations

o Complete the integration of Venus and Titan environment models into AAHFS. 

o Incorporate Mars aerobraking flight data as a “truth" atmosphere archive.

x Incorporate uncertainties and perform Monte Carlo analyses of AADS performance 

o Identify uncertainties in “truth” and onboard modeling. 

o Perform Monte Carlo assessment of “truth” simulations.

x Stress-test AADS in POST2 and AAHFS 

o Atmospheric random noise and bias. 

o Initial ground errors. 

o Modeling errors. 

o Fault management logic. 

x Conduct trade and sensitivity studies 

o EE functionality versus AeroNav. 

o Effect of data buffer frequency on the EE orbit knowledge. 

o Accelerometer sensitivities (bias, noise, scale factor, alignment, etc.). 

o Analyze corridor width versus maneuver frequency. 

o Evaluate maneuver execution performance sensitivities. 

o Solar radiation pressure. 

x Benchmark flight-like processor limitations 

x Assess comparison of AADS decisions with mission simulations using ground-based 

operations teams 

o Run ORT simulating a stressing aerobraking environment. 

o Run ODY spacecraft in AADS using ODY atmospheric data to closely recreate an 

ODY mission. 

o Run MRO spacecraft in AADS using MRO atmospheric data to closely recreate an 

MRO mission. 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

14 of 97 

NESC Request No.: TI-09-00605 

Phase 3 would incorporate the AADS onto a flight-like processor to study the implications of 

operating the AADS in a real-time environment. This phase would quantify the computational 

resources required to support AADS in a flight mission and increase the TRL to a six. 

In Phase 4, the AADS would be installed onto a spacecraft that will use aerobraking, and then 

AA would operate during flight in a shadow-mode, where all the steps for AA are performed, but 

the commands are not executed elevating the TRL to between seven and eight. The onboarddetermined commands would be validated against aerobraking decisions made by the ground 

staff. 

6.0 Problem Description and Proposed Solution

To reduce propellant required, NASA uses aerobraking to decelerate an orbiting spacecraft into a 

desired science orbit around a target planet or moon with an appreciable atmosphere. Detailed 

descriptions of both conventional and AA are described in reference 3. The problem addressed 

herein is how to reduce the cost and risk of the conventional aerobraking mission phase by 

moving much of the decision making onboard the spacecraft. Doing so reduces the cost of the 

DSN and ground support while reducing risk by allowing precisely targeted maneuvers more 

often. Initial development of the onboard software is described in reference 3. The primary 

problem for the Phase 2 effort was to verify, through analysis, that AADS is robust and to 

demonstrate its capability in flight-like environment. 

The proposed solution includes: (1) developing a flight-like environment (i.e., using mission 

observed aerobraking pass densities) (2) upgrading the AADS algorithms to work in the flightlike environment, (3) including additional dispersions, trade studies and stress testing to verify 

that AADS can perform adequately in a wide range of simulated conditions. The description of 

how each facet of the solution was developed and used to verify the robustness and demonstrate 

the capability of AADS is described in the next section. 

7.0 Data Analysis

To demonstrate the robustness of AADS, several model improvements were made since Phase 1. 

Details of the modifications to the AEs, EEs, and MEs are described in Section 7.1. This section 

also includes a description of the updates made to the thermal model and the introduction of an 

alternative EE technique within the AeroNav architecture as an alternative AADS 

implementation. Section 7.2 summarizes the updates to the POST2 and the AAHFS simulations 

and additional analysis capabilities utilized by the simulations for Phase 2. 

Section 7.3 summarizes the results of the Monte Carlo analysis while Section 7.4 describes the 

results of processor loads testing and benchmarking. Additional trades to evaluate the robustness 

of the AADS due to unique parameters are described in Section 7.5. Finally, the overall AADS 

performance demonstration in a flight-like environment is described and results are presented in 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

15 of 97 

NESC Request No.: TI-09-00605 

Section 7.6. It is based on these analyses that the recommended next step for AADS is 

implementation in a flight-like processor to study the implications of operating in a real-time 

environment. 

Section 7.7 lists aspects of the AADS development that were not considered as part of Phase 2 

and describe what remains as future work to prepare for Phase 3. 

7.1 AADS Phase 2 Model Improvements 

7.1.1 Atmosphere Estimator 

During the implementation of the Phase 1 AE in AADS, it was found that the AE statistical 

model was providing uncertainties of the estimated density and scale height that were much 

smaller than the natural variability seen during previous aerobraking missions. To improve the 

uncertainty estimation, AE was redesigned to use deviation between the predicted and observed 

values from the number (N) orbits prior to the one being predicted (i.e., orbit N+1). Previous 

mission data is not used for prediction as scaled season, altitude, time, and location on the planet 

does not translate well for accurate prediction. A number of averaging methods were tested 

against previous aerobraking mission data sets. Methods included arithmetic mean, geometric 

mean, median and harmonic mean. The goal was to find a method that permitted an unusually 

large change in density to increase the uncertainty, but not to influence the uncertainty over too 

long a period of time. For example, if the algorithm averages over the previous seven orbits and 

then the simulation experienced an orbit with an unusually high density then the method would 

provide a high mean and standard deviation for the next six predictions. This method is intended 

to protect against situations that occurred during the ODY aerobraking mission on orbit 106 

when the observed heat rate was near the immediate action line after several orbits with 

considerably lower heat rates. 

The second study included selecting an N that is large enough to provide a reasonable set of 

statistics (mean and standard deviation), but was not so large as to not recognize a change in the 

variability (e.g., periapsis precessing over the polar vortex). It became clear that different 

approaches provided the best estimates depending on the latitude of periapsis, but the differences 

were smaller than the predicted variability and not statistically significant. To maintain 

simplicity, the arithmetic mean and the standard deviation from the mean averaged over seven 

orbits (N=7) was selected. This approach was tested during the ORT and performed as well as 

ground-based human operations teams. 

7.1.2 EE using the SP 

During Phase 2, the EE algorithm was renamed the SP to more clearly convey its functionality, 

and was completely redesigned to address several shortcomings of the original Phase 1 design. 

While the results obtained in Phase 1 served as an adequate proof-of-concept, and the EE 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

16 of 97 

NESC Request No.: TI-09-00605 

demonstrated suitable performance in Phase 1 testing in terms of orbital integration accuracy and 

execution speed, the software architecture and implementation did not allow ready conversion to 

a flight environment (i.e., real-time processing). Furthermore, a review of the coding practices 

and algorithms used for the EE showed that performance improvements (orbital integration 

accuracy, software execution speed, software memory footprint) were possible with a change in 

software architecture and modifications to the orbital integration algorithms. 

A major change was made to the architecture of the EE algorithm used in Phase 1 to better 

facilitate use in a real-time environment. Running AADS on a flight processor was not a 

requirement for Phase 2, but will be necessary in subsequent phases of the project. As such, the 

software architecture modifications were made to the software during Phase 2 in an effort to save 

the substantial expense that would come from an architecture change in a later project phase. In 

Phase 1, an external process, presumed to exist in the host spacecraft flight software, would 

buffer measurement data for batch processing by AADS. In Phase 2, the SP was designed to run 

synchronously with the IMU/star tracker output data rate, thereby minimizing the software 

necessary to use AADS on a host spacecraft. Rather than buffering this measurement data, the 

host spacecraft need only call AADS at a fixed rate (i.e., 10-Hz). Because each spacecraft may 

have a different IMU and/or star tracker, these rates are user-selectable, ensuring maximum 

flexibility of the SP software. The architecture used for AADS in Phase 2 is shown in 

Figure 7.1-1. In the figure, t, a, q, h and x represent time, acceleration, quaternions, altitude, and 

state data respectively. The capital letters denote buffered data. 

The primary function of the SP is to provide predictions of the 3-DOFspacecraft position and 

velocity for use in the AE and ME. This requires a high-fidelity integration of the spacecraft 

translational equations of motion. Ground operations provide the initial condition for the 

integration (termed a "ground update"), and the spacecraft uses this initial state coupled with 

onboard force models and measurement data to propagate the ground update forward in time. 

Because the desired integration interval is long (defined by this project to be no more often than 

once every seven days), the integration must be performed as accurately as possible because the 

position errors grow quadratically with time. 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

17 of 97 

NESC Request No.: TI-09-00605 

Figure 7.1-1. AADS Phase 2 Block Diagram 

The SP is divided into two functions: (1) state determination (SD) or determining the current 

spacecraft translational state, and (2) orbit propagation (OP) or propagating of the current state 

forward in time to predict a future orbital condition (periapsis or apoapsis). The SD process 

integrates the last known state forward in time using the IMU measurements and onboard gravity 

models to propagate the state up to the current time. Once on each orbit, the OP process takes 

the current state and predicts ahead one orbit to provide estimates of the subsequent periapsis and 

apoapsis states to the AE and ME tasks for corridor maneuver computations. The SD runs in 

real-time, although at a low rate (~0.01-Hz) as it continually updates the translational state based 

on the sensed accelerations. The OP is a batch process that only runs once per orbit to feed the 

AE and ME algorithms. 

To accurately perform the SD, all accelerations acting on the vehicle must be included in the 

integration. These accelerations can be broadly grouped into two categories: gravitational and 

non-gravitational. In general, the gravitational accelerations are modeled in the software, and the 

non-gravitational accelerations are measured with the accelerometers. The dominant 

gravitational acceleration is due to the central body, but other perturbing bodies may also be 

modeled. The dynamics of the motion due to gravity change slowly over time, allowing for a 

large integration time step. Further, the gravitational accelerations are computationally 

expensive, so there is a strong desire to limit the number of evaluations of these models. In 

contrast, the non-gravitational accelerations, which can change quickly with time, particularly 

aerodynamic drag and thrust acceleration, necessitate a small integration step. The competing 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

18 of 97 

NESC Request No.: TI-09-00605 

step size requirements for the orbital integration create what is known as a computationally stiff 

system. 

The EE solved this stiff system by ignoring the evaluation of the non-gravitational accelerations 

during the exo-atmospheric portion of the orbit and integrating the equations of motion with a 

large, variable time step (tens or hundreds of seconds). During those times when the nongravitational accelerations were substantial (during the drag pass and ∆V maneuvers), the 

integrator ran with a small, fixed integration step (0.1 second). While this method is reasonably 

accurate, the high-frequency evaluation of the gravity model(s) during the drag pass and burns is 

unnecessary and significantly increases computational cost. 

The solution to the stiff orbital integration problem used by the SP is to separate the fast and 

slow dynamics, and integrate them with different methods and time steps. This hybrid method 

capitalizes on the special structure of the integration problem, y’’ = f(t,y) + g(t). The slow 

dynamics, f(t,y), are due to the gravitational accelerations and are evaluated using a variable-step 

Runge-Kutta-Nystrom (RKN) method that allows minimal computation of the expensive 

harmonic gravity models. The fast dynamics of the non-gravitational terms, g(t), are not explicit 

functions of the state so their contribution to the integrated state can be computed separately and 

included in the full state integration only when required by the RKN integrator. This high-rate 

integration of the non-gravitational accelerations is accomplished with a simple fixed-step 

trapezoidal integration of the input accelerometer data at (or close to) the measurement data rate. 

This hybrid (fast/slow) integration used in the SP allows rapid fixed-step integration of the fast 

dynamics, and the variable-step RKN used to integrate the slow dynamics takes steps only as 

required to maintain the specified integration tolerance. This serves to minimize the 

computational burden without sacrificing accuracy in the spacecraft SP. 

The SP architecture described above and shown in Figure 7.1-1 greatly reduced the AADS 

memory requirements and computational burden required to accurately integrate the orbit. The 

largest computational burden in the EE or SP is the harmonic gravity model evaluations. The SP 

only evaluates these models as necessary (throughout the entire orbit) to maintain tolerance, 

whereas the EE computes these models dozens of times each second during the drag pass and 

burns. Depending on the time step of the variable-step integration and the duration of the drag 

pass, the improved structure of the SP can cut the number of gravity model evaluations by two to 

ten times over the number required by the EE, greatly speeding the execution of the orbital 

integration. By integrating the high-rate acceleration data in real-time with a simple integration 

method, the SP eliminates the need to buffer this data throughout the drag pass. A much smaller 

buffer (~100 seconds) of high-rate ∆V data is all that is necessary to perform the low-rate 

integration in the SP versus the large (~4000 seconds) of high-rate accelerometer and quaternion 

data that was required to execute the EE. This architecture change reduces the memory 

requirements of AADS by a factor of nearly 100. 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

19 of 97 

NESC Request No.: TI-09-00605 

7.1.3 EE: AeroNav 

AeroNav is a stripped down version of the Jet Propulsion Laboratory (JPL) Navigation AutoNav 

program. It contains the trajectory integration module with some specific spacecraft dynamic 

modeling capabilities for autonomous on-board aerobraking/aerosampling applications. 

AeroNav consists of a RG78 integrator and key spacecraft dynamic modeling functions. It can 

run onboard the spacecraft and a ground-based simulator. The full AutoNav program has been 

flown on several spacecraft, including Deep Space 1, Deep Impact, and Stardust nominal and 

extended missions. 

The overall plan for JPL participation in the AA task was to: (1) develop a stand-alone EE for 

use with AADS and POST2 to serve as a validation point, and (2) feed AADS functions, such as 

the AE and ME back into the current AutoNav software to produce a functioning onboard system 

for AA. This would combine the current capabilities for virtual machine language (VML)-based 

spacecraft commanding of turns and propulsive maneuvers with knowledge of the atmosphere. 

The final step would be operation of this new AeroNav in testbed mode, with atmospheric 

modeling based on past operations. 

These two tasks proceeded in parallel until the complications of simultaneously modifying and 

stripping out the EE/SP required more attention. The AutoNav-based EE has been developed 

and showed good agreement with AADS test cases. The additions to the estimator are described 

below. To date, the updates to the eventual full AeroNav have been to upgrade to VML 3.0 and 

to setup a testbed for development. 

In future development, incorporating the AE and ME into AutoNav to produce a flight-capable 

system should be a priority. The lessons learned in Phases 1 and 2 will be used to develop logic 

algorithms capable of dealing with aerobraking-specific operations, such as corridor control and 

automated pop-up maneuvers. Running in testbed mode would provide a full verification of 

AADS. 

During AADS Phase 2, several modifications to AutoNav were made to reproduce the EE 

function. These include low-thrust ion propulsion, executive functions to control the spacecraft 

(e.g., the commanding of attitude changes or propulsive maneuvers), low-level message-passing 

functions, optical image processing, orbit-determination filtering, and maneuver-determination 

functions. The AeroNav delivery retains trajectory integration and the associated spacecraft 

dynamic modules, drag pass determination and event finding. 

High-level function calls are used to issue commands to AeroNav and to return results. Each 

high-level service is expressed by a C-language function. Each function may have input 

arguments, and may provide output arguments. Most return a code indicating either success or 

non success. 

AeroNav issues file interface commands to store or retrieve intermediate results in a Unix-like 

file system (both binary and text files). This has been implemented in order to provide flexibility 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

20 of 97 

NESC Request No.: TI-09-00605 

with multiple languages and environments. It also keeps a number of controlling parameters in 

text files. 

The output spacecraft ephemeris file contains ephemeris polynomials for the spacecraft upon 

which AeroNav is running. It can be regenerated after each spacecraft event, such as a burn or a 

drag pass. It can also be re-regenerated after each maneuver computation, to incorporate the 

upcoming maneuver just prepared. This and other ephemeris files are stored in files of 

Chebyshev polynomial coefficients. Positions and velocities can be evaluated from these 

polynomials instead of integrating a trajectory numerically. These binary files have provisions 

for storing ephemeris coefficients for more than one body. Also, an ephemeris can be segmented 

time-wise, allowing lower-degree polynomials to be used for different arcs. 

The arc for each body contains three sets of polynomials coefficients, for the X-, Y-, Z-position 

components relative to a specified origin, and with respect to the International Celestial 

Reference Frame (ICRF)1

. Polynomials for velocity coefficients are not stored, as they can be 

derived from the position coefficients. 

A number of files in this format are used to provide necessary inputs for the integration. The 

planetary ephemeris file contains ephemeris polynomials for solar system bodies, such as the 

Sun, the planets, and their natural satellites. A convention allows storing an ephemeris of the 

barycenter of a planet/satellite system, as well as ephemerides of the center of each body. 

The body parameter file specifies integration center body information, which includes gravity 

parameters, radius, gravitational constants, and body rotation parameters. It is a text-based input 

file and must be loaded in the system before calling the ephemeris creation function. 

The binary maneuver file contains one record for every maneuver. Each record contains the 

ET2000 time of the maneuver (from midnight January 1, 2000), and the 'V components of any 

maneuver scheduled during the planning cycle. These components are expressed in 

kilometers/second in the ICRF coordinate system, with the Sun as the origin. The maneuver file 

should be updated as a maneuver is computed or designed. Note that a finite-burn model is 

implemented in the latest version, but only an impulsive model is described for this task. 

