www.nasa.gov 

2006 Near-Earth Object Survey and Deflection Study www.nasa.gov

These artist’s concepts of Tempel 1 simu-

late an optical view of the comet (left), next 

to the simulated infrared view (right). The 

images illustrate the comet’s shape, reflec- tivity, rotation rate and surface temperature, 

based on information from NASA’s Hubble 

Space Telescope and Spitzer Space 

Telescope. 

Measurements from the Great Observato-

ries indicate that the comet is a matte black 

object roughly 14 by 4 kilometers (8.7 by 

2.5 miles), or about one-half the size of 

Manhattan. 

At the time of these early observations, 

March 25-27, 2005,Tempel 1 was still far 

enough away from the Sun that it had not 

yet developed its characteristic halo of 

evaporating gas. 

Hubble and Spitzer observed the comet in 

visible and infrared light, respectively. The 

comet appeared only as an unresolved dot 

due to the great distance, but its general 

shape, size and color could be deduced 

from the way the visible and infrared bright-

ness varied over time. 

The animation simulates an optical view 

of the comet, followed by its appearance 

in infrared. Spitzer detects the comet’s 

infrared energy or heat, depicted by the 

reddish glow. The movement of the comet 

has been dramatically sped up to 13 sec-

onds per rotation versus its normal rotation 

of about 40 hours. As the comet slowly 

rotates, the sunlit side heats up while the 

dark side cools down. 

Image Credit: 

NASA/JPL-Caltech 
May 9, 2007 

Context and Errata to the NEO Study Analysis Report 

In the 2005 Budget Authorization Act, the U.S. Congress directed the NASA Administrator 

to provide an analysis of alternatives to detect, track, catalogue, and characterize potentially 

hazardous near-Earth objects (NEO). In addition, the legislation required the Administrator 

to submit an analysis of alternatives that NASA could employ to divert an object on a likely 

collision course with Earth. A study team derived requirements and figures of merit from the 

congressional direction, and used these factors to evaluate the alternatives. The team 

developed a range of options from public and private sources and then analyzed their 

capabilities, performance, life-cycle costs, schedules, and development and operations risks. 

This document collects the detailed results of those analyses and was prepared initially as a 

draft of the final report to Congress. 

As the study progressed, it became clear that integrating the full extent of the analysis 

material with programmatic assessments would be ambitious. In October 2006, a decision 

was made to prepare a consolidated report to Congress,1

 and work on this document stopped. 

Therefore, this report is an incomplete draft, and some facts and dates related to the 

publication of the final report itself are incorrect. In December of 2006, NASA made the 

decision to distribute printed copies of this document to NASA senior management and to the 

study team to replace older drafts that had circulated. 

Notes and Corrections 

• References to the date and disposition of the final report to Congress are incorrect. 

• This draft contains statements regarding planned assets of other agencies (NSF, USAF) 

that have not been coordinated with and do not represent the views of those agencies. 

• The costs presented in this document are “architecture trade costs”, appropriate for 

comparison but not for budgeting. 

• This document does not present programmatic options as required by Congress. 

• This document contains significant examples where analyses representing different 

strategies and points of view are presented together. These differences are not 

necessarily reconciled into a comprehensive message in this document. 

• This document refers to a “public workshop”, held in June 2006. This choice of 

nomenclature does not fully represent the structure of the meeting. Abstracts for 

presented papers were solicited from the public, but the meeting was by invitation only 

and contributors were not members of the study team. 

• The “estimated total” of NEOs that may be discovered in Figure 5 should be 10-15% less. 

The scientific community changed the process for estimating the size of NEOs during the 

study, and the figure was not fully updated to match. 

• The results in Section 6.17 and Figure 59 were generated using mixed assumptions for 

launch, orbit transfer, and threat size and the “order of magnitude” cost estimates may be 

incomplete. This figure will be modified if these analyses are republished. 

• A close approach by Apophis is reported as 2022 on page 30, it should be 2021. 

• The reference on page 50 should be to Dearborn [18] not Schweickart [17]. 

 

1

 NASA Near-Earth Object Survey and Deflection Analysis of Alternatives: Report to Congress. March 2007. 

Errata 0-1
May 9, 2007 

• The hypothetical threat Porthos has a diameter of 1 km rather than a radius of 1 km as 

reported in the middle of page 112. 

• The study team intended to differentiate between the terms “deflection alternative”, 

“deflection campaign”, “deflection strategy”, and “mitigation strategy”. In its current 

state, the document does not consistently use these terms and meaning must sometimes 

be drawn from context. These terms are described here for additional clarity. 

• A deflection alternative is a technological solution for deflecting the path of a 

potentially threatening object. These alternatives are the primary focus subject of the 

congressionally directed analysis of alternatives. These alternatives along with 

tracking assets form a “toolkit” for deflecting specific threats. 

• A deflection campaign is the combination of (potentially) multiple missions using 

(perhaps) multiple deflection options to divert a potential threat. Deflection 

campaigns for specific threats may vary widely in the how the alternatives are 

combined into an effective system. Campaigns were not a primary subject of study. 

• A deflection strategy may include activities that take place when no threat is known 

(such as now) including development of technologies, characterization, or “waiting 

for a credible threat”, and these strategies are not the subject of this document. 

• A mitigation strategy considers deflection options, costs, consequences, and other 

factors on how to respond to the general hazard of NEOs and under which 

circumstances to act. A mitigation strategy is formed considering the size-frequency 

curves, public will, available budget, cost-benefit, and many other factors. Mitigation 

strategies were not within the scope of this study. 

• The frequency of resonant returns and keyholes, and how well keyhole events could be 

predicted was a source of significant discussion during the study. While the study team 

found that keyholes represent less than 1% of potential impacts, others have estimated 

their frequency to be considerably higher. This difference does not alter the study 

findings, but could affect the development of deflection or mitigation strategies. 

• The deflection scenarios analyzed are expected to be representative of the range of 

potential threats. However, the scenarios were not intended to convey information 

regarding the predicted frequency of potential threats; this information is presented in 

Figure 2 and Table 1. Figures 37 and 38 may be used to evaluate the performance of the 

deflection options for the range of potential threat masses and deflection velocities (ΔV) 

required. 

• While the VD17 orbit may intersect keyholes in the future, it was not treated as such for 

these analyses. Scenarios were not meant to comprehensively represent actual threat 

scenarios, and keyholes are represented by the analysis of Apophis. Keyhole cases may 

be further analyzed by using Figures 37 and 38 for specific deflection performance 

requirements. 

Errata 0-2
DRAFT Pre-Decisional Material

2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material

, 

2006 Near-Earth Object 

Survey and Deflection Study 

December 28, 2006 

NASA HQ, PA&E 
DRAFT Pre-Decisional Material

2006 Near-Earth Object Survey and Deflection Study 

2 DRAFT Pre-Decisional Material 

For additional information, please contact the Office of Program Analysis and Evaluation (PA&E) at 

NASA Headquarters in Washington, D.C. 
DRAFT Pre-Decisional Material

2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 3 
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2006 Near-Earth Object Survey and Deflection Study 

4 DRAFT Pre-Decisional Material 

Table of Contents 

1 Executive Summary................................................................................................. 12

1.1. Survey Analysis of Alternatives ................................................................... 12 

1.2. Deflection Analysis of Alternatives.............................................................. 15 

1.3. Summary of Findings.................................................................................... 16 

1.4. Summary of Program Options ...................................................................... 17 

2 Participants .............................................................................................................. 18

2.1. Study Leadership .......................................................................................... 18 

2.2. Study Team ................................................................................................... 18 

2.3. Analyses........................................................................................................ 18 

2.4. Acknowledgements....................................................................................... 18 

3 Introduction and Overview .................................................................................... 20

3.1. Purpose.......................................................................................................... 20 

3.2. Statement of Need......................................................................................... 20 

3.3. Direction ....................................................................................................... 20 

4 Background .............................................................................................................. 22

4.1. Asteroids and Comets ................................................................................... 22 

4.2. Population of Near-Earth Objects................................................................. 23 

4.3. The Probability of Impact and Distribution of Sizes .................................... 25 

4.4. NASA Spaceguard Survey............................................................................ 26 

4.5. Small Near-Earth Object Survey Science Definition Team Report.............. 27 

5 Survey Program Analysis of Alternatives ............................................................. 30

5.1. Introduction................................................................................................... 30 

5.2. Derived Requirements and Definitions......................................................... 30 

5.2.1. Survey Goal .......................................................................................... 30 

5.2.2. Detect, Track, Catalogue, and Characterize.......................................... 30 

5.2.3. Resonant Returns and Keyholes ........................................................... 31 

5.3. Figures of Merit ............................................................................................ 32 

5.4. Development of Alternatives ........................................................................ 33 

5.4.1. Overview............................................................................................... 33 

5.4.2. Alternatives Definition Methodology................................................... 34 

5.4.3. Contributions of the Science Definition Team Report ......................... 34 

5.4.4. Alternatives Design Approach.............................................................. 35 

5.5. Alternative Approaches to Meeting Search Requirements........................... 35 

5.5.1. Ground-Based Optical Systems ............................................................ 35 

5.5.2. Space-Based Optical Systems............................................................... 35 

5.5.3. Space-Based Infrared Systems.............................................................. 36 

5.6. Detection and Tracking Survey Alternatives Considered............................. 36 

5.6.1. Survey Alternatives Not Analyzed and Rationale and Why................. 39 

5.6.2. Survey Alternatives Analyzed .............................................................. 39 

5.7. Survey System Elements Schedules ............................................................. 44 

5.8. Survey Performance Simulation Results ...................................................... 45 

5.9. Orbital Uncertainty ....................................................................................... 51 

5.10. Performance of Orbit Determination Alternatives........................................ 53 

5.10.1. Precision Orbit Determination Using Survey Assets............................ 53 
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5.10.2. The Role of Precovery in Precision Orbit Determination..................... 54 

5.10.3. Precision Orbit Determination Using Radar......................................... 55 

5.10.4. Precision Orbit Determination Using a Transponder In-Situ ............... 57 

5.11. Detection and Tracking System Element Costs............................................ 58 

5.12. Object Characterization Trade Space............................................................ 58 

5.13. Introduction to Characterization ................................................................... 60 

5.13.1. Object Characterization Parameters Useful for Deflection .................. 60 

5.13.2. The Role of Characterization Using Remote Sensing .......................... 60 

5.13.3. The Role of In-Situ Characterization Missions .................................... 61 

5.14. Object Characterization Strategy .................................................................. 62 

5.15. Characterization Option Architectures ......................................................... 63 

5.16. Characterization Development Schedules .................................................... 65 

5.17. Characterization Elements Cost Estimates ................................................... 65 

5.18. Performance and Cost................................................................................... 66 

5.19. Program Elements......................................................................................... 68 

5.20. Other Factors................................................................................................. 68 

5.20.1. Opportunity Science from the Survey and Characterization ................ 68 

5.20.2. Potential Benefits to Exploration.......................................................... 69 

5.20.3. Potential Synergies with Department of Defense ................................. 70 

5.21. Findings......................................................................................................... 70 