The non-gravitational history file is a binary file describing the non-gravitational forces on the 

spacecraft (as might be inferred by its name). Each record contains the ET2000 time of the 'V, 

and a three-element Cartesian vector of the 'V components. The vector is expressed in 

meters/second in the ICRF coordinate frame (not the spacecraft body frame). Typically, 

accelerometers keep track of thruster firings during deliberate maneuvers, desaturations of the 

momentum wheels, or drag passes. These are buffered by calls to nongrav_gen.c, which 

converts high-frequency IMU acceleration data into a lower-frequency non-grav history file. 

 

1

 International Celestial Reference Frame; see http://www.iers.org/ (search for icrf) 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

21 of 97 

NESC Request No.: TI-09-00605 

The user inputs the IMU data, and the desired time step of the output data. The acceleration data 

are then integrated over each time step and the resultant velocities at the center of each time step 

are calculated and output into the non-grav file at each time. 

A binary OD file contains the ET2000 time and the spacecraft state (position and velocity). This 

file is used only to store the uploaded OD solution, as AeroNav performs no OD. 

Finally, the controls file is intended to hold parameters for all of AeroNav with exception of 

central body oblateness parameters. Parameters related to the generation of a new spacecraft 

ephemeris, such as code numbers for the planets/barycenters to include gravitational 

perturbations, the masses of each planet, and the spacecraft mass and area (only used for solarradiation pressure calculations) are part of the control file. Also, if desired, there is an overall 

scale factor for the non-grav history 'V. An additional parameter specifies the maximum 

number of non-grav history records to retain. Filenames and directory paths are included in the 

control file. 

A small number of low-level utilities are used to calculate trajectory events; ‘dragpass.c’ finds 

the start and end times of the drag pass. The input table of IMU data is searched for the time 

period during which the acceleration is above a given threshold. The desired threshold is input 

as a fraction of the peak acceleration, and the duration of the drag pass in seconds is returned. 

The ‘find_apses.c’ searches an input spacecraft Chebyshev ephemeris file for all apses occurring 

within a specified time period for a specified spacecraft. The Chebyshev polynomials are 

converted to Cartesian states, allowing RxV to be calculated at each time. This method returns 

the periapsis or apoapsis, but not the maximum or minimum altitude, consistent with AADS 

convention. The outputs are the time, radius, and type (apoapsis or periapsis) of each apsis. 

7.1.4 Thermal Model Update 

The Phase 2 report focuses on AADS algorithm robustness testing and operational performance 

primarily at Mars due to the availability of aerobraking flight data at the planet. However, at the 

end of Phase 1, an error was identified in the Venus thermal model [ref. 3]. Below is a 

description of the approach to and results of the updated Venus thermal model. 

Because of its proximity to the Sun, the temperatures due to solar heating at Venus are nearly 

4.5 times higher as compared to Mars. Likewise, an aerobraking spacecraft will encounter a 

higher aerodynamic heating environment. During Phase 1, a sample spacecraft, MRO, along 

with associated thermal models, was used for the AADS aerobraking assessment at Venus. The 

thermal model nor the spacecraft were designed to handle the thermal loads at Venus, therefore, 

a different approach was used in Phase 2. 

The original approach utilized a single response surface equation (RSE) for the large day-night 

temperature (and also density) range at Venus. In Phase 2 it seemed practical to break the 

density range into separate RSE equations. The first RSE (RSE #1) would cover densities 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

22 of 97 

NESC Request No.: TI-09-00605 

between 2 and 30 kg/km3 and the RSE #2 between densities of 30 and 60 kg/km3. The original 

approach used only a single model, Thermal Desktop®, to develop the RSE. Temperatures were 

calculated using the Systems Improved Numerical Differencing Analyzer/Fluid Integrator 

software, which can be invoked from Thermal Desktop®. The new approach combined two 

models and used the temperatures from Thermal Desktop® in MSC Software PatranTM, a method 

that has heritage from the thermal model development demonstrated on both the ODY and MRO 

aerobraking missions. Results of the new approach are provided below. 

7.1.4.3 Response Surface Goodness of Fit Determination 

A face-centered central composite design (CCD) with 13 factors was generated using the JMP®

statistical software. The CCD had 26 axial points, 10 center points, and 128 points from the 

fractional factorial contribution. JMP® automatically reduces the fraction used to compute the 

fractional factorial contribution as the number of factors increases; in this case the fraction was 

1/64th. The temperatures calculated for each of the 164 total runs for both Mars and Venus were 

analyzed using JMP® where a least-squares-fit was constructed using the stepwise regression 

option in JMP®. The result of the regression is a quadratic equation, one unique to the Mars 

mission scenario and one unique to the Venus mission scenario. The coefficient of 

determination or R2 adjusted value was measured and used to determine how well the assumed 

functional form of the response measures the variability of the supplied data. In this case, the R2

adjusted value measured how well the quadratic response surface represented the variability in 

the temperatures generated by the design of experiment (DOE) cases. For the Venus mission 

scenario, the updated R2

 adjusted value for RSE #1 is 0.9952 and for RSE #2 is 0.9993. An R2

adjusted value greater than 0.9 was desirable, but was not sufficient to determine the goodness of 

fit of the response surface. 

To understand how well the RSE is fitting the response data from the DOE runs, a plot of the 

actual versus predicted values, a plot of the residual versus predicted values, and the model fit 

distributions must be examined. The actual versus predicted plot shows the temperatures 

calculated by the thermal model for the cases described in the DOE plotted against the 

temperatures calculated by the quadratic RSE and is given in Figure 7.1-2 for the RSE #1 and 

Figure 7.1-3 for RSE #2. In the figures and throughout the thermal results section of this report, 

the term “actual” temperatures refers to the temperatures calculated by Patran Thermal® and the 

“predicted” temperatures are those computed by the RSE. Note that the temperatures for the 

Venus mission scenario, again based on a Mars spacecraft, are unrealistically high and are 

therefore scaled to match the maximum temperature calculated for a proposed Venus 

aerobraking spacecraft; a spacecraft which had a more robust thermal design and had solar 

panels tailored to minimize the aerodynamic heating. 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

23 of 97 

NESC Request No.: TI-09-00605 

Figure 7.1-2. Venus Mission Scenario RSE #1 Actual Temperatures versus Predicted Temperatures for 

Densities between 2 and 30 kg/km3 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

24 of 97 

NESC Request No.: TI-09-00605 

Figure 7.1-3. Venus Mission Scenario Actual Temperatures versus Predicted Temperatures for 

Densities between 30 and 60 kg/km3

The solid centerline of the plot in Figure 7.1-2 and Figure 7.1-3 represents a perfect fit of the 

data. The plots show that the data points lie close to the centerline, which indicates a good fit. 

The residual is the error between the temperature calculated by the thermal model and the 

temperature calculated by the RSE. The residual for the maximum solar panel temperature 

versus the predicted maximum temperature is plotted in Figure 7.1-4 for the RSE #1 and 

Figure 7.1-5 for RSE #2. 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

25 of 97 

NESC Request No.: TI-09-00605 

Figure 7.1-4. Venus Mission Scenario RSE #1 Maximum Solar Panel Temperature Residual versus 

Predicted Maximum Temperature for Densities from 2 to 30 kg/km3

Figure 7.1-5. Venus Mission Scenario RSE #2 Maximum Solar Panel Temperature Residual versus 

Predicted Maximum Temperature for Densities from 30 to 60 kg/km3
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

26 of 97 

NESC Request No.: TI-09-00605 

In general, the data points are randomly scattered in Figures 7.1-4 and 7.1-5, which indicates a 

good fit of the temperature data. 

One final check of the goodness of fit is to examine the model fit and model representation error 

distributions. Both are calculated as percentages based on absolute temperatures in Kelvin. The 

model error distributions should approximate a normal distribution with a mean around zero and 

standard deviation less ≤ 1.0. The model fit error is a measure of how well the response surface 

fits the temperature data in the DOE. The model fit error distribution for the maximum 

temperature for the RSE #1 and RSE #2 is plotted in Figure 7.1-6. The model fit error 

distribution for the first Venus RSE is approximately normal and has a mean of -0.0251 and a 

standard deviation of 1.4326. The model fit error distribution for the second Venus RSE is 

approximately normal and has a mean of -0.00054 and a standard deviation of 0.3278. The 

standard deviation for RSE #1 is higher than desirable and may indicate that the density range 

between 2 and 30 kg/km3

 might be further divided into separate RSEs in future analysis. 

Figure 7.1-6. Venus Mission Scenario Model Fit Error Distribution 

The model representation error is how well the response surface fits temperatures calculated by 

the thermal model as percentage error based on difference in temperature (K) for points other 

than those on the DOE. The updated model representation error for RSE #1 and #2 for the 

Venus mission scenario is plotted in Figure 7.1-7. The model representation error distribution 

for the RSE #1 is approximately normal with a mean of -0.6723 percent and standard deviation 

of 2.2192 percent. For RSE #2 the model representation error distribution is approximately 

normal with a mean of -0.0022 percent and standard deviation of 0.4301 percent. Again, RSE #2 

appears to have a better fit than RSE #1. 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

27 of 97 

NESC Request No.: TI-09-00605 

Figure 7.1-7. Venus Mission Scenario Model Representation Error Distribution 

The model fit and model representation errors are accounted for in the RSE when the 

temperature calculation is made from within the AA simulation to provide a conservative 

temperature. Another error is also added as a bias to the temperature calculated by the response 

surface. This error is present because the high-fidelity thermal model will typically not be 

correlated to the aerobraking flight temperature data. This error is typically unknown until the 

first couple of drag passes are made and the flight temperatures and predicted temperatures 

compared. Therefore, a short calibration period is required, but this can be accomplished during 

the walk-in phase of aerobraking, which makes up the first initial orbits where the spacecraft 

periapsis is gradually lowered into the aerobraking altitude corridor. 

One important aspect of response surface modeling that must be emphasized is the RSEs are 

only valid over the range for which they are defined. It must be stressed that even a small 

amount of extrapolation in any factor included in the equation can produce results that are 

invalid. 

The Phase 2 analysis indicated thermal model fits could be improved by moving from a single to 

a double RSE approach. Depending on the specific Venus mission, it may be determined that an 

aerobraking mission at Venus actually requires several different RSE’s to accurately 

approximate temperatures, however, further analysis and sensitivity studies are needed. 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

28 of 97 

NESC Request No.: TI-09-00605 

7.1.5 Maneuver Estimator 

For the Phase 2 analyses, the ME was updated to utilize the 1-sigma uncertainty on density now 

provided by the AE (see Section 7.7.1). When the ME predicts the spacecraft's next atmospheric 

pass maximum heat rate relative to the operational corridor, it also calculates the N-sigma high 

(or low, in the case of the altitude corridor) estimate, where N is specified by an AADS iLoad 

parameter. This additional dataset is output for informational purposes, but can also be used to 

modify the AADS ME maneuver logic. If instructed to do so (through the setting of an 

additional AADS iLoad parameter), the ME can compare the N-sigma estimate against an 

"immediate action line" (also input through an AADS iLoad), and if exceeded, the ME passes 

into its "bias maneuver" logic. The intent of this capability is to further protect the spacecraft 

from any instances where the AE density uncertainty is sufficiently high as to cause concern that 

an immediate action line violation has an increased likelihood of occurring on the next 

atmospheric pass. In this case, the ME will command an altitude raise maneuver such that at 

least the current (previous pass) altitude is maintained. The logic is described below and shown 

in the flowchart in Figure 7.1-8. 

x If the nominal AADS ME logic (looking at only the predicted nominal location with 

respect to the operational corridor) has indicated an altitude raise maneuver is necessary, 

where the increase in periapsis altitude places the next periapsis altitude at or above that 

of the previous pass, that maneuver is commanded; if the altitude raise is insufficient to 

maintain the previous pass periapsis altitude, the maneuver is adjusted such that the 

current periapsis altitude is maintained. 

x If the nominal AADS ME logic has indicated an altitude lowering maneuver is necessary, 

this maneuver is ignored, and instead the ME commands only that maneuver which may 

be necessary to maintain the periapsis altitude estimated from the previous pass; if the 

next periapsis altitude is estimated to be at or above that of the previous pass without a 

maneuver, no maneuver is commanded. 

x If the nominal AADS ME logic has indicated no maneuver is necessary, the ME 

commands only that maneuver which may be necessary to maintain the periapsis altitude 

estimated from the previous pass; if the next periapsis altitude is estimated to be at or 

above that of the previous pass without a maneuver, no maneuver is commanded. 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

29 of 97 

NESC Request No.: TI-09-00605 

Figure 7.1-8. Flowchart of the AADS N-sigma Bias Maneuver Logic 

Pop-up Capability

In addition to the bias maneuver logic, the ME was updated with "pop-up maneuver" logic. In 

this case, if instructed to do so (through the setting of an additional AADS iLoad), the ME can 

compare an estimate of the spacecraft's previous pass with respect to the operational corridor 

against the specified immediate action line, and if the previous pass violated this operational 

limit, the ME would immediately command a maneuver to raise the spacecraft altitude by a 

specified (through the setting of an additional AADS iLoad) number of atmospheric scale 

heights, and AADS operations would cease. In the current implementation, the simulation would 

simply terminate after the pop-up maneuver is executed, however, operationally, the AADS 

calculations could be allowed to continue, though any AADS maneuver commands would be 

ignored by the spacecraft. The intent of this capability is to protect the spacecraft in the event 

that a likely corridor violation on the previous pass was not predicted during the previous AADS 

call. The error in the prediction could have resulted from errors in the AADS execution 

(e.g., AE density predictions and/or SP state predictions) during the previous call, or simply large 

environmental variations, which were not anticipated. In either case, the spacecraft would be 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

30 of 97 

NESC Request No.: TI-09-00605 

placed into a "safe" orbit and await instructions from the ground before returning to nominal 

AADS operation. 

7.1.6 Atmosphere Modeling 

An additional atmosphere modeling update was made for Phase 2 that did not directly affect the 

AADS code, but allowed its evaluation in a realistic atmosphere, in contrast to the nominal 

Gaussian density curves generated using Global Reference Atmosphere Models (GRAMs). To 

evaluate AADS in a flight-like environment, aerobraking atmospheric flight data from Mars 

Global Surveyor [ref. 6], ODY and MRO were obtained from the planetary data system (PDS). 

Because the AADS flight trajectory in the ORT simulations would not coincide with the altitude, 

latitudes, longitudes, season and LTST, nor have the same flight time as the past flight missions, 

a method to scale the flight profiles in both time and altitude was developed. 

The goal of the scaling process is to create a POST2 data table that is a function of altitude and 

density, based on a normalized time from periapsis, for each of the flight missions aerobraking 

passes. Then, by indexing the profiles using an input parameter, the simulation can select from 

any mission’s density profile. Therefore, atmosphere profiles from different missions could be 

used in a single simulation. Likewise, profiles could be used in any order. The objective would 

be to determine if AADS could successfully accommodate the wide range of observed density 

structure. 

The process to scale the profiles required flight time from periapsis and the corresponding timedependent altitude and density from the PDS atmospheres data. Atmospheric scale height 

(e.g., the change in altitude that results in a factor of e change in density), calculated from a fit to 

periapsis data within 10 to 15 km of periapsis altitude during operations for each profile, is also 

used in the scaling and is not time dependent. 

Flight data is tagged as time from periapsis. Because the flight data time from periapsis is not 

the same for the inbound and outbound leg of the trajectory, the first step in the scaling process is 

to normalize time to range from -1 to 1. This is done by taking the shortest leg (either inbound 

or outbound) of the trajectory and by dividing all times by the lowest absolute value. For 

example, if a trajectory has a 200 seconds inbound leg and a 150 seconds outbound leg, so that 

the trajectory time stamps extend from -200 to 150 seconds, all times are divided by 

150 seconds, yielding a range of data from 1.33 to 1. Any points outside the range will be 

ignored in the simulation. Based on the normalized time values, multi-variant POST2 lookup 

tables based on orbit number and normalized time from periapsis are written for both density and 

altitude. Scale height tables are also written as a function of orbit number only. 

In the simulation an altitude trigger is specified (e.g., 200 km) as the point at which an 

atmosphere is activated. This is a valid assumption for both long and short period orbits, as there 

is little appreciable atmosphere above 150 km. When the vehicle is at or below 200 km altitude, 

the flight atmosphere lookup tables are enabled. Otherwise it is assumed there is no atmosphere. 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

31 of 97 

NESC Request No.: TI-09-00605 

When the atmosphere is enabled, the time to periapsis and orbit number are used to lookup 

reference density (ߩ (௢and altitude (݄௢) values from the tables and scale height (hs). These 

values are used, along with current altitude (݄), to calculate current density (ߩ (in an exponential 

atmosphere model: 

To verify that the actual flight density profiles are implemented in the simulation, the profiles are 

compared to POST2 profiles generated by the engineering atmosphere model, MarsGRAM. 

Nominal MarsGRAM aerobraking density passes resemble bell shaped curves indicative of 

density changing with altitude during an aerobraking pass. Figure 7.1-9 compares orbit densities 

from MarsGRAM (top) to actual ODY flight data, and ODY flight data scaled to a change in 

periapsis altitude of -590 m, +1143 m, and +560 m in sample orbits 1, 2, and 3, respectively. 

The scaling maintains observed density profile characteristics that are intended to stress AADS 

while accommodating the variations in time and altitude of the ORT simulations. The flight 

profiles for each mission, MGS and MRO, are scaled in the same manner. 

In Phase 2, all three mission data sets were used to evaluate the performance of AADS in 

realistic flight-like conditions. Results of the evaluation are provided in Section 7.6.1. 

Figure 7.1-9. Scaling Flight-like Atmosphere for the ORT Simulation 

Orb 7 Orb 8 Orb 9 0

2

4

6

8 x 10-8

Density kg/m3

Mars GRAM

ODY 1 ODY 2 ODY 3 0

2

4

6

8 x 10-8

Density kg/m3

Flight Atmospheres

Scaled Density

Table Density
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

32 of 97 

NESC Request No.: TI-09-00605 

7.2 Improved Simulations 

Two simulations were used to evaluate AADS during Phase 1: POST2 and the AAHFS. Full 

descriptions can be found in the Phase 1 report [ref. 3]. The same simulations were used for 

Phase 2, though several improvements were made. This section describes the simulation 

improvements that were made and why and how each simulation was used to evaluate the Phase 

2 AADS. 

7.2.1 POST2 

During Phase 2, the POST2 simulation code was updated from version 1.1.8, which was a 

combined C/Fortran version, to an all-C version. The C version of the performance simulation 

allows for faster run times than Fortran and allows greater flexibility, efficiency, and capability. 

Additional upgrades included adding a Monte Carlo capability and incorporating flight observed 

atmosphere profiles. These upgrades enabled faster compile and run times and afforded the 

opportunity to architect the code in such a way that minimized the differences between the 

AADS and ground-based simulations into a single executable. This greatly simplified efforts to 

evaluate AADS for the ORT. 

Once all upgrades were implemented and verified, the POST2 simulation was used to complete 

aerobraking mission simulations. However, prior to describing the mission simulations, key 

aerobraking terms are defined below. 

Maneuver Menu: During ODY and MRO aerobraking operations, a preset menu of maneuver 

sizes was stored onboard the spacecraft to minimize the chance for error when commanding an 

ABM. The same strategy was used for the Phase 2 AADS ORT analysis. The maneuver menu 

values are based on those used for ODY and MRO. They are 0.05, 0.1, 0.2, 0.3, 0.5 and 0.6 m/s. 

The sign of the magnitude was also selected providing capability to both raise and lower the 

periapsis altitude to generate lower or higher heat rates, respectively. 

Corridor: The corridor is defined as the upper and lower limits on the value of the aerobraking 

parameter. For all AA Mars simulations, the upper limit is maximum heat rate on an aerobraking 

drag pass. Venus used maximum solar array temperature and Titan used minimum altitude. The 

lower limit serves to keep the mission in the atmosphere to finish in the specified amount of time 

while the upper corridor provides desired or required margin to thermal limits of the spacecraft. 

These values can and do change throughout an aerobraking mission depending on many 

parameters including, but not limited to, the location of periapsis, the position relative to the 

glide slope, the desire to maintain spacecraft safety for a specified duration, etc. The corridor 

values are input parameters set within the iLoads data set for AADS, and they are also set as 

input parameters to all ground-based aerobraking simulations. 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

33 of 97 

NESC Request No.: TI-09-00605 

Immediate Action Line: The immediate action line is intended as a fail-safe option. The value 

of the line is set at a desired margin from the upper corridor. If the spacecraft experiences an 

unforeseen environment, such as a sudden increase in density that affects peak heat rate, the 

spacecraft commands a maneuver calculated by AADS at the next apoapsis to raise the periapsis 

altitude (lowering the heat rate) and protect the spacecraft from further harm. 

Aerobraking Margin: Aerobraking Margin: The dominant unknown for aerobraking is the 

atmospheric density profile. Density is a primary factor for spacecraft survival, as it is a main 

factor for both structural loads (dynamic pressure) and thermal (heat rate). The ratio of the 

density required to reach the immediate action line and the upper corridor is defined as margin. 

This density ratio for the Mars aerobraking missions was converted to either equivalent heat rate 

margin (if thermal is the designing issue – e.g., ODY and MRO), or dynamic pressure 

(e.g., MGS) for reporting purposes. Note that it is expected that the spacecraft will exceed the 

upper corridor on many orbits due to the atmospheric uncertainty, but should not exceed the 

immediate action line. For ODY, the aerobraking margin was 100 percent and was considered 

an aggressive mission; MRO maintained over 200 percent margin. 

During Phase 2, four different types of simulations were used to characterize and evaluate AADS 

performance and robustness. The application of each of these simulations as they pertain to the 

AADS assessment is explained in more detail in Section 7.6. Definitions are provided below: 

Reference Mission Run-out: An aerobraking mission simulation from walk-in where the 

spacecraft makes small-predetermined maneuvers to slowly dip periapsis altitudes, within the 

atmosphere, to a predefined final apoapsis altitude. The reference spacecraft flies through a 

GRAM atmosphere on each aerobraking pass and a maneuver is allowed at each orbit apoapsis. 

For the reference mission run-out, each maneuver is targeted such that the aerobraking controlled 

parameter, in this case, the maximum periapsis heat rate, is targeted to a specified value on every 

aerobraking pass. For the purposes of this study, the target heat rate value of 0.1125 W/cm2

 was 

selected such that the final apoapsis altitude of 2800 km was achieved after approximately 

70 days of aerobraking. This mission simulation provides the reference glide slope, or an 

apoapsis vs. time curve, that is used as a target value by the other simulations that use AADS. A 

reference mission run-out is a standard practice for all aerobraking missions and provides a 

guideline (glideslope) by which real-time analyses can be compared. 

Operational Mission Run-out: A mission run-out is a simulation of a spacecraft through the 

aerobraking phase that has applied realistic constraints as traditional aerobraking ground 

simulations. Historically, during ground-based aerobraking operations, maneuvers at apoapsis 

are limited to once during a specified time increment (e.g., every 48 or 72 hours) to minimize 

staff workload. The operational mission run-out simulation (similar to the reference mission runout simulation in that it flies the reference spacecraft through MarsGRAM atmospheres from 

walk-in through a final apoapsis altitude of 400 km) only allows a maneuver on specified 

apoapses and imposes limits on the aerobraking parameter (e.g., heat rate). The heat rate limits, 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

34 of 97 

NESC Request No.: TI-09-00605 

also called the corridor, have two main purposes: (1) to maintain spacecraft safety and (2) to 

control the length of the mission. Further explanation of the corridor is provided later in the 

section. At an apoapsis designated for a maneuver, the simulation predicts ahead (propagates the 

orbit) for the specified time to determine if, on any pass through the atmosphere, the heat rate 

limits are violated. For this study, the heat rate limits were adjusted so that the reference and 

operational missions end at approximately the same time. Therefore, the purpose of the 

operational mission run-out is to obtain an initial guess for the corridor limits that will be used 

for the ORT. Additionally, maneuvers are only made when a predicted maximum heat rate is 

outside the corridor limit. The maneuver size is not optimized to provide the exact target heat 

rate (e.g., 0.1125 W/cm2

), rather it calculates the maneuver size required to reduce the predicted 

orbit’s maximum heat rate back to the middle of the corridor. A detailed description of the 

corridor as it pertains to aerobraking is provided in the section below. 

Figure 7.2-1 shows the maximum heat rates near periapsis for the reference and operational 

mission run-outs. The operational mission run-out determined that the initial lower and upper 

corridors set at 0.85 and 0.14 W/cm2 respectively, would provide an approximate operational 

mission ending in nearly 70 days. The immediate action line was defined as 100 percent of the 

upper corridor or 0.28 W/cm2

 to provide margin similar to the Mars ODY aerobraking mission. 

Figure 7.2-1. Reference and Operational Mission Run-outs used to Establish Criteria for Phase 2 

AADS Evaluation within an ORT 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

35 of 97 

NESC Request No.: TI-09-00605 

ABM Sweep: This simulation considers only a short subset of the whole aerobraking mission, 

known as the aerobraking planning period (APP). The APP is the specified future time span 

(e.g., up to 1 week) for which a maneuver decision needs to be made and the subsequent passes 

through the atmosphere are evaluated to verify there are no violations of the corridor. The 

simulation contains a menu of 'V options available at the next available apoapsis opportunity 

(see Maneuver Menu below). The simulation runs a separate trajectory for each option, 

performing the maneuver at the specified apoapsis and allowing the orbits to propagate through 

subsequent aerobraking passes until the specified time is met. The resulting max heat rates from 

each simulation are plotted and evaluated. A ground-based maneuver recommendation is made 

based on the simulated maneuver trajectory. The decision process is outlined in more detail in 

Section 7.6.1.3. This simulation provides analysis with which to base the maneuver 

recommendation. 

Weekly Reset: The weekly reset simulation is similar to the reference mission run-out in that it 

allows a maneuver on every apoapsis and targets a specified heat rate. However, the simulation 

is initiated multiple times with the latest spacecraft state and simulates flight assuming the 

spacecraft flew exactly the upper, middle and lower corridor heat rates. This simulation offers 

an assessment of the end game performance margins. 

AADS Mission Run-out: The AADS mission run-out is similar to the reference mission run-out, 

but it differs in how and when a maneuver is performed. The AADS uses its prediction of the 

atmosphere and the SP to determine on which apoapsis to perform a maneuver and how big the 

maneuver will be. The AADS mission run-out can allow a maneuver on every orbit. 

Spacecraft Flight Simulation: Also known as the “truth” simulation, this simulation emulates 

the “real” environment and the spacecraft flight though it. The simulation uses the same 

spacecraft parameters, but flies through observed Mars atmosphere aerobraking density profiles 

rather than MarsGRAM. The simulation is intended to emulate the NAV function of uploading a 

maneuver to the spacecraft and the DSN communication link providing the latest spacecraft state 

to the AADS and ground-based simulations. 

7.2.2 AA High-Fidelity Simulation 

The bulk of the truth model development for the AAHFS was done in Phase 1. This simulation 

environment is capable of 6-DOF simulations for aerobraking missions at Mars, Venus or Titan. 

In Phase 1, each of these central bodies used the GRAMs (MarsGRAM, VenusGRAM, and 

TitanGRAM), but Phase 2 added the capability to run simulations by using the 

nondimensionalized flight atmospheric data described in Section 7.1.6. This allows 6-DOF 

testing of the AADS with atmosphere data that reflects characteristics that have been observed 

by prior aerobraking missions (ODY, MRO and MGS). Although the AAHFS was not used as a 

part of the ORT conducted in late January 2013, this capability was put in place to allow 6-DOF 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

36 of 97 

NESC Request No.: TI-09-00605 

testing of ORT-like mission scenarios. Test results using these flight atmospheres are provided 

in the next section. 

7.2.2.1 AAHFS AADS Testing 

In Phase 1, testing in AAHFS was performed for Mars aerobraking scenarios. Following the 

development of the SP in Phase 2, AADS was tested in AAHFS for Mars, Venus and Titan. All 

bodies were tested using 6-DOF simulations, although the attitude dynamics were simplified by 

reducing the magnitude of the aerodynamic torques. This simplification was made to accelerate 

the development of the test cases since a substantial effort would be required to modify the 

vehicle aerodynamic characteristics and attitude control system design to produce well-behaved 

rotational motion during (and following) the drag pass. This control system design effort for the 

notional vehicle used for AADS testing was deemed beyond the scope of this project, and this 

design would follow established procedures and is not considered germane to demonstrating an 

AA capability. Any mission using AADS would use the detailed vehicle and attitude control 

system design, and the AAHFS could easily be adapted to these characteristics. 

AAHFS –AADS at Mars 

Initial testing in a simulated Mars environment highlights the performance advantages of the SP 

used in Phase 2 over the EE used in Phase 1. Figure 7.2-2 shows the performance of the SP at 

Mars; the figure shows errors in the AADS translational state when compared with the truth 

model orbital integration. The SP is able to accurately maintain the current spacecraft state via 

the state determination (SD) process, as the reconstructed periapsis time drifts by about 

0.3 second in the worst-case over a 28-day integration arc. (The term “reconstructed” is used in 

this context to mean a state prediction that makes use of historical accelerometer measurements, 

and this state prediction is at the current time. This is opposed to a predicted state, which may be 

as much as one orbit into the future.) This shows that given near-perfect acceleration 

measurements by the IMU, the onboard integration can be nearly as accurate as one done on the 

ground. This is markedly (~10-30 times) better than the performance demonstrated by the EE in 

Phase 1 within this same simulation. The predicted timing (generated by the OP) for periapsis 

and apoapsis show similar accuracy. The spikes in the periapsis timing prediction occur when 

the spacecraft executes a corridor control maneuver, as the OP has no knowledge of the 

maneuver in the orbit prediction process. Therefore, the apoapsis maneuver introduces a 

perturbation into the predicted periapsis timing, but it does not affect the apoapsis timing as the 

maneuver immediately follows apoapsis. Figure 7.2-3 shows a similar trend for the 

reconstructed periapsis altitude, indicating that the altitude error is less than 40 m after 28 days 

of propagation. This accuracy is well below that required to perform the AE and ME tasks, even 

with ground updates 28-days apart (four times the study-defined requirement of weekly ground 

updates). Of course, the accuracy in the integration is driven by errors in the accelerometer 

measurements, which are not included in this simulation, but this case does demonstrate 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

37 of 97 

NESC Request No.: TI-09-00605 

excellent accuracy of the SP. Using the SP does not markedly improve the corridor 

performance, as shown in Figure 7.2-4, as the uncertainty in the corridor control variable 

(e.g., heat rate) is dominated by the atmospheric variability and not driven by errors in the 

periapsis altitude or velocity predictions. However, while the error introduced by atmospheric 

variability does not appreciably grow with time, the state prediction error grows quadratically 

with the integration interval, so increasing intervals between ground updates can lead these SP 

error terms to have a significant effect on corridor performance. 

Figure 7.2-2. SP Apses Timing Performance at Mars in AAHFS without Accelerometer Errors 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

38 of 97 

NESC Request No.: TI-09-00605 

Figure 7.2-3. SP Periapsis Altitude Performance at Mars in AAHFS without Accelerometer Errors 

Figure 7.2-4. AADS Corridor Performance at Mars in AAHFS without Accelerometer Errors 

AAHFS –AADS at Titan 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

39 of 97 

NESC Request No.: TI-09-00605 

AAHFS testing was conducted with a mission scenario at Titan, and these results are shown in 

Figures 7.2-5 through 7.2-7. As with Mars, the SP is able to accurately propagate an initial state 

over a ground update interval of 28 days. Apses timing errors remain small during this 28-day 

integration interval; although, the larger maneuvers necessary at Titan (a factor of ~50 larger 

than those at Mars) introduce a correspondingly larger spike in the periapsis prediction accuracy 

from the OP. Periapsis altitude errors remain below 70 m worst-case, which is negligible when 

compared with the 150 km altitude corridor used in the simulation. This corridor is shown in 

Figure 7.2-7, and the AAHFS results are nearly identical to the POST2 results shown in the 

Phase 1 final report. The Titan atmosphere does not show the same variability as at Mars, and 

the smooth nature of the corridor plot reflects this atmospheric uniformity. 

Figure 7.2-5. SP Apses Timing Performance at Titan in AAHFS without Accelerometer Errors 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

40 of 97 

NESC Request No.: TI-09-00605 

Figure 7.2-6. SP Periapsis Altitude Performance at Titan in AAHFS without Accelerometer Errors 

Figure 7.2-7. AADS Corridor Performance at Titan in AAHFS without Accelerometer Errors, 

Overlaid by POST2 Results from Phase 1 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

41 of 97 

NESC Request No.: TI-09-00605 

AAHFS –AADS at Venus 

AAHFS testing was conducted with a mission scenario at Venus, and these results are shown in 

Figures 7.2-8 through 7.2-10. As with Mars and Titan, the SP is able to accurately propagate an 

initial state over a ground update interval of 28 days. Apses timing errors remain small during 

this 28-day integration interval. Periapsis altitude errors remain below 15 m worst-case. Venus 

simulations use a temperature corridor, exercising the thermal RSE software inside the AADS. 

For this case, there are no truth model results shown in Figure 7.2-10, as there is no thermal truth 

model running in AAHFS. For this case, it is useful to compare the AAHFS results to the 

POST2 cases shown in the Phase 1 final report; such a comparison shows that the AADS 

performance remains consistent between AAHFS and POST2, as expected. 

Figure 7.2-8. SP Apses Timing Performance at Venus in AAHFS without Accelerometer Errors 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

42 of 97 

NESC Request No.: TI-09-00605 

Figure 7.2-9. SP Periapsis Altitude Performance at Venus in AAHFS without Accelerometer Errors 

Figure 7.2-10. AADS Corridor Performance at Venus in AAHFS without Accelerometer Errors 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

43 of 97 

NESC Request No.: TI-09-00605 

The results presented in this section demonstrate consistent performance for AADS when 

comparing results from AAHFS with POST2 for a variety of conditions. The POST2 results are 

shown in the Phase 1 final report. Three different central bodies were used (Mars, Titan, and 

Venus), as well as three different corridor types (heat rate, altitude, and temperature). These 

results also demonstrate that when the SP provides nearly perfect accelerometer data, the AADS 

orbit propagation can retain sufficient accuracy over long time spans (28 days or longer) to have 

a negligible contribution to corridor control errors. Note that the simulated accelerometer data 

does not perfectly represent the true acceleration as software latencies and quantization error are 

an inevitable part of the IMU modeling, and these data in the above simulation results do not 

contain noise, bias, alignment or scale factor errors which be present in flight accelerometers. 