6 Analysis of Deflection Alternatives ........................................................................ 72

6.1. Introduction and Background ....................................................................... 72 

6.2. Definitions..................................................................................................... 73 

6.2.1. Deflection Campaigns........................................................................... 73 

6.2.2. Launch Energy and C3.......................................................................... 73 

6.2.3. Momentum Exchange Efficiency (beta, β)........................................... 73 

6.2.4. Specific Impulse (Isp)........................................................................... 73 

6.3. Derived Requirements .................................................................................. 73 

6.3.1. Derived Deflection Distance Requirement ........................................... 74 

6.3.2. Derived Reliability Requirement .......................................................... 74 

6.3.3. Derived Characterization Requirements............................................... 74 

6.3.4. Post-Action Effectiveness Assessment Requirements.......................... 75 

6.4. Deflecting a Potentially Hazardous Object................................................... 75 

6.5. Deflection Mission Timelines....................................................................... 76 

6.6. Figures of Merit ............................................................................................ 77 

6.7. Deflection Alternatives Trade Space ............................................................ 78 

6.7.1. Deflection Alternatives Trade Space .................................................... 78 

6.7.2. Deflection Alternatives Analyzed......................................................... 79 

6.8. Technology Readiness and Robustness ........................................................ 80 

6.8.1. Technology Readiness .......................................................................... 80 

6.8.2. Overall Effectiveness............................................................................ 80 

6.8.3. Readiness and Effectiveness Summaries.............................................. 80 

6.9. Linkage of Characterization and Mitigation................................................. 81 

6.10. Characterization Options Capabilities Matrix .............................................. 83 

6.11. Deflection Performance Analyses................................................................. 84 

6.11.1. Deflection Performance Analysis Methodology................................... 84 
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6.11.2. Deflection Performance Analysis Assumptions ................................... 84 

6.11.3. Deflection Performance Analysis Results ............................................ 85 

6.12. Deflection Mission Operations Risk............................................................. 89 

6.12.1. Overview............................................................................................... 89 

6.12.2. Notional Deflection “Campaign”.......................................................... 89 

6.12.3. Parametric Reliability Calculations ...................................................... 90 

6.12.4. Precursor Characterization Not Required ............................................. 90 

6.12.5. Precursor Characterization Required .................................................... 91 

6.12.6. Characterization Reduces Common-Cause Failure Rate...................... 92 

6.12.7. Summary............................................................................................... 93 

6.13. Possible Scenarios – Application of the Alternatives................................... 94 

6.13.1. Scenario - Apophis (Deflect before 2029)............................................ 97 

6.13.2. Scenario - Apophis (Deflect after 2029)............................................. 101 

6.13.3. Scenario - Asteroid VD17................................................................... 103 

6.13.4. Scenario – Small Asteroid with Satellite ............................................ 105 

6.13.5. Scenario – 1 km-Class Asteroid.......................................................... 109 

6.13.6. Scenario – 1-km diameter Long Period Comet................................... 111 

6.14. Fragmentation ............................................................................................. 115 

6.15. Deflection Systems Schedule Analysis....................................................... 116 

6.16. Deflection System Cost Analysis................................................................ 117 

6.17. Performance and Effectiveness vs. Cost..................................................... 118 

6.18. Other Factors............................................................................................... 118 

6.19. Findings....................................................................................................... 118 

7 Integrated Analyses ............................................................................................... 120

7.1. Warnings..................................................................................................... 120 

7.1.1. The Torino Scale................................................................................. 120 

7.1.2. Warnings Generated by the Survey .................................................... 121 

7.1.3. The Role, Effect, and Effectiveness of Warning Systems .................. 122 

7.2. Framework for Action Thresholds.............................................................. 122 

7.2.1. Phases of Discovery............................................................................ 122 

7.2.2. The Two Epochs of the NEO Problem ............................................... 124 

7.2.3. Amortized Cost of Unnecessary Threat Response Missions .............. 125 

7.3. Survey and Deflection Strategy Options..................................................... 126 

8 Summary ................................................................................................................ 128

8.1. Survey Analysis of Alternatives ................................................................. 128 

8.2. Deflection Analysis of Alternatives............................................................ 132 

8.3. Summary of Findings.................................................................................. 134 

9 Definition of Terms................................................................................................ 136

10 References .............................................................................................................. 138

11 Appendices ............................................................................................................. 146
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List of Figures 

Figure 1. Asteroid Orbit Types...................................................................................................... 22

Figure 2. Frequency of NEOs by Size, Impact Energy, and Magnitude ....................................... 23

Figure 3. History of Asteroid Discoveries..................................................................................... 24

Figure 4. Frequency of Threats by Object Size and Population.................................................... 26

Figure 5. Spaceguard Survey Status as of 10/1/06 ........................................................................ 27

Figure 6. SDT Goal of 90% of Objects Larger than 140 meters ................................................... 28

Figure 7. Detection, Tracking, and Cataloguing Figures of Merit ................................................ 32

Figure 8. Characterization Figures of Merit .................................................................................. 33

Figure 9. Complete Survey Alternatives Trade Tree..................................................................... 40

Figure 10. Pared Survey Alternatives Trade Tree ......................................................................... 42

Figure 11. Multi-Element Detection, Tracking, and Cataloguing Systems Analyzed .................. 43

Figure 12. Detection and Tracking Nominal Development Schedules ......................................... 44

Figure 13. Survey Performance for Selected Alternative Systems................................................ 48

Figure 14. Constituents of Ground-Based Survey Schedules to 90% ........................................... 48

Figure 15. Constituents of Space-Based Survey Schedules to 90%.............................................. 49

Figure 16. Possible Effect of Venus Flyby on 90% Completeness Dates ..................................... 49

Figure 17. Survey Performance of Ground-Based Combinations................................................. 50

Figure 18. Survey Performance of Ground and Space Based Combinations ................................ 50

Figure 19. Orbit Prediction Accuracy Improves with Length of Observation .............................. 51

Figure 20. Comparison of Precision Orbit Capabilities of Survey Alternatives ........................... 53

Figure 21. Value of Precovery Analysis........................................................................................ 55

Figure 22. Orbital Precision of PHOs Observed with and without Radar..................................... 56

Figure 23. The Benefits of Radar Data in the Orbit Determination Process ................................. 57

Figure 24. Percentage of Asteroids within Range of Radars......................................................... 57

Figure 25. Survey System Elements Costs through 2020 ............................................................. 58

Figure 26. Complete Characterization Trade Tree ........................................................................ 59

Figure 27. Distribution of Types of NEO Characterization Performed......................................... 61

Figure 28. Characterization Systems Development Time ............................................................. 65

Figure 29. Characterization Alternatives Life-cycle cost Results ................................................. 65

Figure 30. Survey Combinations % Completeness vs. Cost Through 2020.................................. 66

Figure 31. Survey Combinations Date of 90% vs. Life-cycle cost thru 90%................................ 67

Figure 32. PHO Approach Uncertainties and Deflection Goal ..................................................... 75

Figure 33. Illustration of PHO Deflection..................................................................................... 76

Figure 34. ΔV Required to Deflect a Hypothetical Asteroid......................................................... 76

Figure 35. Deflection Figures of Merit.......................................................................................... 78

Figure 36. Deflection Alternatives Trade Tree.............................................................................. 78

Figure 37. Deflection Performance of Impulsive Alternatives..................................................... 87

Figure 38. Deflection Performance of Slow Push Alternatives.................................................... 88

Figure 39. Illustration of Types of Deflection Mission Redundancy ............................................ 90

Figure 40. Chance of Deflection Failure if Characterization is Not Required .............................. 91

Figure 41. Chance of Deflection Failure when Characterization is Enabling ............................... 92

Figure 42. Chance of Deflection Failure if Characterization is Enhancing................................... 93

Figure 43. Momentum Capability of Impulsive Alternatives Applied to Scenarios ..................... 95

Figure 44. Momentum Capability of Slow Push Alternatives Applied to Scenarios .................... 96

Figure 45. ΔV to Avoid the Keyhole in 2029................................................................................ 98

Figure 46. ΔV required to Deflect Asteroid Apophis after 2029 Keyhole.................................. 102

Figure 47. ΔV Required to Deflect Asteroid VD17 .................................................................... 104

Figure 48. ΔV Required to Deflect a Hypothetical 200 m Asteroid............................................ 106
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Figure 49. Launch Opportunities for Optimistic Launch Constraints ......................................... 107

Figure 50. Opportunities for More Realistic Launch Constraints ............................................... 107

Figure 51. Athos Deflection Vehicle........................................................................................... 108

Figure 52. Velocity to Deflect a Hypothetical 1-km Asteroid .................................................... 110

Figure 53. ΔV to Deflect a Hypothetical 1 km Long Period Comet ........................................... 113

Figure 54. Frequency of Observed Comets by Size .................................................................... 113

Figure 55. C3 required to Reach Porthos as a Function of Time of Flight.................................. 115

Figure 56. Deflection System Estimated Development Schedules ............................................. 116

Figure 57. Deflection Concepts Development and First Mission LCC....................................... 117

Figure 58. Deflection Mission Recurring Cost............................................................................ 117

Figure 59. Deflection Performance vs. ROM Development Cost............................................... 118

Figure 60. The Torino Impact Hazard Scale ............................................................................... 120

Figure 61. Representative Warnings for a Survey Completed by 2020 ...................................... 122

Figure 62. Amortized Cost of Unnecessary Threat Response Missions ..................................... 125

Figure 63. Decision Tree for Survey and Deflection Strategy Options...................................... 126

Figure 64. Year of Reaching 90% Completeness vs. Life-Cycle Cost........................................ 130

Figure 65. Deflection Performance vs. ROM Development Cost............................................... 133

Figure 66. NEO Study Organizational Structure......................................................................... 151

Figure 67. Analysis of Alternatives Organizational Structure .................................................... 151

Figure 68. Study Calendar........................................................................................................... 152

Figure 69. Radar Observations of Asteroid (216) Kleopatra....................................................... 180

Figure 70. Radar Images of Asteroid (4179) Toutatis................................................................. 181

Figure 71. Examples of 1-year and multi-year Earth-return orbits ............................................. 183

Figure 72. Geometry of the Gravity Tractor Concept ................................................................. 211

Figure 73. Illustration of Space Tug acting on a PHO ................................................................ 216

Figure 74. PHO/Earth Encounter Geometry................................................................................ 224

Figure 75. Buildup of Baseline – Existing + Shared PS4 + Shared LSST .................................. 227

Figure 76. Buildup of Baseline + Dedicated PS8........................................................................ 228

Figure 77. Buildup of Baseline + Dedicated LSST..................................................................... 228

Figure 78. Buildup of Baseline + 0.5m IR @ L1 ........................................................................ 229

Figure 79. Buildup of Baseline + 0.5m IR @ Venus .................................................................. 229

Figure 80. Data Flow for Spaceguard Survey ............................................................................ 237

Figure 81. Statistical Distribution of PHA Orbit Periods........................................................... 240

Figure 82. Statistical Distribution of PHA Orbit Inclinations .................................................... 241

Figure 83. Statistical Distribution of PHA Orbit Eccentricities ................................................. 241

Figure 84. Approximate Distribution of Asteroid Types ........................................................... 242

Figure 85. Infrared Signatures of Asteroid Types ...................................................................... 243

Figure 86. Momentum Exchange Efficiency Depends on Porosity ........................................... 244

Figure 87. Rotation Rates as compared with Diameter - Fast Rotation Barrier ......................... 245

Figure 88. C3 of Current and Planned Launch Vehicles............................................................ 247

Figure 89. Required launch C3 for Apophis for a Fixed Launch Date....................................... 248

Figure 90. Launch C3 for Fixed Launch Date............................................................................ 249

Figure 91. Launch C3 Launch and/or Arrival Dates Vary ......................................................... 249

Figure 92. Realistic Intercept Constraints – Fixed Launch Date................................................ 250

Figure 93. Realistic Intercept Constraints – Variable Launch/Arrival Date .............................. 250

Figure 94. Breakout of Search Element Costs............................................................................. 252

Figure 95. Characterization Alternatives Life-cycle cost Results ............................................... 253

Figure 96. Breakout of Space-based Characterization Element Costs ........................................ 253
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List of Tables 

Table 1. Impact Frequencies and Typical Consequences.............................................................. 25

Table 2. Description of Ground-based Survey Alternatives Considered ...................................... 37

Table 3. Description of Space-based Survey Alternatives Considered ......................................... 38

Table 4. Description of Data Management Alternatives Considered............................................ 38

Table 5. Rationale for Detection and Tracking Alternatives not Analyzed ................................. 41

Table 6. Start Date for Ground-based Survey Elements ............................................................... 45

Table 7. Start Date for Space-based Survey Elements .................................................................. 45

Table 8. Ground-based Survey Performance................................................................................. 46

Table 9. Space-based Survey Performance ................................................................................... 46

Table 10. Survey Performance of Combinations........................................................................... 47

Table 11. Additional Benefits of Space Systems .......................................................................... 47

Table 12. The Magnitude of Various Forces Acting on Asteroids................................................ 52

Table 13. Characterization Alternatives Not Analyzed - Rationale .............................................. 58

Table 14. Characterization Capability Options ............................................................................. 63

Table 15. Possible Characterization Option Architectures and Timelines .................................... 64

Table 16. Summary of Detection, Tracking, and Cataloguing Alternatives ................................. 66

Table 17. Summary of Characterization Capability Options ........................................................ 67

Table 18. Potentially Hazardous Object Mission Timeline........................................................... 77

Table 19. Impulsive Deflection Alternatives Considered ............................................................. 79

Table 20. Slow Push Deflection Alternatives Considered ............................................................ 79

Table 21. Impulsive Alternatives Readiness and Effectiveness Summary.................................... 81

Table 22. “Slow Push” Alternatives Readiness and Effectiveness Summary ............................... 81

Table 23. Characterization Required for Impulsive Alternatives.................................................. 82

Table 24. Characterization Required for Slow Push Alternatives................................................. 82

Table 25. Deflection Alternatives Enabled by Characterization Options...................................... 83

Table 26. Apophis before 2029 Scenario Description................................................................... 97

Table 27. Apophis Keyhole – Deflection Performance Design Point 1........................................ 99

Table 28. Apophis Keyhole – Deflection Performance for Design Point 2 ................................ 100

Table 29. Apophis after 2029 Scenario Description.................................................................... 102

Table 30. Asteroid Apophis 2034 - Deflection Performance of Alternatives ............................. 103

Table 31. VD17 Scenario Description......................................................................................... 103

Table 32. Asteroid VD17 - Deflection Performance of Alternatives .......................................... 104

Table 33. 200 Meter Class Asteroid Scenario Description ......................................................... 106

Table 34. 200 Meter Asteroid - Deflection Performance of Alternatives ................................... 108

Table 35. 1-km Class Asteroid Scenario Description.................................................................. 110

Table 36. 1-km Asteroid - Deflection Performance of Alternatives ........................................... 111

Table 37. 1-km Comet Scenario Description .............................................................................. 112

Table 38. Long Period Comet - Deflection Performance of Alternatives ................................... 114

Table 39. Phases after an Object is Discovered........................................................................... 123

Table 40. Phases after the Survey is Complete ........................................................................... 124

Table 41. Summary of Detection, Tracking, and Cataloguing Alternatives ............................... 129

Table 42. Summary of Characterization Capability Options ...................................................... 131

Table 43. List of Vail NEO Workshop Attendees and Contributors........................................... 156

Table 44. Mapping of Characterization Technique to Derived Information ............................... 179

Table 45. Possible Flyby Characterization Objectives................................................................ 182

Table 46. Possible Rendezvous Characterization Objectives...................................................... 184

Table 47. Readiness for Technology Areas for Impulsive Techniques....................................... 226

Table 48. Readiness for Technology Areas for for Slow Push Techniques ................................ 226
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Table 49. Generalization of Taxonomic Classes and Densities ................................................. 243

Table 50. Space-Based Survey Systems Life-cycle cost Breakout ............................................. 252

Table 51. Space-Based Remote Characterization Systems LCC Breakout................................. 254

Table 52. Space-Based In-Situ Characterization Systems LCC Breakout .................................. 254

Table 53. Deflection Systems Life-Cycle Cost Breakout............................................................ 254

Table 54. Life-cycle costs of Characterization Capability Options............................................. 255

Table 55. Planned or Existing Large Observatories .................................................................... 256

Table 56. Proposed Very Large Observatories............................................................................ 257

Table 57. Postulated Extremely Large Observatories ................................................................. 257

Table 58. Characterization and Deflection Technology Needs Compared with Exploration ..... 267
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12 DRAFT Pre-Decisional Material 

1 Executive Summary 

In the 2005 Budget Authorization Act, the U.S. Congress directed the NASA 

Administrator to provide an analysis of alternatives to detect, track, catalogue, and 

characterize potentially hazardous near-Earth objects (NEO). Congress required that the 

Administrator submit a program by December 28, 2006 to survey 90% of the potentially 

hazardous objects measuring at least 140 meters in diameter by the end of 2020. In 

addition, the legislation required the Administrator to submit an analysis of alternatives 

that NASA could employ to divert an object on a likely collision course with Earth. 

A study team, led by the Office of Program Analysis and Evaluation (PA&E), derived 

requirements and figures of merit from the Act, and used these factors to evaluate the 

alternatives. The team developed a range of options from public and private sources and 

then analyzed their capabilities, levels of performance, life-cycle costs, schedules, and 

development and operations risks. This document presents the detailed results of these 

analyses. A summary report was submitted to Congress in December of 2006. 

1.1. Survey Analysis of Alternatives 

Detection and tracking alternatives identified by the study team included optical systems 

located on the ground and optical and infrared assets located in space. For ground-based 

alternatives, the study team considered sharing planned observatories such as 

PanSTARRS 4 (PS4), funded by the U.S. Air Force, and the Large Synoptic Survey 

Telescope (LSST), partially funded by the National Science Foundation. The team also 

considered new NASA-funded facilities that would be dedicated to the search for 

hazardous objects and would be based on these planned observatories. Although cost 

margin was applied to alternatives that leveraged planned assets, programs that rely on 

these projects may carry additional cost and schedule risk. Specific results include: 

• An architecture, which combines the sharing of the planned PS4 and LSST 

systems with a second, dedicated NASA-funded LSST, was the only groundbased alternative able to meet the congressional goal of identifying 90% of the 

hazardous objects by 2020. This combination is estimated to have a life-cycle cost 

of $820M ($FY06). 

• A shared PS4, a shared LSST, and a dedicated NASA-funded PS8 were able to 

catalog 90% of hazardous objects by 2024, with a life-cycle cost of $560M. 

• A dedicated, NASA-funded observatory based on LSST’s design was also able to 

catalog 90% of potentially hazardous objects in 2024 without the contributions of 

other programs. Its estimated life-cycle cost is $870M. 

Space-based search alternatives were located in low-Earth orbit, at Sun-Earth Lagrange 

points, and in heliocentric Venus-like orbits. Only an infrared system operating in a 

Venus-like orbit was able meet the congressional goals without the contribution of shared 

ground-based assets. All space-based alternatives were able to meet the goals when 

combined with a ground-based baseline of a shared PS4 and a shared LSST. 
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A space mission failure could delay achieving the 90% goal by up to 5 years, after which 

the catalog could be completed with shared ground-based assets. Infrared systems 

operating in space could provide more accurate size estimates of up to 80% of objects in 

the catalog. Observatories located in a Venus-like orbit are the most efficient at finding 

objects inside Earth’s orbit, a potentially underestimated population. Additionally, by the 

end of 2020, infrared systems in Venus-like orbits can find 90% of the objects measuring 

over about 80 meters, exceeding the 140-meter requirement. Finally, space-based systems 

have much less uncertainty in the date of reaching 90% due to their superior sensitivity. 

Selected space-based alternatives include: 

• A 0.5-meter infrared system operating in a Venus-like heliocentric orbit 

completes 89% of the survey by 2020 which is within the uncertainty of the 

analysis. This system has a life-cycle cost of $840M ($FY06). 

• A similar 0.5-meter infrared system operating in a Venus-like orbit and working 

in concert with a shared PS4 and a shared LSST completes 90% of the survey in 

2018, with a life-cycle cost of $1B through 2018. 