7.2.2.2 Embedded MATLAB (EML) Version of AADS 

With the proof-of-concept Phase 1 design complete, a more holistic approach was possible to 

reduce software redundancy and streamline the algorithms. The functional decomposition of 

AADS in Phase 1 allowed different institutions to bring their expertise to bear on the segments of 

the AA problem, but with a completed software design, streamlining opportunities were 

apparent. In order to realize additional performance gains, an effort was undertaken to convert 

all AADS code to EML. This process allowed restructuring algorithm boundaries, allowing 

further reductions of the code memory footprint and computational overhead. The block 

diagram for this implementation is shown in Figure 7.2-11. With the additional algorithm 

modifications to the AE made in Phase 2, the terms in the regressor and observation matrices in 

the least-squares estimation problem (ALS and bLS) can be built recursively, eliminating the need 

to batch process a large vector of input data in the AE, and thereby obviating the need to pass 

any buffered data to the AE. Although not readily apparent in the block diagram, by converting 

to EML, the software tasks are easily grouped by execution rate, making this version of the 

software more amenable to adaptation for an embedded environment. It also streamlines the 

software interfaces and data structures, eliminating redundancy and making more consistent, 

readable code. EML also offers rapid, near push-button C-code and executable generation, 

suitable for a variety of host environments, affording developers the ease of the MATLAB 

coding and debugging environment and the simulation performance of compiled code. 

Ultimately, this makes the code more portable and adaptable for use in a variety of flight mission 

host spacecraft. 

Because this EML version of AADS was constructed late in Phase 2, this code was only tested in 

the AAHFS environment, and not in POST2. Plots for the EML results for Mars, Venus, and 

Titan are qualitatively identical to the C-version of AADS, and are therefore not shown. Minor 

algorithm modifications described above prevent quantitatively identical results, as expected. 

Benchmarking the execution speed of the EML AADS against the C-version should be 

performed in further development of AA. 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

44 of 97 

NESC Request No.: TI-09-00605 

Figure 7.2-11. Block Diagram for the EML Version of AADS 

7.3 Monte Carlo Analysis 

The objective of the AADS Monte Carlo analysis was to demonstrate that AADS is robust to 

environmental dispersions like maneuver errors, and aerodynamic and atmosphere uncertainties. 

The analysis was performed using 2000 3-DOF POST2 simulations of the AADS mission runout. It should be noted that there was no Monte Carlo of the AADS module (i.e., no multiple 

simulations on the spacecraft computer). Each AADS mission run-out simulation was evaluated 

to determine AADS performance degradation as compared to a nominal AADS mission run-out. 

The full list of Monte Carlo parameters considered is shown in Table 7.3-1. 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

45 of 97 

NESC Request No.: TI-09-00605 

Table 7.3-1. Monte Carlo Parameters 

Parameter Nominal Perturbation Distribution Rationale

Maneuver 

Pointing Errors 

Inertial 

DEJ

+/- 5 deg Normal Angular offset applied to apoapsis velocity vector. 

Maneuver 

timing errors 

0.0 +/- 100 s Normal Estimated apoapsis time +/- perturbation to see 

effect of starting burn early/late. 

'V (m/s) / 

Maneuver Burn 

Duration 

AADS 

calculated 

'V 

+/- 5 % Normal Normally, engine is "on" until AADS calculated 

'V is achieved. To handle burn duration and 'V 

parameters, allow applied 'V= AADS calculated 

'V + perturbation. 

Lift Coefficient 

Multiplier 

1 +/- 10 % Normal Larger than expected multiplier provides 

sensitivity analysis. 

Drag 

Coefficient 

Multiplier 

1 +/- 10 % Normal Larger than expected multiplier provides 

sensitivity analysis. 

Atmosphere 

Random Seed 

1 1-29999 Integer 

Uniform 

Value used by the atmosphere program to 

determine perturbation of the density and wind 

profiles. The range that is allowed by the Mars 

atmosphere program used in the simulation. 

Dusttau 

(Mars only)

0.5 0.1:0.9 Uniform This determines the dust loading and thus the 

density and wind profiles that the vehicle will 

experience. This range provides large variability, 

but would not include dust storms. 

Histogram results of the primary output parameters of interest are provided. Note that all of the 

Monte Carlo input parameters are varied normally except for those related to the atmosphere; 

however, the distribution in the output parameters was not expected to be Gaussian due to the 

fact that many of the processes involved in aerobraking (e.g., utilizing a corridor) are nonGaussian in nature. Monte Carlo analyses were performed to evaluate the effect of the corridor 

on the distributions by flying a narrow corridor at the target heat rate value. The results verified 

the corridor was a primary driver of the atypical distributions. Further analysis could also be 

performed to determine the extent of atmosphere effects. 

Results of the AADS Monte Carlo analysis indicate that AADS is robust to typical spacecraft 

aerodynamic, spacecraft maneuver, and atmospheric variations experienced at Mars. The final 

aerobraking duration for all 2000 cases fell within +/- 7 days of the nominal AADS mission runout aerobraking duration. The statistics are shown in Figure 7.3-1. The green dot in the 

histograms denotes nominal AADS mission run-out value while the red dot denotes the mean of 

the 2000 Monte Carlo simulation. 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

46 of 97 

NESC Request No.: TI-09-00605 

Figure 7.3-1. AADS Monte Carlos Results: Aerobraking Duration 

The variation in the local true solar time (LTST) of the equatorial ascending node crossing at the 

end of each Monte Carlo AADS mission run-out was less than 15 minutes (3-sigma) from the 

nominal, see Figure 7.3-2. This is consistent with the actual performance criteria imposed on the 

ODY and MRO aerobraking missions for LTST at the end of aerobraking. Note that AADS 

meets the performance criteria used in previous missions without any intervention once it is 

started during the aerobraking phase of the mission. In actual flight operations, the corridor 

could always be adjusted in a weekly reset meeting such that AADS could end closer to the 

mission desired LTST if the performance criteria for another mission is less than 15 minutes. 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

47 of 97 

NESC Request No.: TI-09-00605 

 

Figure 7.3-2. AADS Monte Carlos Results: Final LTST 

The 3-sigma variation in the number of AADS commanded maneuvers, required over the 

duration of each AADS mission run-out Monte Carlo case, was approximately 13, with the 

maximum case at 80 (versus 46 for the nominal AADS mission run-out). This variation is quite 

small given the nominal AADS mission run-out required ~425 orbits, and AADS can perform a 

maneuver each orbit apoapsis if necessary, see Figure 7.3-3. 

Figure 7.3-3. AADS Monte Carlo Results: Number of Maneuvers 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

48 of 97 

NESC Request No.: TI-09-00605 

Despite the range in number of maneuvers required per AADS mission run-out case, the actual 

variation in total 'V was just under 3 m/s (3-sigma) across all 2000 cases, with a maximum case 

approximately 8 m/s higher than the nominal AADS case. This variation in total 'V is small 

enough to be accommodated in a typical aerobraking 'V budget, see Figure 7.3-4. 

Figure 7.3-4. AADS Monte Carlo Results: Total ABM 'V 

Figure 7.3-5 shows an example of the total heat rate dispersion for a 2000 case Monte Carlo 

analysis utilizing the MarsGRAM atmosphere. The black dots show a comparison to the 

nominal mission, and the black dashed lines represent the upper and lower corridor limits used 

by the AADS. The typical immediate action line threshold of 100 percent above the upper 

corridor in this case would be located at a heat rate of 0.34 W/cm2

; thus all of the perturbed cases 

have heat rates that stay below the immediate action line. Figure 7.3-6 shows a histogram of the 

same data where the instances of heat rate in 0.01 W/cm2

 bins can be seen. The statistical mean 

heat rate (0.139 W/cm2

) is right at the middle of the corridor, and the 99 percentile heat rate is 

19 percent above the upper corridor limit. These two plots, along with the other Monte Carlo 

results, suggest that AADS can accommodate significant variation in flight parameters while 

successfully completing missions with flight-like criteria. 

Figure 7.3-7 shows the corresponding Monte Carlos periapsis timing errors through the mission. 

Increased timing error demonstrates propagated errors in the SP. The goal with the AADS 

weekly updates is to minimize the periapsis timing estimate. It is noted that even though the 

timing errors for the 1 and 99 percentile cases are nearly twice the periapsis timing errors in the 

nominal case, the performance capability of AADS is not compromised. This is evidenced by 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

49 of 97 

NESC Request No.: TI-09-00605 

the previous discussion of the Monte Carlo analysis results for total 'V, aerobraking duration, 

glide slope, number of maneuvers, and number of occurrences above the upper corridor limit. 

 

Figure 7.3-5. AADS Monte Carlo Heat Rate Distribution 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

50 of 97 

NESC Request No.: TI-09-00605 

Figure 7.3-6. AADS Monte Carlo Results: Heat Rate Statistics 

Figure 7.3-7. AADS Monte Carlo Results: Periapsis Timing Errors 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

51 of 97 

NESC Request No.: TI-09-00605 

Additional values that could be perturbed in future Monte Carlo analysis include, spacecraft 

mass, Mars dust storm variables and initial state conditions. Likewise, trade studies, using 

randomly selected normalized MGS, ODY, or MRO atmospheres, as described in Section 7.1.6 

might provide insight to the performance using realistic profiles instead of MarsGRAM. These 

studies will determine which aspects of the AADS are most vulnerable. 

7.4 Processor Load Analysis/Benchmarking 

Demonstrating that processor capabilities can support the software requirements is critical to 

ensuring the AADS utility in a flight application. Although the code as written is not yet suitable 

for implementation on a flight processor, the software is mature enough to perform a preliminary 

assessment of the feasibility of its use in flight. It is time-consuming and expensive to port 

AADS to a real flight processor; this task is beyond the scope of Phase 2. However, simple 

benchmarking was done using the existing code running in a workstation that provides a rough 

order-of-magnitude assessment of the quantities of interest. Phase 1 identified characterizing the 

memory requirements and central processor unit (CPU) loading for the AADS a top priority. 

During Phase 1, there was a concern that the memory requirements for the AADS would make 

execution in a flight processor difficult. At that time, the AADS memory requirements were 

largely due to the substantial buffers required for the attitude and acceleration data throughout 

the drag pass for later batch integration by the EE. These buffers were on the order of 

3 megabyte (MB), which would complicate the memory handling in a flight processor. In 

addition to the buffered data from the spacecraft to the EE, there were smaller buffers of data 

from the EE to the AE that were approximately 0.5 MB. Because of the architecture changes 

made when converting the EE to the SP in Phase 2, these memory issues were resolved. Both of 

the large buffers noted above have been eliminated, and the memory footprint of AADS has been 

reduced by over an order of magnitude. 

Assessing the processor loading for the AADS is essential to demonstrating the computational 

feasibility of AA. Actually testing the AADS on a flight processor is prohibitively expensive, 

but by performing timing tests of AADS running in a workstation, an assessment was made 

about the expected processor loading running in an embedded environment. The approach taken 

for this study was simply to run the AADS in the AAHFS simulation and to time each portion of 

the AADS that would correspond to a flight software task. These times obtained by the 

workstation testing were scaled based on the ratio of the clock speed of the workstation-to-flight 

processors to estimate the CPU time each task would require on a flight processor. Comparing 

this scaled time with the expected CPU time available for each task provided an estimate of the 

processor loading for each task. The testing described in this section used the AADS running on 

a workstation PC with a clock speed of 2670-MHz. As an example, the MESSENGER mission 

flight processors had a clock speed of 25-MHz, so the ground test times were scaled up by 

107 times to estimate the equivalent CPU time required for a notional flight processor. The 

flight processor clock speed used in this analysis is considered conservative, as MESSENGER 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

52 of 97 

NESC Request No.: TI-09-00605 

processors were purchased nearly a decade ago, and current-day missions are routinely running 

faster than 25-MHz. 

In its current form, AADS is not explicitly decomposed into separate tasks as would be necessary 

for flight software. As a result, the benchmarking must only consider those functions that are 

related to an expected software task. This means that the benchmarking must carefully collect 

timing information on software functions that are only a part of the expected three tasks that 

comprise AADS. The tasks are: a high-rate task (notionally 10-Hz) for acceleration/attitude data 

processing, integration, and buffering; a low-rate task (notionally 0.01-Hz) for evaluation of the 

gravity models, integration of the spacecraft orbit, and preparation of the least-squares estimation 

matrices; and a batch process (executed once per orbit) that performs the orbit prediction and the 

subsequent AE and ME tasks. Each task will have different processor priorities and temporal 

allocations for completion. 

The highest-rate process found in the AADS is the input data processing and high-rate 

integration function in the SP. There are few calculations associated with this software task, as it 

merely performs a trapezoidal integration of the acceleration data and buffers the resultant ∆V 

for use in the low-rate integrator. Although this task runs at a high rate, the processing required 

is minimal (a few multiplications and sums) and was not characterized, as it is by far the least 

concerning of the three processing tasks. It is well known that substantially more complicated 

algorithms are routinely running in spacecraft flight software at markedly higher rates (greater 

than 100-Hz), so no effort was devoted to benchmarking this task. 

The low-rate integration task is the most challenging from a processor loading point of view. 

The SP is architected such that the low-rate integration attempts to step across a 100-s time 

period. During the exo-atmospheric flight, these integrations can routinely be performed in one 

or two integration steps (each of which requires 16 evaluations of the gravity models). During a 

drag pass, the integration steps can be much shorter (a few seconds each) requiring many more 

integration steps and many more gravity model evaluations. The processor loading will increase 

during each drag pass as the dynamics necessitate smaller integration steps to maintain tolerance, 

and because the variability in the atmosphere can require some back stepping in the integrator 

(essentially wasted integration steps). Benchmarking this task was performed at Mars, with a 

21-degree and order gravity model, which is currently assumed to be the highest-fidelity gravity 

model considered for AADS testing (higher ordered gravity models are available, but do not 

offer significant enough increase in accuracy over the 21-degree and order field for purposes of 

aerobraking analysis). The results of this benchmarking throughout an entire Mars aerobraking 

mission are shown in Figure 7.4-1. Each drag pass sees a spike in the execution time for the 

low-rate integrator due to the shortened time steps. As the mission progresses, these spikes 

become more frequent as the orbital period drops and become more prolonged as the drag pass 

duration increases, but the peak value is not markedly affected. In the worst case, a flight 

processor would require approximately 1.6 seconds to complete the low-rate integration. Note 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

53 of 97 

NESC Request No.: TI-09-00605 

that the resolution of the workstation timing data is 1-ms, so the scaled time is quantized at 

107 ms. Since this task is assumed to run at 0.01 Hz, this reflects a processor loading (for this 

task only) of less than 2 percent. This result suggests that the processor loading for the low-rate 

integration is not an issue. If higher degree and order gravity models are used, there could be 

some increase in the expected peak processor loading, so any mission using the AADS would 

have to perform a CPU loading versus integration accuracy trade study. 

Figure 7.4-1. AADS Low-rate Integrator State Determination Scaled Time for Mars Case using 21x21 

Central Body Harmonic Gravity Model 

The AADS batch process that performs the orbital propagation and the atmosphere and 

maneuver calculations is by far the most computationally expensive task to complete and it will 

run at the lowest software priority. Despite the challenges that this task presents, the batch 

process is not expected to pose an issue for processor loading as there is a substantial amount of 

time to complete this task. It is assumed that the task will be triggered shortly after the drag pass 

exit, and must complete with sufficient time to configure the spacecraft for any corridor control 

maneuver. In the worst case, this affords the task approximately 25 percent of the minimum 

aerobraking orbit period to complete, or ~30 minutes. Timing tests were done to characterize the 

expected processor time required to complete this task, and the results are shown in Figure 7.4-2. 

For this task, the orbital propagation takes longer in the long period orbits, so the worst-case is 

early in the mission (when the task has much longer to complete). In all cases, the batch process 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

54 of 97 

NESC Request No.: TI-09-00605 

would require much less than 1 percent of the available CPU, presenting no issue for the timely 

completion of this task. 

Figure 7.4-2. AADS Low-rate Integrator Orbit Propagator Scaled Time for Mars Case using 21x21 

Central Body Harmonic Gravity Model 

7.5 Other Trades and Sensitivity Analysis 

7.5.1 Accelerometer Sensitivities

The analysis described in Section 7.2.2.1 is limited to cases without error sources that would 

affect the orbital integration. There are several notable error sources that could substantially 

affect the SP, including OD, accelerometer, and gravity model errors. Each of these error 

sources has been explored independently, but further analysis is needed to understand the 

coupling between these errors. 