• A 0.5-meter infrared system operating at Sun-Earth L1 in conjunction with the 

baseline finishes 90% of the survey in 2020. Its life-cycle cost is $1.1B. 

Infrared systems with a 1.0-meter aperture complete the survey about 1 year earlier than 

the 0.5-meter alternatives described above, and have life-cycle costs about $300M higher. 

Optical systems with 1.0-meter and 2.0-meter apertures in Venus-like orbits, combined 

with the baseline ground-based systems, completed the survey by 2017 and 2019 

respectively, with life-cycle costs in excess of $1.7B. The visible system with a 2.0-meter 

aperture progressed more slowly than the 1.0-meter system due to differences in 

development time. Acquisition of new systems was assumed to start October 1, 2007, and 

delays in funding will affect the ability of some alternatives to meet the 90% 

completeness goal by the end of 2020. 

Congress provided two objectives for characterizing potentially hazardous objects. The 

first objective, to “assess the threat,” requires analysts to determine the orbit and 

approximate the mass of each hazard. Detection and tracking systems with judicious 

follow-up are all able to provide warning, and some are able to provide very good size 

and mass estimates. Systems operating in the visible spectrum are limited by a factor of 

two for size estimates, resulting in a factor-of-eight uncertainty in mass. Infrared systems 

provide data for much more accurate size estimates. 

If detection systems must characterize the catalog, the time to complete the survey to a 

90% completion level will be extended. Furthermore, the costs of these systems may 

increase $100M-$400M to accommodate filters and additional data processing. In 

addition, the smallest and faintest objects may remain visible to sensors only for a few 

days or weeks. Therefore, if characterization is required and it is not performed by 

detection systems, either formal relationships with extant observatories for “on demand” 

access must be negotiated or new dedicated characterization facilities will be needed. 

Radar may quickly and precisely characterize and determine the orbit of about 10-25% of 

the objects of interest within 5 years of their detection. While the number of objects 

observed by radar increases with time, the relative value of radar to precisely determine 
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2006 Near-Earth Object Survey and Deflection Study 

14 DRAFT Pre-Decisional Material 

the orbits of the full catalog declines over the same period. Orbits determined from 

optical data alone will nearly match the accuracy of radar-improved orbits after decades 

of observation. Therefore, the utility of radar is limited to a relatively few “short 

warning” cases that may be of very high interest during the survey. Up to $100M in 

funding (not included in detection and tracking life-cycle costs) may be required to 

maintain radar capability through 2020, as NASA and National Science Foundation 

funding for existing radars is currently in flux. 

The second objective of characterization is to “inform mitigation.” Depending on the 

mitigation strategy selected, this objective may require information beyond the size and 

orbit of potential threats. This information may include the structure, porosity, rotation 

rate, material composition, and surface features of the threats. The deflection alternatives 

considered are sensitive to the maximum mass that needs to be deflected, but some 

alternatives are orders of magnitude less sensitive than others. 

Characterization by remote sensing provides some information about the diversity of 

objects in the population. From this information, analysts build models that can be used 

to infer a limited number of characteristics of a particular object. Only in-situ encounters 

can provide the definitive observations necessary to calibrate the remote observations. 

More importantly, only in-situ visits can obtain the information needed by some of the 

deflection alternatives to mitigate a specific threat. For credible threats with sufficient 

warning, it is expected that in-situ characterization will always be performed to both 

confirm the probability of impact (with a transponder, for example) and to characterize 

the potential threat if deflection is necessary. 

This study has determined that it is premature to set specific characterization 

requirements to enable mitigation until a mitigation strategy has been determined; 

therefore, the study has developed characterization options that provide a range of 

capabilities. These options included the use of detection and tracking assets, dedicated 

ground and space systems for remote observation, and in-situ missions to inform 

mitigation of threats with sufficiently high impact probabilities. These options have lifecycle costs ranging from $50M-$8B ($FY06) over several decades. 

It is expected that during the 5-10 years of a survey, a total of 500,000 objects will be 

discovered by more than 2 million individual observations. About 21,000 of these objects 

will measure 140 meters or larger and be tracked as potentially hazardous. Although this 

study uses an estimate of the population of potentially hazardous objects based on 

statistical projections, the actual number of objects will not affect the date of reaching the 

90% goal as long as the objects are approximately distributed in orbits as predicted. 

This volume of observations will require a data-processing capability that is 100 times 

more capable than current cataloguing systems. After objects are detected, the system 

must be able to obtain follow-up observations, store and distribute collected data, and 

analyze these data for observed but previously undetected objects. Currently, 

uncompensated or under-funded analysts perform many of these functions. Such an 

approach likely will not remain viable. Finally, either the NASA Survey or otherwise 

funded activities, such as PS4 and LSST, are expected to produce impact warnings at a 

rate that is 40 times greater than what is experienced today. This much higher rate of 

warnings will start as early as 2010. 
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1.2. Deflection Analysis of Alternatives 

The study considered a wide range of techniques to divert a threatening object. These 

alternatives were broadly classified as “impulsive” if they acted nearly instantaneously, 

or “slow push” if they acted over an extended period of time. Launch, orbit transfer, 

technology development, and object characterization requirements were developed for 

each of these alternatives. They were applied to a set of five scenarios representing the 

likely range of threats over million-year timescales. 

The use of nuclear explosives was found to be the most effective alternative in the near 

term. While an explosion on or below the surface of a threatening object is 10-100 times 

more effective than a detonation above the surface, the standoff detonation would be less 

likely to fragment the target. Nuclear options require the least amount of detailed 

information about the threatening object. A nuclear standoff mission could be designed 

knowing only the orbit and approximate mass of the threat, and most impulsive missions 

could be carried out incrementally to reach the required amount of deflection. Additional 

information about the object’s mass and physical properties would perhaps increase the 

effectiveness, but likely would not be required to accomplish the goal. The study 

examined conventional explosives, but found they were ineffective against most threats. 

Kinetic impact alternatives are the most effective non-nuclear option, transferring 10-100 

times less momentum than nuclear options for a fixed launch mass. Impact velocities, 

varying from 10-50 km/s, produced a factor-of-three variation in deflection performance. 

In addition, kinetic impacts also are sensitive to the porosity, elasticity, and composition 

of the target and may require larger performance margins if these characteristics are not 

well determined. 

Slow push techniques analyzed in this study included a gravity tractor, which would alter 

the course of an object using the gravitational attraction of a massive spacecraft, and a 

space tug, which would attach to an object and move it using high-efficiency propulsion 

systems. An attached space tug has generally 10-100 times more performance than the 

gravity tractor, but it requires more detailed characterization data and more robust 

guidance and control and surface attachment technologies. Slow push techniques were 

determined to be useful in relatively rare cases (fewer than 1% of expected threat 

scenarios), but these techniques could be effective in instances where small increments of 

velocity (less than 1 mm/s) could be applied to relatively small objects (less than 200 

meters in diameter) over many decades. 

The level of risk reduction required of a deflection campaign needs to be clearly 

understood, as it has a first-order impact on cost and complexity. While this report uses a 

goal of reducing the probability of impact to 1 in 1 million, this is not a nationally or 

internationally accepted threshold. Additionally, when designing the deflection campaign, 

planners must take into account that launch vehicles and interplanetary spacecraft fail at 

relatively high rates (2-5% for launches; 10+% for spacecraft) and that deflection 

approaches may not perform as designed. Planning for many flights of multiple 

spacecraft designs launched from several different launch vehicles may be necessary to 

achieve the reduction in impact probability projected to be required. 
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2006 Near-Earth Object Survey and Deflection Study 

16 DRAFT Pre-Decisional Material 

1.3. Summary of Findings 

• Combining optical ground-based observatories currently under development with 

a dedicated ground-based asset can reach the congressional goal by the end of 

2020. Life-cycle cost for this architecture, including a robust data-management 

and data-analysis infrastructure, is estimated to be $820M through 2020. 

• Space-based infrared systems, combined with shared ground-based assets, could 

reduce the overall time to reach the 90% goal by up to 3 years, with life-cycle 

costs of $1.0-$1.3B through 90% completion. Space systems have additional 

benefits and risks over ground-based alternatives, and are generally more capable 

(sensitive) than ground based alternatives. 

• Radar systems cannot contribute to the search for potentially hazardous objects, 

but may be used to rapidly refine tracking and to determine object sizes for a few 

objects of potentially high interest. Existing radar systems are oversubscribed, and 

funding to operate these systems may be in flux. A budget for radar is not 

included in the detection and tracking life-cycle costs. 

• Determining an object’s mass and orbit are required to determine whether it 

represents a threat and to inform deflection alternatives. Beyond these parameters, 

characterization requirements and capabilities are tied directly to the mitigation 

strategy selected. Life-cycle costs for the characterization options vary by billions 

of dollars depending on the mitigation strategy pursued. 

• While several countries have capable programs to study near-Earth objects, none 

of these efforts has materially influenced the results of the study team. 

• Nuclear standoff explosions are assessed to be 10-100 times more effective than 

the non-nuclear alternatives analyzed in this study. Other techniques involving 

nuclear explosives may be more effective, but they run an increased risk of 

fracturing the target. They also carry higher development and operations risks. 

• Kinetic impactors are the most mature approach and could be used in some 

scenarios, especially for objects that consist of a relatively small, solid body. 

• Slow push deflection techniques are the most expensive, and their ability to both 

travel to and divert a threatening object is limited unless mission durations of 

many decades are available. 

• Deflection campaigns may need to be 100-1,000 times more reliable than current 

space missions to meet mitigation requirements. 

• Many potentially hazardous objects (30-80%) are in orbits that are beyond the 

capability of current or planned launch systems. Therefore, if these objects need 

to be deflected, swingby trajectories or on-orbit assembly of modular propulsion 

systems may be required to augment launch vehicle performance. 
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1.4. Summary of Program Options 

The figure below shows a decision tree for selected strategy options for both parts of this 

study. In Step 1, a detection and tracking architecture is selected. Ground-based, spacebased, and combinations of ground- and space-based systems can meet the congressional 

goal of detecting 90% of potentially hazardous objects 140 meters and larger by 2020. In 

Step 2, it is expected that search systems can provide a portion of the orbit and mass 

characterization required to assess the threat, but options for enhancing this capability 

including radar are presented. In Step 3, mitigation strategy options (including perhaps 

deflection demonstrations) are chosen informing the choice of further characterization in 

Step 4 (if required). 

Infrared in Venus-like orbit + Baseline 97% $1.1B

Infrared @ L1 + Baseline 91% $1.1B

Infrared in Venus-like orbit 89% $0.84B

Dedicated ground-based 85% $0.66B

Dedicated ground-based + Baseline 90% $0.82B

Shared ground-based (Baseline)* 83% $0.31B

Option FY06 $ through 2020 LCC thru 2020*

Infrared in Venus-like orbit + Baseline 97% $1.1B

Infrared @ L1 + Baseline 91% $1.1B

Infrared in Venus-like orbit 89% $0.84B

Dedicated ground-based 85% $0.66B

Dedicated ground-based + Baseline 90% $0.82B

Shared ground-based (Baseline)* 83% $0.31B

Option FY06 $ through 2020 LCC thru 2020*

2018 $1.0B

2020 $1.1B

2021 $0.86B

2020 $0.82B

2024 $0.87B

2026 $0.52B

Year 90% LCC to 90%*

2018 $1.0B

2020 $1.1B

2021 $0.86B

2020 $0.82B

2024 $0.87B

2026 $0.52B

Year 90% LCC to 90%*

1. Select Detection and Tracking Architecture other similar options exist

~$150M cataloguing costs in all options 1-2% uncertainty in relative performance

2. Add Orbit and Mass Characterization 

Dedicated space $1-$1.5B

Radar for some threats < $100M

Dedicated ground $100-$300M

Augment search systems $100M-$400M

Use search systems $0B

Options FY06 $ LCC thru 2020*

Dedicated space $1-$1.5B

Radar for some threats < $100M

Dedicated ground $100-$300M

Augment search systems $100M-$400M

Use search systems $0B

Options FY06 $ LCC thru 2020*

3. Develop a Mitigation Strategy

Cooperate with international efforts $50M+

Develop nuclear explosives technologies ?

Perform deflection demonstration (by 2014) $100M+

Coordinate deflection technologies w/ ESMD $0+

No dedicated preparation $0+

Options FY06 $ LCC thru 2020

Cooperate with international efforts $50M+

Develop nuclear explosives technologies ?

Perform deflection demonstration (by 2014) $100M+

Coordinate deflection technologies w/ ESMD $0+

No dedicated preparation $0+

Options FY06 $ LCC thru 2020

4. Develop Characterization Further if Needed for the Mitigation Strategy ($0-$8B)

 * LCC = Life-cycle cost 
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18 DRAFT Pre-Decisional Material 

2 Participants 

2.1. Study Leadership 

• Bill Claybaugh, Study Lead – NASA/HQ/PA&E 

• Dan Mulville, Study Coordinator – NASA (ret) 

• Marcus Shaw, Study Executive – The Aerospace Corporation/PA&E 

• Lindley Johnson, NEO Observation Program Manager 

• Ed Barker, Detection Lead – NASA/JSC 

• Don Yeomans, Detection Co-Lead – NASA/JPL 

• Tom Morgan, Characterization Lead – NASA/HQ/SMD 

• Rob Gold, Characterization Co-Lead – APL/JHU 

• Vern Weyers, Deflection Co-Lead – NASA (ret) 

• Larry Ross, Deflection Co-Lead – NASA (ret) 

2.2. Study Team 

• Members of the 2006 NEO study working groups provided significant technical 

input, technical direction, and review of the study results, see Appendix B for the 

full study team, schedule, and structure 

2.3. Analyses 

• The JPL NEO Program Office (Steve Chesley and Paul Chodas) performed 

survey performance, precision orbit determination performance, and warning 

prediction analyses. 

• David Bearden and Matt Hart, of The Aerospace Corporation, led concept 

development, deflection performance, and cost, risk, and schedule analyses. 

2.4. Acknowledgements 

• Survey performance model validation was performed using results from 

MIT/Lincoln Labs and led by Grant Stokes 

• Additional technical analyses and an independent technical review were 

performed by NASA LaRC and led by Dan Mazanek 

• Members of the June 2006 NEO Public Workshop held in Vail, Colorado 

provided significant technical content and many detection, characterization, and 

deflection concepts (See Appendix D) 

• The NEO Survey Science Definition Team Report provided the foundation for 

this study (See Section 4.5) 

• A Red Team composed of Jay Greene (NASA, retired), General Lester Lyles 

(USAF, retired), Dr. Lucy McFadden (University of Maryland), and Liam 

Sarsfield, and provided independent review of the study structure, derived 

requirements, study results, and this final report. 
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20 DRAFT Pre-Decisional Material 

3 Introduction and Overview 

3.1. Purpose 

The Administrator of the National Aeronautics and Space Administration (NASA) 

submits this report pursuant to the requirements stated in the George E. Brown Jr. NearEarth Object Survey Act (“the NEO Survey Act”), which Congress passed as part of 

Public Law No: 109-155: the 2006 NASA Authorization Act of 2005 documented in 

Appendix A and Reference [1]. 

In this Act, the Congress directed that the NASA Administrator initiate a Near-Earth 

Object (NEO) Survey program to detect, track, catalogue, and characterize objects larger 

than 140 meters in diameter, with a perihelion distance of less than 1.3 AU (Astronomical 

Units) from the Earth. The Survey would warn of and inform attempts to mitigate the 

hazard. The Survey is to be 90% complete within 15 years after the date of enactment. 

3.2. Statement of Need 

In the NEO Survey Act, Congress made the following findings: 

• Near-Earth objects pose a serious and credible threat to humankind, as many 

scientists believe that a major asteroid or comet was responsible for the mass 

extinction of the majority of the Earth’s species, including the dinosaurs, nearly 

65,000,000 years ago. 

• Similar objects have struck the Earth or passed through the Earth's atmosphere 

several times in the Earth’s history and pose a similar threat in the future. 

• Several such near-Earth objects have only been discovered within days of the 

objects’ closest approach to Earth, and recent discoveries of such large objects 

indicate that many large near-Earth objects remain undiscovered. 

• The efforts taken to date by NASA for detecting and characterizing the hazards of 

near-Earth objects are not sufficient to fully determine the threat posed by such 

objects to cause widespread destruction and loss of life. 

3.3. Direction 

The NEO Survey Act [Appendix A] amended the Space Act of 1958 such that 

The Congress declares that the general welfare and security of the United States 

require that the unique competence of the National Aeronautics and Space 

Administration be directed to detecting, tracking, cataloguing, and characterizing 

near-Earth asteroids and comets to provide warning and mitigation of the potential 

hazard of such near-Earth objects to the Earth. 

In addition, it states 

The Administrator shall plan, develop, and implement a Near-Earth Object Survey 

program to detect, track, catalogue, and characterize the physical characteristics of 
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near-Earth objects equal to or greater than 140 meters in diameter to assess the 

threat of such near-Earth objects to the Earth. It shall be the goal of the Survey 

program to achieve 90 percent completion of its near-Earth object catalogue (based 

on statistically predicted populations of near-Earth objects) within 15 years after the 

date of enactment of this Act. 

The Congress has directed the NASA Administrator to deliver the following by one year 

from the law’s enactment: 

(1) An analysis of possible alternatives that NASA may employ to carry out the 

Survey program, including ground-based and space-based alternatives with 

technical descriptions. 

(2) A recommended option and proposed budget to carry out the Survey program 

pursuant to the recommended option. 

(3) Analysis of possible alternatives that NASA could employ to divert an object on a 

likely collision course with Earth. 
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4 Background 

4.1. Asteroids and Comets 

Asteroids and comets are the two types of potentially hazardous objects discussed in this 

study. Astronomers distinguish these bodies on the basis of their appearance. Moving 

objects that appear as a star-like point of light are known as asteroids. The existence of 

asteroids was not known until about 200 years ago when telescopes became powerful 

enough to detect them. Moving objects that appear diffuse or those that have visible tails 

are known as comets, and because of their distinctive tails, people have known about 

comets since antiquity. It has taken several generations of improvements in telescope 

design to detect and understand the small bodies that orbit near and periodically collide 

with Earth. 

Differences in their appearance reflect in part a difference in their composition. 

Generally, asteroids are relatively rocky or metallic objects without atmospheres, while 

comets are composed in part of volatiles such as water ice that vaporize when heated. 

Comets that are far from the Sun or those that have lost most of their volatiles often look 

like an asteroid. A volatile-rich object will develop an atmosphere only when heated 

sufficiently by a relatively close approach to the Sun. 

The asteroids are categorized as Apollos, Atens, Amors, and Interior Earth Objects 

(IEOs), depending on whether their orbits cross Earth’s orbit with a period of more than 

1 year, cross Earth’s orbit with a period of less than 1 year, exist completely outside the 

Earth’s orbit, or exist completely within the Earth’s orbit, respectively. The distribution 

of these objects in the NEO population is shown in Figure 1. 

Apollo

Semimajor Axis ≥ 1.0 AU 

Perihelion ≤ 1.02 AU

Earth Crossing

Amor

1.02 AU < Perihelion ≤ 1.3 AU

Apollo 62% of known asteroids

Type Near-Earth Population

IEO 6 known asteroids

Amor 32% of known asteroids

Aten 6% of known asteroids

Apollo 62% of known asteroids

Type Near-Earth Population

IEO 6 known asteroids

Amor 32% of known asteroids

Aten 6% of known asteroids

Aten

Semimajor Axis < 1.0 AU 

Aphelion ≤ 1.0167 AU

Earth Crossing

Inner Earth Objects (IEOs) 

Aphelion < 0.983 AU 

Always inside Earth’s orbit 

(aka Apohele)

Figure 1. Asteroid Orbit Types 
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2006 Near-Earth Object Survey and Deflection Study 

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Near-Earth Objects (NEOs) are asteroids and comets in orbits that allow them to enter 

Earth’s neighborhood, defined by astronomers as having a perihelion (closest approach to 

the Sun) of less than 1.3 AU (Astronomical Units or approximately the mean distance 

between the Sun and Earth). Extinct comets may make up 5-15% of the NEO population, 

and some may retain volatiles. [2] 

More relevant to this report is the definition of Potentially Hazardous Objects (PHOs), 

asteroids and comets that have a potential to someday impact the Earth. A PHO is an 

object in our solar system that passes within 0.05 AU (about 7.5 million km) of Earth’s 

orbit and is large enough to pass through Earth’s atmosphere; that is, about 50 meters and 

larger. Approximately 21% of the NEOs of any given size class are expected to be 

potentially hazardous. 