The SP relies on an accurate initial condition for its orbital integration, since there are no 

external measurements available to correct the integration. As a result, the integration is 

performed as a dead reckoning, and the only mechanism to correct drift in the onboard state is to 

perform a ground update. It is helpful to understand the uncertainty associated with the onboard 

state given an initial uncertainty associated with the ground update state. In other words, the 

navigation process that defines the ground update is uncertain, and this leads to greater 

uncertainty as the onboard state is propagated forward in time. The growth in this state 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

55 of 97 

NESC Request No.: TI-09-00605 

covariance was computed using the Analytical Graphics, Inc. OD tool-kit and is shown in 

Figure 7.5-1 for a sample 7-day orbit propagation. The analysis assumes a typical initial 

uncertainty associated with the spacecraft state (1 m position uncertainty per axis and 1 mm/s 

velocity uncertainty per axis). The uncertainty growth shown in this figure demonstrates that 

after 7-day propagation, a typical initial state uncertainty can result in periapsis altitude errors of 

~1 km. Cases were run with and without accelerometer errors. The two curves are largely 

indistinguishable, as the growth in the altitude uncertainty is dominated by the initial navigation 

error, and only mildly affected by the accelerometer error sources. 

Figure 7.5-1. Ground Update State Uncertainty Propagated over 7-days 

A limited assessment of accelerometer errors is described in this section, but a more extensive 

study is necessary to fully understand the impact these errors have on the state uncertainty, 

particularly for long propagation time spans. Because AAHFS has a full IMU model with 

adjustable noise, bias, alignment, and scale factor errors, this model is readily available for 

testing. Error parameters were set consistent with Northrop Grumman Scalable Space Inertial 

Reference Unit used for MESSENGER to quickly assess feasibility of AADS in the presence of 

realistic accelerometer error conditions. This is an identical analysis to that shown in the Phase 1 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

56 of 97 

NESC Request No.: TI-09-00605 

final report, although this testing used the updated Phase 2 AADS (running the SP instead of the 

EE). Qualitatively, the AADS is able to maintain reasonable corridor performance when using a 

7-day ground update interval, as shown in Figure 7.5-1. Quantitatively, the drift in the onboard 

SP from the truth model (due to the accelerometer errors) is much more substantial than the 

error-free case, shown in Figures 7.5-2 and 7.5-3. These errors were not applied to previous 

study. 

The drift in the orbit propagation due to accelerometer errors, presented in Figures 7.5-2 and 

7.5-4, is a notional example. Most of the accumulated error in the SD process of the SP is due to 

uncompensated accelerometer bias, as this error causes drift in the state. Because this bias is 

fixed in the spacecraft body frame, its orientation with respect to the planet-relative velocity 

direction will also remain fixed. This will ultimately result in a bias in the drag pass ∆V 

estimation, leading to nearly monotonic growth in the periapsis timing estimate (and to a lesser 

degree, periapsis altitude). The envelope of possible orbit errors were not characterized by this 

analysis, as only sample accelerometer biases were used that characterize one possible outcome 

with a reasonable accelerometer bias. The possible biases errors themselves map into an 

ellipsoid in the body frame, which in turn translates into an error ellipsoid in the ∆V estimated at 

each drag pass. Further analysis would be required to characterize the worst-case state errors 

associated with the accelerometer errors. The AAHFS is sufficient to conduct this analysis, but 

it was not undertaken in Phase 2 due to time constraints. This analysis would require numerous 

simulations to perform the Monte Carlo or sigma-point analysis and each simulation requires 

more than 12 hours to complete. 

Figure 7.5-2. Corridor Performance for Phase 2 AADS with 7-day Ground Update and Accelerometer 

Errors 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

57 of 97 

NESC Request No.: TI-09-00605 

Figure 7.5-3. Apses Timing Performance for Phase 2 AADS with 7-day Ground Update and 

Accelerometer Errors 

Figure 7.5-4. Periapsis Altitude Performance for Phase 2 AADS with 7-day Ground Update and 

Accelerometer Errors 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

58 of 97 

NESC Request No.: TI-09-00605 

7.5.2 Gravity Model Sensitivities 

All simulations described above and in the Phase 1 report used identical gravity models in the 

truth model and the AADS. Whatever gravity model would be used in flight would not perfectly 

represent the true gravity conditions, and one way to assess the performance of AADS to gravity 

modeling errors is to introduce a difference between the AADS gravity model and the truth 

model. This study gives some insight into what fidelity model is required in flight to maintain 

good SP performance, in terms of maintaining an accurate periapsis timing estimate and altitude. 

Table 7.5-1 summarizes the analysis performed at Mars using the EML version of AADS, 

running with a 7-day ground update. In all cases, the truth model central body harmonic gravity 

model used a degree and order 21. When the SP uses the same 21x21 model, the errors after 

7 days remain negligible, which is expected for this error-free case. As the degree and order of 

the gravity model is decreased, there is a nonlinear growth in the periapsis metrics. The lowestfidelity gravity model tried (3x3) shows errors that are quite large, and the corridor control for 

that case showed several prolonged excursions outside the design corridor that were only 

corrected with a ground update. This assessment provided a rough understanding of the gravity 

model fidelity that is required for reasonable AADS performance. From these results, it appears 

that low-fidelity gravity models are insufficient to ensure adequate corridor performance and 

given the modest increase in SP execution times required for higher-fidelity model evaluation, a 

good balance can be achieved between processor loading and integration accuracy. 

This analysis would need to be repeated for any host mission, as it depends heavily on the central 

body, the reference trajectory path relative to central body gravity anomalies, and the accuracy of 

the gravity model at a particular degree and order. These trades can be quickly performed using 

AAHFS or POST2, as the software parameters that define the gravity model fidelity are easily 

adjusted. 

Table 7.5-1. Impact of Gravity Model Errors on SP Accuracy and Timing for 7-day Ground Update 

at Mars when Compared Against Fixed 21 degree and order Gravity Truth Model 

AADS Mars Harmonic 

Gravity Model 

Degree and Order

Worst-case 

Reconstructed 

Periapsis Timing Error 

(s) 

Worst-case Periapsis 

Altitude Error (m)

AADS Execution Time 

on Ground 

Workstation (ms)

21x21 0.15 3 15 

17x17 2 120 10 

13x13 6 320 6 

9x9 9 680 5 

3x3 100 2400 3 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

59 of 97 

NESC Request No.: TI-09-00605 

7.6 AADS Performance Assessment 

During Phase 1, extensive testing of AADS was performed using an engineering model of the 

atmosphere with dispersions to demonstrate feasibility of the algorithm. Phase 1 concluded that 

AADS had passed the initial feasibility tests, but needed to be rigorously tested in an operational 

flight-like environment. In Phase 2, AADS performance was verified using a more stressing 

aerobraking scenario than was experience by the Mars ODY aerobraking mission which was the 

most aggressive (maintaining the least amount of thermal margin) to date. To test the AADS in 

the most realistic sense, an ORT was used. An ORT is designed to evaluate the ability of a team 

to perform all mission critical tasks prior to the commencement of a mission. Because there is 

no identified flight vehicle for this testing, the Phase 2 demonstration is not an ORT in the most 

proper sense. However, the concept of the ORT provides an ideal testing environment for the 

AADS. The key to a successful ORT is the simulation and processes that recreate the flight and 

operational experience. By focusing the AADS Phase 2 analysis as an ORT at Mars, the 

simulation could take advantage of observed density profiles from the three previous aerobraking 

missions at the planet (MGS, ODY and MRO) and demonstrate the AADS capabilities in a 

realistic mission operations setting. The performance and advantages of the AADS are evaluated 

by running two aerobraking ORTs simultaneously: one in the historical ground-based manner 

(ORT1) and one using AADS (ORT2). So, not only is the AADS prediction and reaction being 

compared with what happens on a spacecraft (ORT2), but it is also compared to the decisions 

that a ground-based team would have made with a similar spacecraft and environment (ORT2 

versus ORT1). Additional variations of ORT2 were run after the initial assessment to evaluate 

additional AADS features including atmosphere dispersions in the maneuver logic (ORT2a) and 

the effect of using constant corridors (ORT2b). 

The ORT scenario was deliberately challenging, in that it was designed to aerobrake from an 

18-hour period orbit to an orbit with a 2800 km apoapsis altitude in ~70 days with a margin of 

only 100 percent (see the definition provided in Section 7.2.1), the same as for Mars ODY. The 

ground-based team met each simulation “day” during ORT1 (e.g., after each 24 hours of 

simulated mission elapsed time), and decided on the maneuver strategy for that day. Every week 

of mission elapsed time, the ground-based team also met, examined the latest mission run-out, 

and decided if the corridor needed to be adjusted. The AADS Team met only weekly. Both 

ORTs successfully achieved the final mission requirements, but the AADS Team required fewer 

resources, namely no daily meetings, and only needed weekly updates from the Atmosphere 

Team for the mission run-outs. 

The reference mission, operational procedures, simulations, and results for the Phase 2 

aerobraking ORTs are described in Section 7.6.1. 

In addition to the ORT, AADS was used for additional analyses, including simulated replication 

of the ODY and MRO aerobraking missions. Discussions of these simulations are provided in 
NASA Engineering and Safety Center

Technical Assessment Report

Document #:

NESC-RP09-00605

Version:

1.0

Title:

Autonomous Aerobraking (Phase 2)

Page #:

60 of 97

NESC Request No.: TI-09-00605

Section 7.6.2 and 7.6.3, respectively. In summary, the AADS has shown to be a robust 

technique that offers more flexibility than the current ground-intensive method. 

7.6.1 Mars ORT

As previously mentioned, an ORT is designed to evaluate the ability of a team to perform all 

mission critical tasks prior to the commencement of a mission. The key to a successful ORT is 

the simulation and processes that recreate the flight and operational experience. Historically, 

real-time aerobraking operations involve dozens of individuals working at several NASA 

Centers, universities, and industry locations to download data, analyze the environment and 

spacecraft, to make maneuver decisions and finally, upload data to the spacecraft. 

Table 7.6-1 shows a listing of the typical aerobraking operations teams and their tasks during 

ground-based aerobraking operations. A successful aerobraking ORT must realistically 

“simulate” a core of each of these elements.

Table 7.6-1. Aerobraking Operations Ground-based Analysis Teams 

Team Purpose 

Spacecraft Evaluate and report on spacecraft health and safety 

Navigation Build and upload maneuver decision to spacecraft 

Flight Mechanics Perform analysis to inform maneuver decisions and evaluate mission 

performance margins 

Atmosphere Advisory Evaluate atmosphere conditions and report effect on mission and 

maneuver decisions 

Thermal Monitor spacecraft temperatures and infer heating for each orbit 

Accelerometer Convert accelerometer data into atmosphere data 

Aerodynamics Compare flight instrument data with model data, make updates as 

needed 

For the Phase 2 evaluation of the AADS, ORTs 1 and 2 utilized the Accelerometer, Flight 

Mechanics and Atmosphere Advisory Teams. The spacecraft is assumed to be healthy and safe 

unless otherwise indicated by derived aeropass maximum heat rate values. The navigation and 

accelerometer team tasks will be included in the spacecraft flight simulation (SFS). It is assumed 

that the SFS or “truth” simulation accurately represents the thermal and aerodynamic properties 

thus eliminating the need for representation from those teams during the ORTs. 
NASA Engineering and Safety Center

Technical Assessment Report

Document #:

NESC-RP09-00605

Version:

1.0

Title:

Autonomous Aerobraking (Phase 2)

Page #:

61 of 97

NESC Request No.: TI-09-00605

7.6.1.1 ORT Reference Mission 

Both ORTs attempted to follow the same ODY-like reference mission. The reference mission, 

initialized on April 9, 2009, was designed to reduce an initial orbit of 18 hours (270,000 km 

apoapsis altitude) to 3 hours (2800 km apoapsis altitude) in approximately 70 days. The 

reference mission ends at 2800 km because the AADS is not currently designed to accommodate 

collision and avoidance (COLA) considerations or the walk-out scenario, nor is it equipped with 

a lifetime constraint algorithm. Therefore, neither ORT will consider this late phase of 

aerobraking. The mission assumes a walk-in from 200 km to approximately 95 km periapsis 

altitude over six orbits, performing three maneuvers (one every other orbit) until the maximum 

heat rate is near the target value of 0.112 W/cm2

. The reference design mission simulation then

calculates the precise 'V required to obtain the target heat rate on each subsequent pass through 

the reference atmosphere (MarsGRAM2010). A plot of the periapsis heat rate versus time for 

the reference mission is shown in red in Figure 7.2-1 in Section 7.2.1. The simulation provided a

reference glide slope, or LTST of the ascending node versus time (as aerobraking duration in 

days or orbit number) plot, that was be used by the ground-based and AADS simulated missions 

to evaluate mission progress. The goal of both missions was to end at the reference mission final 

LTST of 10:46 hours +/-15 minutes. The 15-minute margin was a requirement set by the science 

instrument operability for ODY, but it also provided flexibility in terminating aerobraking during 

the end-game phase of the mission. The reference glide slope, with the 15-minute target bounds,

is shown in red in Figure 7.6-5 in Section 7.6.1.4. 

All simulations used a spacecraft model with the same reference area and aerodynamics as the 

MRO spacecraft, but the mass of the vehicle was reduced from ~1300 kg to 850 kg in order to 

lower the ballistic coefficient and allow the reference mission to complete the main aerobraking 

phase in the desired 70 days.

7.6.1.2 Atmosphere Team Analysis 

The Atmosphere Team function, preserved from past aerobraking missions, was utilized in both 

ORTs. Since neither ORT simulated a priori knowledge of the atmospheric density before entry, 

the MarsGRAM 2010 model was used for predicting the environment. The model assumed

smooth bell-shaped density curves that increase as altitude decreases. However, the actual flight 

atmosphere in the SFS or “truth” simulation, taken from the ODY mission, resulted in a 

difference between actual and predicted flight density profile. An example of the difference 

between the model and observed density profiles is shown in Figure 7.6-1. The Atmosphere 

Team monitored these differences and provided model updates (different for each ORT) in the 

form of a density multiplier (MarsGRAM parameter A0) to be applied only to those analyses 

predicting future aerobraking performance. The multiplier was derived from taking the ratio of 

the integrated density from the SFS (e.g., blue line in Figure 7.6-1) simulation to the integrated 

density from the MarsGRAM prediction (using A0=1). The multiplier was evaluated after each 
NASA Engineering and Safety Center

Technical Assessment Report

Document #:

NESC-RP09-00605

Version:

1.0

Title:

Autonomous Aerobraking (Phase 2)

Page #:

62 of 97

NESC Request No.: TI-09-00605

orbit, but was only changed daily for ORT1 and weekly for ORT2. Figure 7.6-2 shows the 

multiplier for each orbit and the value used in the simulation for both ORTs. 

Figure 7.6-1. Profiles of Observed Density (blue), Model Density with A0 =1 (solid red) and the Model 

Density Scaled to Match area under Observed Density Profile (dashed red) for a sample AADS Mission 

Orbit 209 
NASA Engineering and Safety Center

Technical Assessment Report

Document #:

NESC-RP09-00605

Version:

1.0

Title:

Autonomous Aerobraking (Phase 2)

Page #:

63 of 97

NESC Request No.: TI-09-00605

Figure 7.6-2. A0 Multipliers Observed and used in the Simulations 

7.6.1.3 ORT1: Historical Ground-Based 

To offer a realistic comparison to the operational advantages of using the AADS, a historical 

ground-based ORT was used to baseline the “typical” amount of work (number of tasks and time 

needed to perform them) required during aerobraking for evaluating spacecraft data and 

recommending maneuvers (e.g., to perform the tasks that AADS performs automatically). The 

walk-in phase of the mission, which is similar in both ORTs, executed three maneuvers over six 

orbits to gradually dip into the atmosphere, increasing the peak heat rate to be within the 

corridor. The corridor was determined in the operational mission run-out, described in 

Section 7.2.1, which resulted in an initial guess for the corridor to be 0.085 W/cm2 for the lower

bound and 0.14 W/cm2 for the upper bound. The goal of the ground-based Flight Mechanics

Team was to utilize the atmosphere predictions from the Atmosphere Team and perform analysis 

to determine when and what size of maneuvers should be made to keep the spacecraft heating 

within the required corridor in order to ensure that the mission followed the predetermined 

reference glide slope. 

During ORT1, the Flight Mechanics and Atmosphere teams performed two types of analyses to 

support maneuver decisions: daily and weekly. A description of both analyses is provided 

below. 
NASA Engineering and Safety Center

Technical Assessment Report

Document #:

NESC-RP09-00605

Version:

1.0

Title:

Autonomous Aerobraking (Phase 2)

Page #:

64 of 97

NESC Request No.: TI-09-00605

ORT1: Daily Analysis 

Daily analysis evaluates the potential effect of all maneuver menu options and utilizes the ABM 

sweep simulation described in Section 7.2.1. The simulation implements the density multiplier 

(A0) from the Atmosphere Team and the results were used to make a maneuver recommendation 

to the mission manager, who ultimately decides if a maneuver will be performed (i.e., uploaded 

to the spacecraft). During a real mission, the daily analysis begins by evaluating the “no 

maneuver” option, which is then used to determine a subset of maneuver options to evaluate. 

The likely optimal maneuver option is then further evaluated using Monte Carlo analysis to 

assess how the 1 and 99 percentile trajectories compare to the corridor. However, in the interest 

of time for the condensed Phase 2 ORT1 schedule (where a day of aerobraking was simulated in 

1 hour), all maneuver options were run simultaneously and no additional Monte Carlo analysis 

was performed. Figure 7.6-3 shows the results of the maneuver options on orbit 95 starting from 

the state at apoapsis of orbit 93. The thick red line at 0.28 W/cm2 denotes the immediate action

line and the constant red lines at 0.085 and 0.14 W/cm2 denote the upper and lower corridor

targets, respectively. The vertical black-dashed lines denote the planning period (the next 48 

hours) for which the spacecraft must remain within the corridor limits in the event that no 

maneuver can be uploaded. The potential effects of performing each maneuver (based on the 

MarsGRAM atmosphere and the density multiplier provided from the Atmosphere Team) are 

shown. As expected, a large negative maneuver, which would decrease the periapsis altitude 

(increase heat rate indicator), would likely result in heat rates above the upper corridor. 

Likewise, a large positive maneuver (i.e., 0.6 m/s) would likely raise periapsis altitude such that 

the resulting heat rate on subsequent orbits would be below the lower corridor. In this particular 

example, performing no maneuver would result in the spacecraft heat rates falling below the 

lower corridor throughout the planning period and is highlighted using a black-dashed line. For 

this case, the ground team chose a maneuver of -0.10 m/s during ORT1. 
NASA Engineering and Safety Center

Technical Assessment Report

Document #:

NESC-RP09-00605

Version:

1.0

Title:

Autonomous Aerobraking (Phase 2)

Page #:

65 of 97

NESC Request No.: TI-09-00605

Figure 7.6-3. ORT1 Daily Analysis 

ORT1: Weekly Analysis 

Like the daily analysis, the ground-based weekly analysis was based on past aerobraking mission 

flight experience. The purpose of the weekly analysis was to evaluate the margin to the targeted 

final orbit defined by the reference mission. This margin is determined by flying three simulated 

missions that target exactly the heat rate of the upper, middle, and lower corridor. The 

simulations are initiated with the latest spacecraft state at apoapsis and end at 2800 km apoapsis 

altitude (the point at which the use of AADS will cease). The results are then compared to the 

reference mission to determine if the corridor needs to be adjusted in order to meet the desired 

end conditions. Figure 7.6-4 shows the orbital period results of the three simulations that are 

initialized with an apoapsis state on day 16. Flying a mission near the upper corridor will result 

in a mission approximately 10 days shorter than the reference. However, flying the rest of the 

mission near the lower corridor will result in the aerobraking mission taking about 10 days 

longer than the reference. Figure 7.6-5 also shows the LTST margin for this same simulation, 

where the horizontal red lines denote the desired target range at the end of the mission (EOM).

In this case, if the mission continues to fly near the lower corridor, the mission will miss the 

target LTST of the ascending node by 20 minutes. Therefore, a near-term maneuver strategy 
NASA Engineering and Safety Center

Technical Assessment Report

Document #:

NESC-RP09-00605

Version:

1.0

Title:

Autonomous Aerobraking (Phase 2)

Page #:

66 of 97

NESC Request No.: TI-09-00605

should focus on remaining in the upper part of the corridor for a few orbits in order to pull the 

lower corridor mission back into the final mission limits during the next weekly run-out. 

The team may also choose to change the corridor limits based on the results of the weekly 

analysis. Depending on the results of the run-out, several options exist for modifications to the 

maneuver strategy for the upcoming week, such as: 

1. Modifying the upper corridor.

a. Increasing it reduces the margin to the immediate action line, but may serve as

a viable option to fly more aggressively in regions of the atmosphere that are

known to have less variability.

b. Decreasing the upper corridor would reduce the rate of change of the LTST,

which would be useful if the spacecraft was getting too far ahead of the

reference glide slope.

2. Modifying the lower corridor.

a. Increasing the lower corridor can allow the mission to move closer to the

reference if the mission is behind schedule. It also can result in additional

maneuvers being required.

b. Decreasing the lower corridor can lengthen the mission duration and reduce

the number of required maneuvers.

For ORT1, the corridor was modified three times throughout the 70-day mission. Figure 7.6-6 

shows the corridor with both the predicted and actual heat rates observed on each orbit. Because 

of the lower than expected densities (A0 were near 0.2 for most of the mission), much of the 

mission flew near or below the lower corridor (see the blue dots in Figure 7.6-6). To remain on 

the target glide slope, the lower corridor was raised. Additionally, because of the low densities, 

the ground team also decided to raise the upper corridor. This effectively reduced the thermal 

margin, but did so to make effective use of the atmosphere when risk was perceived to be low. 

Near the EOM, when the orbit period is less than 3 hours, there are 8 orbits occurring between 

maneuver opportunities. Maneuvers are determined based on the latest atmosphere trends

provided from the Atmosphere Team and the ABM sweep. However, in the ground-based 

ORT1, the variability in the atmosphere near the EOM was not captured in the model and 

therefore the heat rate on three orbits violated the immediate action line before a maneuver could 

be performed to raise the periapsis altitude effectively reducing the periapsis heat rate. 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

67 of 97 

NESC Request No.: TI-09-00605 

Figure 7.6-4. ORT1 Weekly Analysis: Orbit Period 

Figure 7.6-5. ORT1 Weekly Analysis: LTST 
NASA Engineering and Safety Center

Technical Assessment Report

Document #:

NESC-RP09-00605

Version:

1.0

Title:

Autonomous Aerobraking (Phase 2)

Page #:

68 of 97

NESC Request No.: TI-09-00605

Figure 7.6-6. ORT1 Heat Rate versus Days of Aerobraking 

ORT1 Maneuver Decision-making Process

During the ODY aerobraking mission, a maneuver could be performed only once per day. 

However, if a maneuver was made, it was designed such that under nominal conditions the next 

maneuver could not occur within the next 48 hours unless a maneuver was required the next day 

to protect the spacecraft. This policy mitigated the work of the ground team to build, test and 

upload maneuver commands. A similar policy was adopted for ORT1. The SFS was stopped 

every 24 hours. The state at the nearest apoapsis was written to file and provided to the ORT1 

Flight Mechanics Team as the spacecraft state file. The team would select on orbit apoapsis 

within the following 24 hours that was eligible to perform a maneuver should it be deemed 

necessary. In keeping with ODY and MRO fight experience, whenever possible, the orbit was 

selected to occur during the prime shift (between 8 a.m. and 5 p.m. local time). Once an 

apoapsis was designated as available for a maneuver, the daily analysis was initiated. The Flight 

Mechanics Team would consider the effect of each maneuver menu option on the propagated 
NASA Engineering and Safety Center

Technical Assessment Report

Document #:

NESC-RP09-00605

Version:

1.0

Title:

Autonomous Aerobraking (Phase 2)

Page #:

69 of 97

NESC Request No.: TI-09-00605

trajectory over the next 48 hours. The maneuver decision was often straightforward, as in the 

case when the no maneuver option kept the spacecraft within the corridor. Other times, the no 

maneuver case would fall outside the corridor on one or more orbits, at which time the team had 

to evaluate the effectiveness of the overall aerobraking strategy, including: 

1. How well the mission was following the reference glide slope; if the mission is

running behind, the strategy may be to perform maneuvers to keep the maximum heat

rates near the top of the corridor.

2. How the orbital mechanics of the trajectory impact periapsis altitude on subsequent

orbits; as the periapsis latitudes dropped, periapsis altitudes increase (due to the

Martian gravity field for the latitudes flown in the ORT) resulting in the desire to

remain at the lower part of the corridor because natural precession would increase

heating on subsequent orbits.

3. Number of corridor violations; if only one orbit exceeded a corridor did it warrant

using a smaller/larger maneuver?

4. Looking ahead; what would the situation look like tomorrow? Should a small

maneuver today be traded for a larger maneuver tomorrow (or vice versa)?

All of these criteria were used to determine whether or not a maneuver would be made. In the 

end, ORT1 opted to perform 22 maneuvers requiring 4.8 m/s 'V. The solid green circles in 

Figure 7.6-6 denote the orbits where maneuvers were commanded. The total ORT1 mission 

duration was 70.4 days. The results of the decisions made for ORT1 are compared to ORT2 in 

Section 7.6.1.5. 

7.6.1.4 ORT2: AADS with the SP

Because AA is only intended for use in the main phase of aerobraking, and because the AE

requires a data archive used to make atmosphere predictions, the ORT2 AADS simulation begins 

at orbit 14 of the reference mission, approximately 10 days after the first walk-in orbit. 

The AADS eliminates the need for nearly all of the teams listed in Table 7.6-1 because the 

software moves the maneuver decision-making process onboard the spacecraft. However, while 

AADS eliminates the need for all of the daily analysis, to evaluate mission margin, the Flight 

Mechanics Team still performs the weekly analysis on the ground to determine whether corridor 

adjustments are required to meet the overall aerobraking constraints. The weekly analysis occurs 

in the same manner as described for ORT1. The Atmosphere Team provides an update to the 

model A0 based on the comparison of the observed and modeled densities from the previous 

week and considerations for the week ahead. The updated A0, together with the spacecraft 

apoapsis state at the end of the week, is used to initiate three simulations — one each targeting 

the heat rate of the upper, middle, and lower corridor. The results are used to determine what 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

70 of 97 

NESC Request No.: TI-09-00605 

changes, if any, should be made to the corridor over the next week to keep the mission on track 

(i.e., with respect to the glide slope). 

Due to the similarity with ORT1, week 3 run-out analysis, shown in Figures 7.6-4 and 7.6-5, 

plots are not provided for ORT2 at the same point in the mission. As with ORT1, ORT2 analysis 

at the end of week 3 indicated that the lower corridor could be raised slightly to speed up the 

mission. Eventually, starting after week 4, the upper corridor was also lowered, effectively 

increasing the margin to the immediate action line, a move contrary to the decisions made in 

ORT1. 

Figure 7.6-7 shows the heat rate corridor plot for the ORT2 simulated mission. It is noted that 

the heat rate on orbit 39 and 48 exceeded the immediate action limit line. In ORT2, maneuvers 

were commanded based only on the predicted value of the heat rate at the next periapsis (open 

red circles in the plot). If the heat rate was within the corridor, no maneuver was performed, and 

if outside the corridor a maneuver was performed to place the next periapsis heat rate at the 

specified point within the corridor. As a note, ODY also had one orbit that reached the 

immediate action line. The maneuver logic and pop-up capability described in Section 7.1.5 are 

fully operational. However, the maneuver logic based on the 3-sigma high-density prediction 

was not active in the simulation, but the 3-sigma heat rates were calculated and are shown in 

Figure 7.6-8. 

Figure 7.6-7. ORT2 Heat Rate Corridor 
NASA Engineering and Safety Center

Technical Assessment Report

Document #:

NESC-RP09-00605

Version:

1.0

Title:

Autonomous Aerobraking (Phase 2)

Page #:

71 of 97

NESC Request No.: TI-09-00605

Figure 7.6.8. ORT2 with the 3-Sigma High Heat Rate Predictions 

For a region of the atmosphere with considerable uncertainty, between orbits 40 and 60, the 

3-sigma high-predicted heat rates are well above the immediate action line. Therefore, ORT2

was re-run using the maneuver logic, which considered the 3-sigma values (ORT2a). The 

mission was completed utilizing a similar corridor profile, while successfully protecting against 

the immediate action limit violations. The corridor selections in ORT2a are similar to those used 

in the original ORT2, but the overall mission profile is different because of the additional biasing 

maneuvers. A plot of the heat rates observed for ORT2a are shown with the corridors and 

predicted heat rates in Figure 7.6-9. The figure also denotes the orbits for which the AADS 

commanded a maneuver. Also, it should be noted that for this analysis (and all subsequent 

analyses), the pop-up logic was enabled, but no pop-up maneuvers were ever determined to be 

necessary. 

It is important to note that while ORT1 experienced periapsis heat rates that exceeded the 

immediate action limit near the EOM due to the ability to only make a maneuver once per day, 

the AADS ORTs did not because a maneuver can be made each orbit. AADS can maintain 

periapsis heat rates closer to the corridor limits in similar atmosphere conditions. This is another 

way AADS reduces mission risk. 
NASA Engineering and Safety Center

Technical Assessment Report

Document #:

NESC-RP09-00605

Version:

1.0

Title:

Autonomous Aerobraking (Phase 2)

Page #:

72 of 97

NESC Request No.: TI-09-00605

Figure 7.6-9. ORT2a with the Density 3-Sigma Correction 

Finally, for comparison to the AADS ORT2 and ORT2a, which modified the corridor throughout 

the mission, a third AADS ORT simulation was considered that kept the upper and lower 

corridor constant for the entire mission (ORT2b). The corridor was set to the values determined 

from the operational mission run-out (0.085 and 0.14 W/cm2

). With the 3-sigma density

maneuver logic enabled, ORT2b also had no immediate action violations. A full ORT mission 

comparison is provided in the next section. 

For comparison, a summary of the ORT1 and ORT2 process is shown in Figure 7.6-10.
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

73 of 97 

NESC Request No.: TI-09-00605 

Figure 7.6-10. Data Flowchart for the Two Types of ORTs; EOM 

7.6.1.5 ORT Performance Comparison 

As was expected, the AADS ORT (ORT2) required much less ground support (both in terms of 

time and effort) and ended with a LTST closer to the reference mission than the ground-based 

ORT (ORT1). However, the required amount of ground support is a qualitative metric and is 

difficult to measure in the AA Phase 2 simulated and condensed ORT environment. Therefore, 

quantitative metrics are identified to compare the overall aerobraking mission performance 

between the ORTs. 
NASA Engineering and Safety Center

Technical Assessment Report

Document #:

NESC-RP09-00605

Version:

1.0

Title:

Autonomous Aerobraking (Phase 2)

Page #:

74 of 97

NESC Request No.: TI-09-00605

Primary metrics used for comparison of the operational approaches included: 

1. Total maneuver 'V required

2. Total number of maneuvers

3. Flight corridors

4. Ability to match reference mission end conditions (e.g., LTST)

Table 7.6.2 compares key aerobraking parameters in each ORT. In ORT2a, which includes the 

density 3-sigma uncertainties in the maneuver logic, 23 more maneuvers and 75 percent more 

'V (15.3 versus 8.75 m/s) were required than for ORT2, however, no immediate action 

violations occurred. The corridor selections in ORT2a are similar to those used in the original 

ORT2, but the overall mission profile is different because of the additional biasing maneuvers. 

Table 7.6-2. ORT Performance Metrics 

Figure 7.6-11 shows the variations in the flight corridors throughout the aerobraking mission for 

each ORT. ORT1, ORT2, and ORT2a modified corridors weekly based on the weekly run-out 

analysis and were adjusted to keep the mission on track with the reference glide slope to end at 

the desired LTST. ORT2b maintained a constant corridor throughout the mission 

ORT2b required fewer maneuvers and used less 'V than either of the other two AADS mission 

simulations, while still only requiring weekly ephemeris updates. It should be noted here that 

even with a fixed corridor, AADS still meets the final LTST requirements (i.e., target 

+/- 15 minutes). What is lost in the case of flying a fixed corridor for the entire mission is the 

ability to catch up in the event that the spacecraft falls behind, or provides more margin if the 

spacecraft has become ahead. 

Simulation Number of 

Maneuvers 

Maneuver 

'V (m/s) 

Duration 

(days) 

Final LTST 

ORT1 – Ground 22 4.8 70.4 10:39

ORT2 – AADS 33 8.75 69.9 10:43

ORT2a – AADS 

(3 sig atm) 

56 15.3 67.3 10:49

ORT2b – AADS 

(const corridor) 

26 11.1 65.5 10.51
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

75 of 97 

NESC Request No.: TI-09-00605 

Figure 7.6-11. ORT Corridors 

Thus, AADS provides both better control of the final LTST and more flexibility primarily 

because a maneuver can be performed every orbit. 
NASA Engineering and Safety Center

Technical Assessment Report

Document #:

NESC-RP09-00605

Version:

1.0

Title:

Autonomous Aerobraking (Phase 2)

Page #:

76 of 97

NESC Request No.: TI-09-00605

Plots of the glide slope for the reference mission, ORT1 and ORT2 are shown in Figure 7.6-12.

ORT2 was able to follow the reference glide slope more closely than the ground-based ORT1. 

Both ORTs start off a little behind the glide slope (on the right side of the reference mission), but 

AADS has more opportunities to take corrective action to get back on track with significantly 

less analysis and enabled several other corridor and maneuver logic scenarios to be tested. 

Figure 7.6-12. ORT Glide Slope Comparison to Reference 

Figure 7.6-13 provides a comparison of the local time of the equatorial node crossing. As 

mentioned, the target was to end within 15 minutes of the reference mission. All ORT 

simulations were able to meet the requirement. Finally a plot of the local time versus orbit 

period for each ORT compared to the reference is provided in Figure 7.6-14.