4.2. Population of Near-Earth Objects 

A constant power law shown in Figure 2 can approximate the number of NEOs of a 

particular size. The figure shows a hundred-fold increase in the number of NEOs as the 

diameter decreases by an order of magnitude. Figure 2 also shows the approximate 

absolute magnitude (brightness) of the objects, their average impact interval, and the 

approximate impact energy they would deliver in a collision with Earth. [2] In any given 

size class, this estimate is probably accurate to within a factor of two or three, as there are 

not enough observations in some classes to form a statistically valid sample. In this report 

the term PHO will be used to indicate potential threats, with the understanding that those 

smaller than 1 km are predominantly asteroids, because comets do not add substantially 

to the population below 1 km. 

~100k NEOs > 140m

~1k NEOs > 1km

Constant Power Law

Provides Good Fit to Data

Figure 2. Frequency of NEOs by Size, Impact Energy, and Magnitude 

The history of the known asteroid population is one of exponential growth. Astronomer 

Guiseppe Piazzi of Palermo, Sicily, discovered the first asteroid on January 1, 1801. 

Early asteroid discoveries usually occurred by chance while astronomers observed or 
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2006 Near-Earth Object Survey and Deflection Study 

24 DRAFT Pre-Decisional Material 

searched for other objects. The number of new finds increased to five per year by 1865, 

15 per year by 1895, 25 by 1910 and up to about 40 by 1930. By 2006, the number of 

known asteroids was about 340,000 including about 800 PHOs. 

Figure 3 composed of images created by the staff of the Amragh Observatory shows a 

graphical history of asteroid discoveries. For many years, it was assumed that asteroids 

that crossed Earth’s path do not exist because none had been observed. It remains to be 

determined if the number found inside the Earth’s orbit is greater than or less the number 

than that expected because current observational methods are unable to reliably detect 

objects there. Red dots represent asteroids that cross the Earth’s path, while those 

represented by green dots do not. Note that these dots are not to scale and the colors are 

intended only to give a visual cue to the pace of discovery. 

Figure 3. History of Asteroid Discoveries 

Assuming comets have an enhanced signature when they come within 1.3 AU of the Sun, 

the majority of short-period comets probably have been found. The Spaceguard Survey 

has found two or three short-period comets in the last several years (see Section 4.4). 

Comets with orbital periods longer than 20 years only will be found when their perihelion 

(closest approach to the Sun) brings them close enough for their volatiles to vaporize and 

produce the distinctive tail. The total number of near-Earth comets is unknown, but it is 

estimated to be smaller than 1% of the population of near-Earth asteroids. 

To estimate the performance of the detection and tracking alternatives, a synthetic 

population of asteroids and comets was used. This population is based on the best 

available estimate of the orbital distribution of objects. This study used an estimate of the 

NEO population developed by Bottke, et al [4], as did the Science Definition Team [2] 
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2006 Near-Earth Object Survey and Deflection Study 

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report discussed in Section 4.5. Although the synthetic population selected for analysis 

may vary by as much as a factor of two, the actual number of objects will not affect the 

date of reaching the 90% goal as long as the objects are approximately distributed in 

orbits as predicted. Sensitivity to variations in the statistical population or orbital 

distribution of objects was not performed. 

The Minimal Orbital Intersection Distance (MOID) is defined as the closest possible 

distance between the orbits of two objects. Assuming random values for the argument of 

perihelion and the longitude of node, about 20% of NEOs (about 20,000 of 100,000 

NEOs ≥ 140 meters) have an Earth (MOID) smaller than 0.05 AU and, therefore, are 

PHOs. About 1% of the NEOs have a MOID smaller than the Moon’s distance from the 

Earth, and the probability of having a MOID smaller than the Earth’s radius is about 

0.025%. This result does not imply that a collision with Earth is imminent because both 

the Earth and the object still must be at the same location at the same time. See Figure 2 

for the frequency of impacts by size. 

The Aten asteroids spend most of their time inside Earth’s orbit. Currently, they account 

for about 6% of observed near-Earth asteroids and 13% of the Earth-crossing ones. 

However, observational biases may play against their discovery; thus they may be 

underestimated. Another potentially under-observed population of asteroids, called IEO 

(interior Earth objects), could exist and evolve almost entirely inside Earth’s orbit, but 

still present a hazard. [5] These objects can be readily discovered only from orbits well 

inside the Earth’s orbit. Only observations from vantage points inside Earth’s orbit will 

determine if their number is accurately represented in current population estimates. 

4.3. The Probability of Impact and Distribution of Sizes 

The chance that an object 140 meters or larger will strike the Earth in any given year is 

about 0.0002 (about 1 every 5,000 years on average). The random nature of the hazard 

means that it is equally probable that a 140-meter object will hit the Earth in the next 50 

years (~1%) or that the Earth will experience no impacts of that size in the next 23,000 

years (0.999823,000 ~ 1%). The occurrence or absence of past events has no influence on 

the likelihood of future impacts. 

Table 1 developed from Reference [11] reproduces a set of estimated impact frequencies 

and consequences for a range of PHO diameters. 

Table 1. Impact Frequencies and Typical Consequences 

Type of Event 

Diameter of 

Object 

Fatalities 

per Impact 

Typical Impact 

Interval (years) 

High altitude break-up < 50 m ~0 annual 

Tunguska-like event > 50 m ~5,000 250 - 500 

Regional event > 140 m ~50,000 5,000 

Large sub-global event > 300 m ~500,000 25,000 

Low global effect > 600 m > 5 M 70,000 

Nominal global effect > 1 km > 1 B 1 million 

High global effect > 5 km > 2 B 6 million 

Extinction-class Event > 10 km 6 B 100 million 
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2006 Near-Earth Object Survey and Deflection Study 

26 DRAFT Pre-Decisional Material 

Figure 4 projects the frequency of impacts by object size for two populations of hazards, 

those larger than 50 meters and those larger than 140 meters. The plot shows the 

percentage of impacts smaller than the corresponding diameter on the X-axis. For all 

threats (larger than 50 meters), there is about a 70% chance that an impacting object will 

be smaller than 100 meters in diameter and a 95% chance that the object will be smaller 

than 200 meters in diameter. These data show that threats are much more likely to be 

relatively small. 

0%

10%

20%

30%

40%

50%

60%

70%

80%

90%

100%

10 100 1000

Diameter of Object (meters)

Probabilty Threat is Smaller Than X-axis

Population of Hazards

Larger than 50 m

Population of Hazards

Larger than 140 m

Figure 4. Frequency of Threats by Object Size and Population 

Additionally, if the congressionally mandated survey detected nearly 100% of the objects 

measuring 140 meters and larger, the figure projects the size distribution of identified 

impact threats in that population. This information, in conjunction with the object’s orbit, 

launch capability, and deflection performance, may be used as an initial filter to 

determine the percentage of threats that an alternative will be effective at mitigating. 

Rotation rate, binarity, composition, porosity, and companion objects will further narrow 

deflection options. 

4.4. NASA Spaceguard Survey 

In a 1992 report to NASA [6], a team led by David Morrison recommended that a 

coordinated Spaceguard Survey be initiated to discover, verify, and provide follow-up 

observations of Earth-crossing asteroids. This survey was expected to discover within 25 

years 90% of the objects that measured more than 1 km in diameter. Three years later, 

another NASA study led by Eugene Shoemaker [7], recommended a search to discover 

60-70% of the same size objects within 10 years. The team set a goal of obtaining 90% 

completeness within another 5 years. 

In 1998, NASA formally accepted the goal of finding and cataloging 90% of all NEOs 

with sizes of 1 km or larger by 2008. This size was chosen after a study indicated that the 

impact of an object larger than 1 km could cause a global catastrophe [6], whereas 
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smaller objects would likely to have local or regional consequences. The estimated 

population of NEOs measuring 1 km and larger is about 1,100. 

The Near-Earth Object Observation (NEOO) Program, coordinated by NASA’s Jet 

Propulsion Laboratory (JPL) in Pasadena, CA, manages NASA-sponsored efforts to 

detect, track, and characterize potentially hazardous asteroids and comets that approach 

the Earth. The progress of the survey to date is depicted in Figure 5. In addition to 

detecting and cataloging of NEOs, the NEOO Program is responsible for facilitating 

communications between the astronomical community and the public as potentially 

hazardous objects are discovered. 

0

200

400

600

800

1000

1200

Estimated Total = 1100

90% Target

Previous 2000 2002 2004 2006 2008 2010

689

Cumulative # of NEOs > 1 km Discovered

Figure 5. Spaceguard Survey Status as of 10/1/06 

While current NEO surveys are dedicated to finding the largest objects, they also 

serendipitously find many that are sub-km in size. However, the current survey systems 

are not capable of performing a comprehensive search for the smaller objects as specified 

by the current congressional direction and recommended by the Science Definition Team. 

4.5. Small Near-Earth Object Survey Science Definition Team Report 

In August 2002, NASA chartered a Science Definition Team (SDT) to study the 

feasibility of extending the search for objects with smaller diameters. [2] The SDT report 

addressed the following topics: 

Search size limit - The team recommended that a survey produce a catalog that is 90% 

complete for PHOs larger than 140 meters. Figure 6 shows the rationale for selecting 

90% of objects larger than 140 meters – this fraction is calculated to eliminate 90% of the 

remaining sub-global risk while retiring virtually all of the global risk from objects 

greater than 1 km. A search system could be constructed to catalog hazardous objects 

down to the air blast limit (about 50 meters in diameter). However, the team suggested 
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that cataloging down to 140 meters was the preferred approach for the next-generation 

survey, which also would provide warning for 60-90% of objects capable of producing 

potentially dangerous air blasts. 

Comets - The frequency with which long-period comets closely approach the Earth is 

roughly one-hundredth the frequency of asteroids. This is a relatively small risk, and 

therefore should not be included in the goals of the next survey. 

Technical feasibility and schedule - The resources made available to the effort would 

drive the survey; technology would not. The survey could be completed in 7-20 years. 

Survey Goal

90% of sub-global 

hazard eliminated by 90% 

completeness for D>140m

Additional Benefit

~99% of global hazard 

eliminated for D > 1 km

Figure 6. SDT Goal of 90% of Objects Larger than 140 meters 

Survey Alternatives - The team identified a series of specific ground-based, space-based 

and combinations of ground- and space-based systems that could accomplish the nextgeneration search. The team indicated that the choice of specific systems would depend 

on the time allowed for the search and the resources available. 

Cost - The SDT estimated that the survey systems required to eliminate 90% of the risk 

for sub-km NEOs would cost between $236-$397 million ($FY2003) if accomplished 

within 20 years. 
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5 Survey Program Analysis of Alternatives 

5.1. Introduction 

In the NEO Survey Act, the U.S. Congress directed that NASA perform an analysis of 

alternatives and propose a NEO Survey Program to detect, track, catalogue, and 

characterize 90% of potentially hazardous objects 140 meters and larger by the end of 

2020. This chapter provides an analysis of alternatives that may be used to support 

program decisions. The descriptions required by the Act may be found in Appendix E. 