The analyses shown here represent only the beginning of understanding the full set of AADS 

capabilities and its fundamentally different approach to aerobraking when the bulk of the 

maneuver decision-making is moved onboard the spacecraft. 
NASA Engineering and Safety Center

Technical Assessment Report

Document #:

NESC-RP09-00605

Version:

1.0

Title:

Autonomous Aerobraking (Phase 2)

Page #:

77 of 97

NESC Request No.: TI-09-00605

Figure 7.6-13. LTST of the Ascending Node Comparison to Reference 

Figure 7.6-14. ORT LTST versus Orbit Period Comparison to Reference 
NASA Engineering and Safety Center

Technical Assessment Report

Document #:

NESC-RP09-00605

Version:

1.0

Title:

Autonomous Aerobraking (Phase 2)

Page #:

78 of 97

NESC Request No.: TI-09-00605

Dust Storm Stress Case 

In addition to the ODY type mission ORTs, a stress case “mini” ORT was performed to evaluate 

the AADS’s ability to accommodate a dust storm. The MGS mission encountered a global dust 

storm early in its aerobraking mission. Because all the flight reconstructed density profiles were 

available, both ground-based and AADS “mini” ORTs were run using MGS atmosphere profiles 

corresponding to MGS orbits 50 to 70. The ground-based operations did experience some 

increase in workload due to higher variability in the day-to-day aerobraking performance.

Compared to ORT1, there was increased difficulty in correctly predicting and reacting to large 

differences to maintain glide slope while trying to preserve spacecraft safety margin. However,

there was almost no effect to the ground analysis or flight observed for the AADS operations due 

to its capabilities to perform maneuvers to remain in the corridor. 

7.6.1.6 Lessons Learned/Bug Fixes 

There were many lessons learned throughout the week the ORT took place. An assessment of 

the overall workload revealed that the ground-based operations required constant spacecraft 

monitoring and analysis. During the ORT, this task required 3 to 5 people. In contrast, the 

AADS simulation required much less staff support (one person), yet yielded better final mission 

conditions, namely the LTST of the ascending node crossing, at the end of aerobraking, while 

not requiring the thermal margin for the spacecraft to be reduced to compensate for falling 

behind. The primary reason the AADS mission minimizes the chances of falling behind is that 

the AADS can support a maneuver on every orbit, while the ground-based analysis could only 

perform a maneuver that would nominally keep the spacecraft safe for 48 hours (maximum being 

once daily). 

Table 7.6-3 lists the issues identified in one or both of the ORTs during the entire process of 

planning and executing these ORT analyses as a way to evaluate the AADS performance and 

robustness. This table also includes the impact and resolution for each issue. It should be noted 

that all of the issues listed here were found to exist in the ORT analysis setup, truth or spacecraft 

simulation, or in external processing scripts and not within the AADS itself. This greatly 

increases the confidence in the design, implementation and interfaces of the AADS system, and 

indicates it is viable for future development. 
NASA Engineering and Safety Center

Technical Assessment Report

Document #:

NESC-RP09-00605

Version:

1.0

Title:

Autonomous Aerobraking (Phase 2)

Page #:

79 of 97

NESC Request No.: TI-09-00605

Table 7.6-3. Issues Related to ORT Simulation and their Impacts and Resolutions 

Issue Effect Impact Rationale Resolution 

ODY atmosphere profiles 

were not being 

incremented due to error 

in table indexing

MAJOR BOTH Resulted in same atmosphere 

profile being used for each 

orbit, eliminating desired orbitto-orbit variability 

Identified during first 

week of ORT; was 

corrected and the ORT 

analyses restarted; walkin trajectory was NOT 

corrected 

20 orbit offset between 

ORT simulation pass 

number and ODY 

atmosphere profile 

(e.g., ORT pass #87 used 

ODY orbit 107 atmosphere 

profile) 

MINOR BOTH Orbit-to-orbit variability is still 

preserved

No resolution was 

necessary for ORT; 

correction was 

implemented for ODY 

"re-flight" analyses 

Atmosphere profiles were 

being changed at periapsis 

MAJOR AADS Resulted in a different ODY 

profile being used for pre and 

post periapsis during any single 

drag pass; introduced 

atmosphere discontinuity at 

periapsis

Identified at end of ORT 

day 1; was corrected and 

AADS ORT mission and 

analyses from day 1 were 

re-run on day 2 

Tolerance too small to 

allow for proper 

convergence of 

optimization problem in 

mission run-out 

simulations (mid-to-small 

period orbits) 

MINOR BOTH Impact only to mission run-out 

analyses; issue was identified 

and corrected prior to the 

point in the ORT where any 

possibly corrupt data would 

have been used for weekly 

reset decisions 

Tolerance was increased 

to allow for better 

convergence; this was 

done mid-ORT; re-runs of 

prior mission run-out 

analyses were not 

deemed necessary 

Using time since periapsis 

instead of time to periapsis 

at start of drag pass to 

query ODY atmosphere 

profile

MINOR AADS Impact only to very beginning 

of drag pass (~200 km altitude) 

where the atmosphere profile 

table begins; result was 

incorrect interpolation of very 

small density; did not likely 

impact acceleration data into 

AADS, which begins/ends at a 

lower altitude (~180 km)

Required code change to 

AADS simulation (truth) 

was incorporated midORT; re-run of prior 

analyses were not 

deemed necessary due to 

insignificant effect on 

truth trajectory 
NASA Engineering and Safety Center

Technical Assessment Report

Document #:

NESC-RP09-00605

Version:

1.0

Title:

Autonomous Aerobraking (Phase 2)

Page #:

80 of 97

NESC Request No.: TI-09-00605

Issue Effect Impact Rationale Resolution 

MarsGRAM inputs in 

target.inp inconsistent 

with those in 

msn_runout.inp, which 

resulted in different 

atmospheres being used in 

the optimization problem 

of the mission run-out 

analyses when the 

periapsis altitude became 

small (near EOM for low 

period orbits) 

MINOR BOTH Impact only to mission run-out 

analyses and reference 

mission; issue was identified 

and corrected prior to the 

point in the ORT where any 

possibly corrupt data would 

have been used for weekly 

reset decisions

Inputs corrected midORT; re-runs of prior 

mission run-out analyses 

were not deemed 

necessary; impact on 

reference mission 

expected to be small so 

no update made during 

ORT; reference updated 

for relevant post-ORT and 

reflight analyses

Tracked instantaneous 

heat rate instead of a 

running mean to 

determine maximum 

spacecraft thermal 

environment during a drag 

pass

NEUTRAL BOTH Due to the high "noise" level in 

the atmosphere data, the 

maximum heat rate based on 

the instantaneous density 

likely does not reflect the true 

spacecraft thermal 

environment; only impact was 

on the reporting of truth 

environments during weekly 

resets; since thermal limit 

violations were not significant, 

the impact was minimal 

Simulations and postprocessing scripts were 

updated to track the 

running mean heat rate 

to determine "true" 

spacecraft environment; 

implemented and used 

for all relevant post-ORT 

and reflight analyses

Atmosphere profiles 

misaligned with periapsis 

numbers; it was assumed 

that the atmosphere 

profile numbers were 

synced to the 

periapsis/orbit number 

when in fact it appears 

they are synced to the 

orbit number at the start 

of the pass 

MINOR BOTH Impact on ORT analyses is 

insignificant; desired ODY 

orbit-to-orbit variability was 

preserved

Issue corrected for MRO 

reflight analyses; ODY 

reflight analyses were not 

correct since any impact 

on those results are 

assumed to be minor
NASA Engineering and Safety Center

Technical Assessment Report

Document #:

NESC-RP09-00605

Version:

1.0

Title:

Autonomous Aerobraking (Phase 2)

Page #:

81 of 97

NESC Request No.: TI-09-00605

Issue Effect Impact Rationale Resolution 

startup atmosphere 

archive for AADS (AE) 

indexing error for density 

"predictions" of next orbit 

NEUTRAL AADS Results in uncharacteristically 

high density sigma estimates 

when using the startup archive 

data (first 7 orbits of AADS 

operation; the effect is largest 

on orbit 1, minimal on orbit 7, 

and non-existent orbits ≥8); 

only impact is an increased 

possibility of commanding a 

bias maneuver when 

unnecessary (occurred only 

once in post-ORT analyses) 

Startup atmosphere 

archive script updated for 

reflight analyses

Additional lessons learned included the need to establish a precise orbit number to stop the SFS. 

For ORT1, the code was set up to stop at the end of every 24-hour period, then wait to see if a 

maneuver would occur or not and if so, define the specified orbit. However, in hindsight, it

would have been better to have stopped the simulation on an orbit number, a rather 

straightforward modification. This would have allowed for fewer maneuver options considered 

in the ABM sweep and likely implemented. 

Also, after performing the ORT, as was found during ODY operations, having more options in 

the 'V maneuver menu would have allowed more flexibility in selecting the most efficient 

maneuvers to remain in the corridor. 

7.6.2 Historic Aerobraking Missions Replicated 

In addition to the ORT analysis, a POST2 simulation was developed to evaluate the ability of 

AADS to replicate past missions that used ground-based operations. The objective was to 

compare the operational decisions that would have been made by AADS to the decisions made 

by the actual flight team. Two simulations were established, one for ODY and another for MRO. 

These aerobraking missions had different mission designs. ODY was an aggressive mission in 

that to meet the science requirements it had to fly with only a 100 percent heat rate margin 

(during the mission, this was reduced to 80 percent because it was flying too far behind the 

desired glideslope). MRO was a conservative mission with a heat rate margin of 250 percent. 

Another difference was that ODY encountered an atmosphere density that was ~20 percent of the 

nominal predicted whereas MRO encountered an atmosphere density that was ~500 percent of 

the nominal predicted. Thus, ODY had to fly lower periapsis than predicted and MRO flew 

higher periapsis than predicted. 
NASA Engineering and Safety Center

Technical Assessment Report

Document #:

NESC-RP09-00605

Version:

1.0

Title:

Autonomous Aerobraking (Phase 2)

Page #:

82 of 97

NESC Request No.: TI-09-00605

Each mission utilized the respective aerobraking initial state, spacecraft mass prosperities and 

aerodynamics, and observed atmospheres from the actual mission. For simplification, the actual 

flight corridors were assumed to be constant for the analyses presented here, although additional 

cases were run to illustrate the AADS optimization capability demonstrated during the ORT 

analysis. The initial state was taken prior to the start of the walk-in phase of each respective 

mission. The walk-in phase was then manually duplicated (ensuring both the periapsis altitudes 

and heat rates were consistent between the actual mission and the AADS simulation) prior to the 

start of the main aerobraking phase, where the AADS is utilized. It is also noted that, since 

AADS is not intended for use in the last phases of aerobraking (endgame and walk-out where 

both COLA and spacecraft lifetime operations take precedence), both simulations are terminated 

for this analysis at an apoapsis altitude of ~2800 km rather than the required 400 km final 

apoapsis altitude. The comparisons of metrics (e.g., number of maneuvers, 'V, mission 

duration) only include the historical data to the same point in the mission. The following 

sections describe the results of the AADS to replicate the ODY and MRO missions. 

7.6.2.1 AADS ODY Aerobraking Mission Replication 

The ODY aerobraking mission, following Mars orbit insertion on October 24, 1999, was the 

most aggressive aerobraking mission to date, meaning it had the lowest amount of margin to its 

thermal limits. ODY utilized a non-constant heat rate corridor that was initially set with an 

upper limit of 0.32 W/cm2

 and a lower corridor of 0.14 W/cm2. The immediate action line was

0.54 W/cm2. This corridor reduced the 18-hour initial orbit period to 2 hours over 77 days.

Figure 7.6-15 shows the maximum heat rates observed each orbit of the ODY mission. To

simplify the AADS evaluation, the ODY initial corridor values were held constant in the 

simulation. Figure 7.6-16 shows the results AADS following a similar corridor. As mentioned, 

the AADS is only intended for use during the main phase of aerobraking and is terminated early 

in the endgame near orbit 210. Table 7.6-4 provides a comparison of the AADS flight metrics to 

the actual mission values to that point in the mission. The AADS simulation required 10 more 

maneuvers, but used slightly less 'V and ended only a fraction of a day earlier while requiring 

updates only weekly compared to the ground operations that occurred daily. The simulation 

verified that AADS was able to meet aerobraking requirements through the main phase. 
NASA Engineering and Safety Center

Technical Assessment Report

Document #:

NESC-RP09-00605

Version:

1.0

Title:

Autonomous Aerobraking (Phase 2)

Page #:

83 of 97

NESC Request No.: TI-09-00605

Figure 7.6-15. ODY Aerobraking Maximum Heat Rate [ref. 4] 

Table 7.6-4. Comparison of ODY Flight and AADS Replicated Missions 

Simulation Number of 

Maneuvers 

Maneuver 'V 

(m/s) 

Duration 

(days) 

Change in final LTST 

of node (min) 

ODY Flight 25 18.84 68.8 -- 

AADS 

fixed corridor 

35 17.6 68.0 8.0

AADS 

small corridor 

173 32.5 64.5 -0.33 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

84 of 97 

NESC Request No.: TI-09-00605 

Figure 7.6-16. AADS ODY Aerobraking Mission Run-out Maximum Heat Rate 

The AADS mission matched closely to the aerobraking duration and 'V, but it missed the 

desired target LTST of the descending node by about 8 minutes (requirement was 

15 minutes) and is shown in Figure 7.6-17. During the ODY aerobraking phase, ODY fell 

behind the desired glide slope to the point that it might not have met the required LTST 

requirement. ODY flight managers reduced the margin to 80 percent to allow ODY to fly higher 

heat rates to catch up. With AADS, the original margin could be maintained by raising the lower 

corridor, which would result in ODY flying higher in the corridor at the expense of more 

maneuvers. To demonstrate this, a second simulation was run with the corridor collapsed to 

always command the middle of the original corridor at 0.23 W/cm2

. Targeting the middle of the 

original corridor met the science requirements, but at the cost of 150 more maneuvers and twice 

the 'V. Thus, if the spacecraft was using AADS and got behind (e.g., if there was a malfunction 

and the spacecraft had to increase its periapsis altitude while the problem was being resolved) 

AADS could be configured to fly higher in the corridor and catch up without having to reduce 

the thermal margin. For reference, Figures 7.6-18 and 7.6-19 show how the two AADS missions 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

85 of 97 

NESC Request No.: TI-09-00605 

compared to the actual ODY flight with regard to periapsis latitude, altitude, orbit period, and 

LTST. 

Figure 7.6-17. AADS ODY Narrow Corridor Simulation Peak Heat Rates 
NASA Engineering and Safety Center

Technical Assessment Report

Document #:

NESC-RP09-00605

Version:

1.0

Title:

Autonomous Aerobraking (Phase 2)

Page #:

86 of 97

NESC Request No.: TI-09-00605

Figure 7.6-18. ODY Mission versus AADS Glide Slope Comparison 

Figure 7.6-19. ODY Mission Compared to AADS Two Replication Simulations 
NASA Engineering and Safety Center

Technical Assessment Report

Document #:

NESC-RP09-00605

Version:

1.0

Title:

Autonomous Aerobraking (Phase 2)

Page #:

87 of 97

NESC Request No.: TI-09-00605

7.6.2.2 AADS MRO Aerobraking Replication 

The much less aggressive MRO aerobraking mission began on the 24th orbit past Mars orbit

insertion on March 10, 2006. MRO had much lower corridor constraints with an average upper 

limit of 0.18 W/cm2 and approximate lower corridor of 0.11 W/cm2

. The immediate action line

was near 0.4 W/cm2. Figure 7.6-20 shows the maximum heat rates observed on each orbit of the

MRO mission. Again, to simplify the AADS evaluation, the initial corridor values were held 

constant in the simulation. Figure 7.6-21 shows the results AADS following a similar corridor. 

Table 7.6-5 provides a comparison of the AADS flight metrics to the actual mission values. The 

AADS simulation required 18 more maneuvers requiring 8.4 m/s 'V and ended 10 days later 

than the original mission. As was shown for the ODY mission, the AADS was able to meet the 

MRO aerobraking mission requirements using spacecraft updates weekly compared to the daily 

ground operations that were required for the actual mission. See Figures 7.6-22 and 7.6-23. 

Table 7.6-5. Comparison of MRO Flight and AADS Replicated Mission 

Figure 7.6-20. MRO Aerobraking Peak Heat Rates [ref. 5] 

Simulation Number of 

Maneuvers 

Maneuver 'V 

(m/s) 

Duration 

(days) 

Change in final LTST of 

node (min) 

MRO Flight 24 16.0 157.2 -- 

AADS 42 24.4 161.0 11.3 
NASA Engineering and Safety Center

Technical Assessment Report

Document #:

NESC-RP09-00605

Version:

1.0

Title:

Autonomous Aerobraking (Phase 2)

Page #:

88 of 97

NESC Request No.: TI-09-00605

Figure 7.6-21. AADS MRO Aerobraking Mission Run-out Maximum Heat Rate 

Figure 7.6-22. MRO versus AADS Glide Slope Comparison 
NASA Engineering and Safety Center

Technical Assessment Report

Document #:

NESC-RP09-00605

Version:

1.0

Title:

Autonomous Aerobraking (Phase 2)

Page #:

89 of 97

NESC Request No.: TI-09-00605

Figure 7.6-23. MRO versus AADS Aerobraking Mission Parameter Comparison 

The purpose of evaluating the various AADS operational scenarios was to understand the 

features and capabilities of the software. The cases examined for the AA Phase 2 ORTs are the 

beginning to understanding the flexibility and capability of AADS. By implementing the 

biasing, the number of maneuvers increased (ORT2a) and relaxing the corridor by keeping it 

constant throughout the mission (ORT2b) the mission used fewer maneuvers with a shorter 

aerobraking phase. Different options could be considered, such as only using the biasing during 

the periods with the most atmosphere variability. Yet, none of the stress cases resulted in a 

modification of the AADS code only in how it was implemented in the mission. The ORT 

demonstrated the power of having the onboard maneuver capability. AADS is the logical 

extension of the current periapsis timing estimator implemented on ODY and demonstrated on 

MRO and is a means to enable a new era of planetary exploration in the solar system. 