5.2. Derived Requirements and Definitions 

Several requirements were explicit in the definition of this study. For example, Congress 

explicitly directed that the Survey catalogue 90% of NEOs whose diameters exceeded 

140 meters by the end of 2020. Derived requirements are levels of performance, cost, 

schedule, or risk that are inferred or derived from explicit requirements. These include 

more specific technical definitions of detection, cataloguing, and characterization as well 

as the synthetic set of NEOs used for statistical analyses of completeness. The derived 

requirements drawn from the congressional language are as follows. 

5.2.1. Survey Goal

Since, by definition, objects that do not pass within 0.05 AU of the Earth are not 

“potentially hazardous,” these objects are assessed (without necessity of discovery) to be 

no threat to Earth. Therefore, the Survey’s goal will be to detect, track, catalogue, and 

characterize 90% of all PHOs greater or equal to 140 meters by the end of 2020, rather 

than 90% of all NEOs as described by the congressional language. This will amount to 

finding at least 18,000 of and expected 20,000 PHOs ≥ 140 meters. 

Limiting the objects to only PHOs will require the Survey to find fewer objects. PHOs 

are brighter (on average) than NEOs because they pass closer to the Earth (by definition). 

Since the subset of NEOs that are PHOs are easier to find, the time to meet the goal will 

be reduced. However, finding 90% of PHOs will provide equally effective “warning and 

mitigation of the hazard.” 

5.2.2. Detect, Track, Catalogue, and Characterize

Congressional direction uses the terms detection, tracking, cataloguing, and 

characterization without specifically defining them. These terms are defined to mean: 

• Detect – Discovery of objects greater than 140 meters in diameter 

• Track – Observing an object twice within 1 week, generally resulting in initial 

orbit determination and ephemeris generation 

• Catalogue – Generating full and precise orbit determination including precovery 

(see Section 5.10.2) and follow-up observations. Cataloguing includes calculating 

uncertainties, publishing the orbit, and archiving the data. A third observation 

within 40 days is required. 
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In the congressionally directed Survey, characterization serves two purposes. 

Characterization “to assess the threat” and “to provide warning” is defined as refining 

the orbit and approximating the mass of PHOs. This primarily will be accomplished with 

detection and tracking assets, although mass may be more accurately refined using other 

follow-up assets. 

Characterization “to provide ... mitigation options” ties characterization requirements to 

mitigation. The deflection options examined in this study are broadly representative, and, 

therefore will help to understand the utility and cost of a range of characterization 

capabilities. However, this analysis does not allow one to define specific characterization 

requirements. 

Therefore, characterization is classified into two parts. 

• Threat characterization is defined as precision orbit determination and risk 

analysis using available data. Threat characterization predicts the probability of 

Earth impact, and provides warning, date, time, relative velocity and estimated 

impact energy. Appropriate alerts are issued from orbit characterization products. 

• Object characterization is defined as all other efforts to obtain information 

needed for mitigation. This may include mass, size, structure, rotation rate, 

material, and determination of the existence of companion objects. 

Object characterization may be accomplished using remote sensing assets as well as insitu space missions. The required characterization capability will be determined by the 

mitigation strategy and deflection approaches selected. For some mitigation strategies, 

marginal remote sensing capability may be adequate to meet objectives. For other 

strategies, models based on remote sensing data may need to be calibrated by in-situ 

visits. Finally, other mitigation strategies may assume that sufficient warning will exist to 

conduct in-situ visits for every threat, which becomes more likely as the catalog is 

completed. Pending the selection of a mitigation strategy, this study examines a range of 

characterization capabilities and the deflection options enabled by each. 

5.2.3. Resonant Returns and Keyholes

A resonant return is created by the gravitational interaction an object during a preceding 

Earth encounter [12]. The highly publicized asteroid, Apophis, which currently passes by 

Earth about every 7 years but is not resonant [13], offers an example of orbit resonance. It 

will make close approaches to Earth in 2013, 2022, 2029, and 2036. 

If Apophis were to obtain a 426-day period (7:6 exterior resonance with the orbital period 

of Earth) due to a very specific gravitational interaction from the 2029 encounter, it 

would return to nearly the same point in space six revolutions and seven years later. At 

that point, Earth would also arrive at that point. Thus, the asteroid would have a resonant 

return in 2036. [14] A keyhole is the small area on an encounter’s target plane that the 

asteroid must pass through to collide on a subsequent encounter. Other examples of 

keyholes are discussed in Reference [15]. 

For the 2029 encounter, Apophis must pass through a keyhole that measures 600 meters 

wide, an occurrence that cannot be confirmed or eliminated by current observations. 

However, as additional measurements are made on each successive close approach, there 
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is a 95% probability that the 2036 impact will be ruled out after 2013 and a 99.8% 

probability that it will be ruled out in 2022. [15] 

Few objects have nearly resonant orbits that lend themselves to keyholes, but these few 

objects offer a particular challenge. Experts postulate that most resonant objects were 

“cleared” by Earth earlier in its 4.5-billion-year history. While additional information 

gained by each pass usually will confirm whether the object will miss Earth, if an object 

becomes resonant by passing through a keyhole, very little time (6 years in the case of 

Apophis) will usually be available to mitigate the threat 

Keyhole scenarios are expected to be “extremely rare,” less than 1% of the total possible 

impacts. This is due to the relatively small size of most keyholes compared with the 

cross-sectional area of the Earth. [16] Therefore, while the Apophis example has been 

well publicized, sophisticated keyhole scenarios such as this one appear to be the 

exception rather than the rule on human timescales. The scenario of deflecting Apophis 

(should it be necessary) is developed further in Section 6.13 and Reference [17]. 

5.3. Figures of Merit 

Figures of merit (FOM) are quantitative or qualitative metrics used to differentiate the 

cost, risk, performance, and other features of the detection, tracking, characterization, and 

deflection concepts and architectures. Metrics common to all concepts include life-cycle 

cost, development time, development risk, and mission risk. Individual metrics for 

performance include the capabilities of the Survey concepts, the relative capability of 

various characterization options, and capabilities of the deflection systems. 

Figure 7 illustrates the metrics used for the distinguishing Survey systems. Secondary 

figures of merit may be used to discriminate between two concepts with similar primary 

FOM results. Performance metrics for detection and tracking concepts include the 

percentage of PHOs detected by the end of 2020, and the time to detect 90% of the PHO 

population. 

Performance

— % of PHOs 

catalogued by 2020

— Time to catalog 90% 

of PHOs

— Precision orbit 

determination 

capability*

Cost

— Program Life 

Cycle Cost at 

70% confidence

Other Factors

— Science data 

acquired*

— Benefits to 

Exploration*

— Qualitative 

value of benefit 

to DoD*

Effectiveness Affordability

— Development Risk

Probability of 

fielding system in 

time to meet need

— Operations Risk

Probability of 

successfully 

operating systems 

to meet need

Risk

Risk Rated Performance Funding Profile

*Secondary FOM

Figure 7. Detection, Tracking, and Cataloguing Figures of Merit 

Risk FOMs are common to each concept. The concept risk FOM includes development 

and mission operations risk. Development risk is the probability that the system will be 

fielded in time to meet the requirement and was normalized for all schedules to the 70% 
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confidence level. Mission operations risk is the probability that the system will operate 

successfully once it has been fielded. 

The life-cycle cost (LCC) figure of merit is the life-cycle cost starting at the end of 

conceptual design and ending at the end of mission operations. Life-cycle costs are 

reported at a confidence level of 70%. The secondary FOMs include identification of 

technology developments or capabilities that may benefit either national science goals, 

NASA Exploration missions, or the Department of Defense (DoD). 

The characterization FOMs are shown in Figure 8. The risk, cost, and other factors are 

identical to those of the detection and tracking concepts, as are the effectiveness and 

affordability metrics. Performance for characterization is a qualitative assessment of how 

each concept informs the deflection options. Some characterization options provide 

robust information, while others provide limited amounts of data at lower confidence 

levels. 

Performance

— Capability that 

evaluates the threat

— Mass and orbit

— Capability that 

supports mitigation

— For “short warning”

scenarios

— For “routine 

warnings” of 

decades

Cost

— Program Life 

Cycle Cost at 

70% confidence

Other Factors

— Science data 

acquired*

— Benefits to 

Exploration*

— Qualitative 

value of benefit 

to DoD*

Effectiveness Affordability

— Development Risk

Probability of 

fielding system in 

time to meet need

— Operations Risk

Probability of 

successfully 

operating systems 

to meet need

Risk

Risk Rated Performance Funding Profile

*Secondary FOM

Figure 8. Characterization Figures of Merit 

Some deflection options require relatively accurate characteristics, while others require 

less information that may be less accurate. This metric illustrates how each 

characterization concept ranks relative to the needs of the deflection missions. 

Characterization is useful only in the context of this study as to how well it informs 

mitigation and its performance metrics are strongly tied to mitigation decisions. 

5.4. Development of Alternatives 

5.4.1. Overview

The study team developed a set of alternatives to evaluate the variation in performance, 

cost, schedule, and risk across the trade space. Alternatives included leveraging current 

assets, collaborating with proposed survey efforts of other government agencies, and 

pursuing new NASA-funded assets dedicated to detection and characterization goals. The 

set of alternatives that the study analyzed included concepts proposed by the private and 

academic sectors. These included presentations from the NEO public workshop (see 

Appendix D), concepts similar to those examined by the NEO Science Definition Team, 

and multiple-element architectures. While some survey systems have the capability to do 
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orbit follow-up and some object characterization, the goal of completing the Survey to a 

90% completeness level by the end of 2020 would prohibit extensive characterization by 

survey systems. After the search is completed and if these assets are still operational, it is 

possible that they could be used to achieve additional characterization requirements. 

5.4.2. Alternatives Definition Methodology

NEO study working groups developed the alternatives considered. Private-sector 

responses to a Request for Information (RFI) on NEO detection and characterization 

approaches and Internet and literature searches for related ideas augmented the Science 

Definition Team concepts. Furthermore, the study team developed architectures 

composed of combinations of multiple elements to investigate their benefits. 

5.4.3. Contributions of the Science Definition Team Report

In addition to the results noted in Section 4.5, the Science Definition Team (SDT) report 

[3] drew the following system comparisons and summaries of SDT results: 

• For a given limiting magnitude, the search system that covers the entire sky the 

quickest is the most efficient for discovering and cataloging undiscovered objects. 