7.7 Open Phase 2 Investigation Topics 

The AA Development Team completed an extraordinary amount of work during Phase 2. 

Simulations that were built in Phase 1 were developed and improved during Phase 2. Many 

trade studies were performed; nominal, off-nominal, and stressing aerobraking scenarios were 

analyzed, and Monte Carlo performance was assessed. In addition, the team developed and 
NASA Engineering and Safety Center

Technical Assessment Report

Document #:

NESC-RP09-00605

Version:

1.0

Title:

Autonomous Aerobraking (Phase 2)

Page #:

90 of 97

NESC Request No.: TI-09-00605

executed an ORT of an aerobraking scenario with a stressing environment and recreated ODYlike and MRO-like aerobraking missions. AADS performed well under all circumstances. There 

are, however, areas of continued study that could be accomplished during a Phase 3 task. These 

items are described below. 

Phase 2 focused on the development of AA at Mars. Although some work was performed in the 

area of Venus atmospheric uncertainty analysis [ref. 7], few simulations were performed to study 

the effect of a stressing environment at Venus. Given additional time and resources, trade 

studies and Monte Carlo assessments could be performed for a Venus-orbiting spacecraft. In 

addition, an ORT-like scenario could be developed if flight-like Venus atmosphere profiles could 

be produced. Additional studies at Titan could be considered for completeness, but it is a less 

probable destination for an aerobraking orbiter. 

AADS was improved in Phase 2 by the architecture and algorithm changes to the orbit 

integration function (formerly the EE, now the SP), and modifications to the AE and ME 

algorithms. Along with these algorithm changes, the software has been streamlined with the 

EML implementation, which further reduces the computational overhead and the memory 

requirements for AADS. The EML version also moves the code closer to a version suitable for 

implementation on a flight processor, although there is still much work to be done in this area. 

An error detection strategy, exception handling strategy, and a fault management strategy need to 

be defined. Some will be mission-dependent, but the software needs a mechanism for reporting 

errors and possibly responding to them. 

Another issue identified in Phase 2 was the characterization of the uncertainty of the control 

variable. When the heat rate is used for the control variable, the AADS computes the uncertainty 

(variance) associated with the atmosphere model fit to help inform the maneuver calculation 

algorithm. This variance can be used to decide whether or not to perform a second category of 

maneuver, which might help further ensure spacecraft safety (instead of simply attempting to 

follow the designed glide slope of the reference mission). While it is largely believed that the 

natural variation in the atmosphere is the dominant source of uncertainty in the corridor control 

variable, there are other uncertainty sources that should be characterized for a more thorough 

understanding of the robustness of AA. Of particular importance are the uncertainties due to the 

orbit propagation. Because this process is only reset via ground interaction, this largely 

determines the duration between ground updates. Initial state uncertainty (in the ground-based 

navigation solution) and uncertainties in the accelerometer data are generally small, but grow 

large through the long orbit propagation arcs. This can lead to uncertainty in periapsis altitude 

and timing predictions which, in turn, affect the atmosphere estimation and corridor control 

variable predictions. 

A final issue identified in Phase 2 is related to the ground update. In current testing, the 

simulations use a current state provided to the spacecraft to reset the onboard ephemeris 

knowledge. This instantaneous state update is unrealistic as there would likely be substantial 
NASA Engineering and Safety Center

Technical Assessment Report

Document #:

NESC-RP09-00605

Version:

1.0

Title:

Autonomous Aerobraking (Phase 2)

Page #:

91 of 97

NESC Request No.: TI-09-00605

time between the ground computation of the orbit and the ability to load a new ephemeris state to 

the spacecraft. This latency may mean that several drag passes have transpired since the end of 

the ground tracking data arc, which degrades the orbit knowledge considerably, as the ground 

does not have an accurate atmosphere model for propagation. In this case, the spacecraft has 

better information about the recent past than the ground. In order for the ground to provide the 

spacecraft the most accurate state, this state may have to be several orbits old, which then 

requires the spacecraft to propagate this delayed state measurement up to the current time. This

propagation may span numerous drag passes that require some method of storage or 

characterization to preserve the accuracy in the ground update. This problem is currently under 

investigation by the team, but will be challenging to solve without introducing substantial code 

changes or new concepts of operations. 

Even with these identified studies to improve performance of AADS, Phase 2 provided key 

enhancements to the AADS and answered numerous important questions. In addition to the 

tasks assigned for Phase 2, an attempt was initiated to recreate the AADS ORT within the 

AAHFS. This was beyond the scope of the initial task, but much progress was made in 

performing this analysis. 

A significant objective of Phase 2 was to provide two working versions of AADS: one using a 

SP developed by Applied Physics Laboratory (APL), the second using a version of the JPLdeveloped AeroNav. All of the analysis shown in this final report was using the SP version of 

AADS. Work was performed in developing and delivering AeroNav to the AADS 

implementation team, but following a unit test of the performance of AeroNav, there was not 

enough time left in Phase 2 to complete the integration of AeroNav into AADS and run the 

analyses. Phase 2 was successful in completing a broad suite of studies with AADS-SP. The 

AeroNav version of the AADS was desired for ease of implementation within a robotic JPL-led 

orbiter mission that would already have been using AutoNav. It is still desired to maintain this 

option, and with further study, AeroNav could be implemented and tested with the same rigor 

that was given to the AADS-SP. 

Any further trade studies that have been identified for the AADS (e.g., 'V sizing, maneuver 

magnitude limitations, maneuver execution performance sensitivities, etc.) will be left for 

implementation for a specific mission. These small studies are too spacecraft-specific to provide 

much benefit for the general AA analysis performed. 
NASA Engineering and Safety Center

Technical Assessment Report

Document #:

NESC-RP09-00605

Version:

1.0

Title:

Autonomous Aerobraking (Phase 2)

Page #:

92 of 97

NESC Request No.: TI-09-00605

8.0 Findings, Observations, and NESC Recommendations

8.1 Findings 

The following findings were identified: 

F-1. The AADS performance at Mars was similar to that of a ground-based team, illustrated 

by the ORT and by recreation of the ODY and MRO aerobraking missions. 

F-2. The ORT demonstrated the daily work requirements utilized for current ground-based 

operations could be extended to every 7th day for AADS. 

F-3. The ORT demonstrated an ephemeris update only once per week is sufficient to maintain 

successful aerobraking with the AADS, which reduces the requirement for DSN from 

continuous coverage to a few hours per week. 

F-4. The ORT demonstrated that the aerobraking corridor could be reduced to require a 

maneuver every orbit with the SP maintaining the required accuracy. 

F-5. The ORT demonstrated the ability to execute the required maneuvers to continue 

aerobraking with the required margins when subjected to a simulated dust storm 

equivalent to that encountered by MGS. 

F-6. AA simulations demonstrated that the number of maneuvers could be traded for thermal 

margin. 

F-7. AA simulations replicated the ODY and MRO missions with no reductions in thermal 

margins. 

F-8. A unit test of the AeroNav orbit propagation showed similar results to AADS (i.e., state 

differences <10 seconds) when tested at two independent locations. 

F-9. The AADS simulations demonstrated robustness to reconstructed atmosphere 

characteristics using MRO, MGS, and ODY flight data. 

F-10. The AADS algorithms are feasible for an onboard implementation in terms of execution 

speed and memory requirements. 

F-11. Benchmarking of the AADS for onboard implementation demonstrated that 

computational frame requirements could be easily met with significant margin. 

F-12. The AADS internal architecture was revised to allow sequential processing of input 

sensor data, eliminating large data buffers. 

F-13. The number and frequency of the gravity model evaluations was dramatically reduced, 

reducing the computational burden on the flight processor of AADS. 

F-14. AADS SP errors were greatly improved over Phase 1. 
NASA Engineering and Safety Center

Technical Assessment Report

Document #:

NESC-RP09-00605

Version:

1.0

Title:

Autonomous Aerobraking (Phase 2)

Page #:

93 of 97

NESC Request No.: TI-09-00605

F-15. The fidelity of the onboard orbital integration was significantly improved by separating 

the integration of the fast dynamics (thrust and drag) from the slow dynamics (gravity 

model terms) and using a numerically efficient method for each integration. 

F-16. In IMU noise-free simulation, ground updates can be spaced greater than 28 days with no 

degradation in ephemeris performance. 

F-17. Corridor performance is not significantly affected by typical maneuver execution errors. 

8.2 Observations 

The following observations were identified: 

O-1. Maintaining two independent and synergistic simulations (POST2 and AAHFS) allowed 

for rapid, parallel trade studies performed in parallel, providing increased quantity of 

Phase 2 analyses. 

O-2. The EE AeroNav is not as mature as the AADS. 

O-3. AADS does not employ algorithms to address latency of ground-based ephemeris 

updates. 

O-4. Venus atmospheric uncertainties are substantially different than those calculated for Mars 

and may have significant impact in Venus aerobraking. 

8.3 NESC Recommendations 

The following NESC recommendations were identified and directed toward the NESC (including 

the technical disciplines of Flight Mechanics, Aerosciences, Passive Thermal, GN&C, Software, 

Loads and Dynamics, and Human Factors) and future NASA Programs and Projects that may 

utilize aerobraking: 

R-1. Future missions that require aerobraking should evaluate the application of AADS. 

(F-1 through F-7, F-9 through F-11)

R-2. The ORT simulation capability developed in the Phase 2 effort could be utilized by future 

aerobraking missions to offer mission designers, project managers and spacecraft 

operators a realistic flight mission rehearsal. (F-1 through F-7)

R-3. Further development of AA should address: 

a. Evaluate Venus AADS performance, including potential ORT analysis and offnominal testing. (F-1 through F-7, F-9, O-4)

b. Implement AeroNav in AADS, provide alternative algorithms for autonomous

aerobraking and perform an ORT. (F-8, O-2)
NASA Engineering and Safety Center

Technical Assessment Report

Document #:

NESC-RP09-00605

Version:

1.0

Title:

Autonomous Aerobraking (Phase 2)

Page #:

94 of 97

NESC Request No.: TI-09-00605

c. Evaluate and improve the suitability of AADS for flight processor

implementation, ensuring code robustness and proper error reporting.

(F-10 through F-17)

d. Develop AADS algorithms to address the latency of ground-based ephemeris

updates. (O-3)

9.0 Alternate Viewpoint

There were no alternate viewpoints identified during the course of this assessment by the NESC 

team or the NRB quorum. 

10.0 Other Deliverables

No unique hardware, software, or data packages, outside those contained in this report, were 

disseminated to other parties outside this assessment. 

11.0 Lessons Learned

No applicable lessons learned were identified for entry into the NASA Lessons Learned 

Information System. 

12.0 NASA Standards and Specifications

No recommendations for NASA standards and specifications were identified as a result of this 

assessment. 

13.0 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. 
NASA Engineering and Safety Center

Technical Assessment Report

Document #:

NESC-RP09-00605

Version:

1.0

Title:

Autonomous Aerobraking (Phase 2)

Page #:

95 of 97

NESC Request No.: TI-09-00605

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. 

14.0 Acronyms List

A0

AA

AADS

AAHFS 

ABM 

AE

AMA 

APL 

APP 

CCD

COLA 

CPU

DOE

DOF

'61

'V

MarsGRAM Density Multiplier 

Autonomous Aerobraking 

Autonomous Aerobraking Development Software 

Autonomous Aerobraking High Fidelity Simulation 

Aerobraking Maneuver 

Atmosphere Estimator 

Analytical Mechanics and Associates 

Applied Physics Laboratory 

Aerobraking Planning Period 

Central Composite Design

Collision and Avoidance 

Central Processor Unit 

Design of Experiments 

Degrees of Freedom 

'HHS6SDFH1HWZRUN

Delta Velocity 
NASA Engineering and Safety Center

Technical Assessment Report

Document #:

NESC-RP09-00605

Version:

1.0

Title:

Autonomous Aerobraking (Phase 2)

Page #:

96 of 97

NESC Request No.: TI-09-00605

EE Ephemeris Estimator 

EML Embedded MATLAB 

EOM End of Mission 

ET2000 Ephemeris Time from midnight January 1, 2000 

GN&C Guidance, Navigation, and Control 

GRAM Global Reference Atmosphere Model 

ICRF International Celestial Reference Frame

IMU Inertial Measurement Unit 

JPL Jet Propulsion Laboratory 

LaRC Langley Research Center 

LTST Local True Solar Time 

MarsGRAM Mars Global Reference Atmosphere Model 

MB megabyte 

ME Maneuver Estimator 

MHz Mega Hertz 

MRO Mars Reconnaissance Orbiter 

MTSO Management Technical Support Office 

N number 

NASA National Aeronautics and Space Administration 

NAV Navigation 

NCSU North Carolina State University 

NESC NASA Engineering and Safety Center 

NIA National Institute of Aerospace 

NRB NESC Review Board 

OD Orbit Determination 

ODY Mars Odyssey 

OP Orbit Propagation 

ORT Operational Readiness Test 

PDS Planetary Data System 

POST2 Program to Optimize Simulated Trajectories (v2) 

RKN Runge-Kutta-Nystrom 

RSE Response Surface Equation 

SD State Determination 

SFS Spacecraft Flight Simulation 

SP State Propagator 

TDT Technical Discipline Team 

TRL Technology Readiness Level 

VML Virtual Machine Language 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #:

NESC-RP09-00605 

Version:

1.0

Title: 

Autonomous Aerobraking (Phase 2) 

Page #:

97 of 97 

NESC Request No.: TI-09-00605 

15.0 References

1. You, T. et al. “Mars Reconnaissance Orbiter Navigation,” AIAA 2004-4745. 2004. 

2. Gladden, R. “Mars Reconnaissance Orbiter: Aerobraking Sequencing Operations and 

Lessons Learned,” AIAA 2008-3353. 2008. 

3. Murri, Daniel G. et al. “Development of Autonomous Aerobraking (Phase 1),” 

December 15, 2011. NESC-RP-09-00605/NASA-TM-2012-217328. 

4. Smith, J. C., “2001 Mars Odyssey Aerobraking” AIAA 2002-4532, 2002. 

5. Long, S. M., et al. “Mars Reconnaissance Orbiter Aerobraking Navigation Operation”, 

AIAA 2008-3349. 2008. 

6. Lyons, D. T., et al. “Mars Global Surveyor: Aerobraking Mission Overview,” Journal of 

Spacecraft and Rockets, Vol. 36, No. 3, May–June 1999. 

7. Tolson, R., Prince, J., Konopliv, A. “An Atmospheric Variability Model for Venus 

Aerobraking Missions.” Control ID#: 1654639: Technical Paper has been accepted for 

presentation at AIAA Modeling and Simulation Technologies Conference, Boston, MA, 

August 19-23, 2013. 
REPORT DOCUMENTATION PAGE Form Approved

OMB No. 0704-0188

2. REPORT TYPE 

Technical Memorandum

 4. TITLE AND SUBTITLE

Development of Autonomous Aerobraking

Phase 2

5a. CONTRACT NUMBER

 6. AUTHOR(S)

Murri, Daniel G.

 7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

NASA Langley Research Center

Hampton, VA 23681-2199

 9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Washington, DC 20546-0001

 8. PERFORMING ORGANIZATION

 REPORT NUMBER

L-20316 NESC-RP-09-00605

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

Phase 1 of the Development of Autonomous Aerobraking (AA) Assessment investigated the technical capability of 

transferring the processes of aerobraking maneuver (ABM) decision-making (currently performed on the ground by an 

extensive workforce and communicated to the spacecraft via the deep space network) to an efficient flight software algorithm 

onboard the spacecraft. This document describes Phase 2 of this study, which was a 12-month effort to improve and rigorously 

test the AA Development Software developed in Phase 1. 

15. SUBJECT TERMS

Aerobraking maneuver; Autonomous Aerobraking; Autonomous Aerobraking Development Software; Deep Space Network; 

NASA Engineering and Safety Center

18. NUMBER

 OF 

 PAGES

102

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)

December 2011 - June 2013

5b. GRANT NUMBER

5c. PROGRAM ELEMENT NUMBER

5d. PROJECT NUMBER

5e. TASK NUMBER

5f. WORK UNIT NUMBER

869021.05.07.01.14

11. SPONSOR/MONITOR'S REPORT

 NUMBER(S)

NASA/TM-2013-218032

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- 08 - 2013