• While locating observing systems in the Northern and Southern Hemispheres 

provides only a modest improvement in cataloging efficiency over collocated 

northern telescopes, north/south systems have increased warning efficiency for 

relatively rare “short warning” situations. 

• Search telescopes of the same type and aperture size located in low-Earth orbit 

(LEO) and at the second Lagrange point (exterior to the Earth on the sun-Earth 

line) have comparable cataloging efficiencies. 

• Detector systems in Venus-like orbits have several advantages over their groundbased counterparts and spacecraft located in the Earth’s vicinity: 

- A shorter orbital period allows them to observe more objects, more often, and 

closer to perihelion. 

- NEOs at heliocentric distances of 1 AU or less will appear brighter due to 

their fuller phases. 

- These systems have an increased ability to detect potentially underrepresented 

population of Aten and IEO asteroids. 

The SDT report ruled out space-based, infrared (IR) search systems because IR 

technology was not sufficiently mature. However, IR detector technologies and systems 

have progressed, and therefore were included in this study. 
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5.4.4. Alternatives Design Approach

Most alternatives were developed from existing designs or concepts. When existing 

designs or concepts were not available, the study team developed spacecraft-sizing 

algorithms to sufficiently develop the concepts for cost, risk, schedule, and performance 

analyses. 

5.5. Alternative Approaches to Meeting Search Requirements 

Broadly, the approaches to finding PHOs can be classified in three categories. Groundbased optical systems use large apertures to scan the sky at night for PHOs. Space-based 

optical systems gather visible light from vantage points near the Earth or in Venus-like 

heliocentric orbits. Space-based infrared systems operate from similar vantage points, and 

use passively cooled infrared detectors to find and track objects. The advantages and 

drawbacks of each system are discussed in the following sections. 

5.5.1. Ground-Based Optical Systems

Ground-based optical systems have several advantages over space-based systems. In 

general, ground-based systems are mostly based on mature technology (some have new 

focal planes) and are relatively easy to maintain and upgrade because they are easily 

accessible. Consequently, these systems can be implemented using a phased approach 

and may take advantage of shared software. This typically means that ground systems 

cost less to build, verify, operate, maintain and upgrade than their space-based 

counterparts. 

Because these optical systems must view through Earth’s atmosphere, ground systems 

have drawbacks. Ground-based optical systems cannot operate during daylight or twilight 

and are subject to interference from weather, atmospheric turbulence, scattering from 

moonlight, and atmospheric attenuation. These systems cannot easily operate close to the 

galactic plane because atmospheric turbulence and scattering cause source confusion. 

Significant atmospheric attenuation in the infrared-spectral region prevents these systems 

from accurately determining NEO sizes. These systems also will have difficulty finding 

objects in inner-Earth or Earth-like, orbits. They have fewer discovery opportunities 

because they are available only at the beginning and end of each night. Additionally, 

ground-based systems have intangible programmatic issues related to site and 

infrastructure maintenance. These issues are made worse if the telescopes are sited on 

foreign territory to achieve the best observing conditions and operate for decades. 

5.5.2. Space-Based Optical Systems

Space-based optical systems are based on mature technologies with a broad base of 

existing spacecraft mission heritage. Like ground-based systems, their advantages are 

primarily based on where they are put to use. Because these systems are in space they can 

access almost the entire sky at any given time with no interference from weather, daylight, 

moonlight, or atmospheric attenuation. Also, they can observe objects in inner-Earth or 

Earth-like orbits more easily than ground-based systems, especially if the detectors are 

located at Sun-Earth L1 or in a Venus-like orbit. 
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Beyond the fact that space-based systems are historically more expensive than groundbased systems, these systems offer there are several additional drawbacks. Getting a 

space-based system into place subjects it to possible launch and deployment failures and 

places it in a hostile environment that results in a shorter lifetime (7 to 10 years). This 

shorter lifetime is an important consideration if a NEO program is expected to continue to 

track objects for extended periods of time. In addition, they are dependent upon 

spacecraft-to-ground data links and unique onboard software. 

5.5.3. Space-Based Infrared Systems

With the exception of technology maturity, space-based infrared systems have the same 

advantages as space-based optical systems. For infrared systems this technology is 

maturing rapidly. Space-based, passively cooled infrared systems also have additional 

advantages. They require smaller apertures than optical systems of equal detection 

efficiency and provide more accurate estimates of object sizes. The object size 

uncertainties are less than 50% compared with 230% for visual detectors. A two-band 

infrared system could lower the size uncertainties to about 20%. These space-based 

systems also are much less affected by the problem of source confusion. There are about 

100 times fewer infrared sources per square degree at an infrared wavelength of 8 

microns compared with the number of visible sources at 0.5 microns. In addition, spacebased infrared systems have lower downlink data rate requirements than space-based 

visible detector systems. Space-based infrared systems were the most capable (sensitive) 

of the alternatives considered. 

Space-based infrared systems suffer from similar drawbacks as space-based optical 

systems. 

5.6. Detection and Tracking Survey Alternatives Considered 

Table 2, Table 3, and Table 4 list the detection, tracking, and cataloguing survey 

alternatives considered in this study. An initial feasibility assessment was performed on 

these concepts to produce a more manageable set of alternatives. The concepts not 

selected for further analysis are italicized and shaded in grey, and the rationale for not 

choosing them is described in Table 5. A corresponding alternatives trade tree is shown 

in Figure 9, and a pared trade tree is shown in Figure 10. 
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Table 2. Description of Ground-based Survey Alternatives Considered 

# Classification Concept Name Description 

1 Visible - Ground LINEAR Lincoln Near-Earth Asteroid Research (LINEAR) survey 

project using two 1m search telescopes. 

2 Visible - Ground NEAT Near-Earth Asteroid Tracking (NEAT) survey project using 

1.2m search telescope. 

3 Visible - Ground Pan-STARRS 1 Panoramic Survey Telescope And Rapid Response System 

(Pan STARRS) survey project using one 1.8m search 

telescope. 

4 Visible - Ground Spacewatch Spacewatch survey project using 1.8m and 0.9m search 

telescopes. 

5 Visible - Ground Spaceguard Combined ground-based detection efforts including 

LINEAR, NEAT, Catalina Sky, Spacewatch and LONEOS. 

6 Visible - Ground Catalina Sky 

Survey 

Catalina Sky Survey project using 0.7 and 1.5 m telescopes 

in the north and a 0.5 m in the southern hemisphere 

7 Visible - Ground VISTA Visible & Infrared Survey Telescope for Astronomy (VISTA) 

is a planned (2007) 4m astronomy telescope. 

8 Visible - Ground Shared LSST Large Synoptic Survey Telescope (LSST) is a planned 

(2014) 8m telescope, will spend 75% in survey mode. 

9 Visible - Ground DCT Discovery Channel Telescope (DCT) is a planned (2010) 

4.2m telescope, both broad and narrow band pass. 

10 Visible - Ground Shared PS4 Four 1.8m Pan-STARRS telescopes searching same spot of 

the sky at a time with an effective aperture of 3.6m; will 

spend 30% in survey mode. 

11 Visible - Ground SST Planned ground-based optical telescope. 

12 Visible - Ground Dedicated LSST Rebuild of Shared LSST, dedicated to NEO search. 

13 Visible - Ground Dedicated PS4 Rebuild of Shared PS4, dedicated to NEO search. 

14 Visible - Ground Dedicated PS8 Proposed system of two PS4 telescopes searching same area 

of the sky at a time with an effective aperture of 5.1m. 

15 Visible - Ground Dedicated PS16 Proposed system of two PS8 telescopes (North and South or 

equator) searching differing regions sky regions at any 

given time, thus doubling the search area. 

36 Visible – Ground LONEOS Lowell Observatory near-Earth Object search program 

using a 0.6 m telescope 

33 Radar - Ground Arecibo Arecibo Radio Telescope is an operational ground-based 

radio telescope with a 305m fixed dish. May be closed by 

2011 without additional support beyond current National 

Science Foundation funding. 

34 Radar - Ground Goldstone Goldstone is an operational ground-based radio telescope 

with a 70m steerable dish. Similar or enhanced capability 

provided if Canberra 70m radar adopted for the Deep 

Space Network. 

35 Radar - Ground Bistatic 100m Proposed system composed of two 100m steerable radio 

antennas to be operated in a bistatic mode. 

* Italicized and shaded concepts were not considered for further detection and tracking analysis. 

Radars were analyzed for their contribution to precision orbit determination capability. 
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Table 3. Description of Space-based Survey Alternatives Considered 

# Classification Concept Name Description 

16 Visible - Space Hubble Hubble Space Telescope is an operational 2.4m optical, 

space-based astronomy telescope. 

17 Visible - Space Kepler Kepler is a planned (2009) 0.95m space-based astronomy 

telescope dedicated to the search of extra-solar planets. 

18 Visible - Space 1m Vis 

LEO/L1/L2 

Concept for a space based 1m optical search telescope in 

the near-Earth region (LEO/L1/L2). 

19 Visible - Space 2m Vis 

LEO/L1/L2 

Concept for a space based 2m optical search telescope in the 

near-Earth region (LEO/L1/L2). 

20 Visible - Space 1m Vis Venuslike Orbit Concept for a space based 1m optical search telescope in 

Venus-like orbit (Heliocentric ~0.7AU from Sun). 

21 Visible - Space 2m Vis Venuslike Orbit Concept for a space based 2m optical search telescope in 

Venus-like orbit (Heliocentric ~0.7AU from Sun). 

22 Visible - Space 3m Vis 

LEO/L1/L2 

Concept for a space based 3m optical search telescope in 

the near-Earth region (LEO/L1/L2). 

23 Visible - Space 4-6m Segmented 

Vis 

Concept for a space based 4-6m segmented, optical search 

telescope in the near-Earth region (LEO/L1/L2). 

24 Infrared - Space Spitzer Spitzer Space Telescope is an operational 0.85m IR, spacebased astronomy telescope. 

25 Infrared - Space JWST James Webb Space Telescope (JWST) is a planned (2013) 

6.5m segmented, IR, space-based astronomy telescope. 

26 Infrared - Space WISE Wide-field Infrared Survey Explorer (WISE) is a planned 

(2010) 0.4cm IR, space-based, non-pointable survey 

telescope. 

27 Infrared - Space 0.5m IR LEO Concept for a space based 0.5m IR search telescope in low 

Earth orbit (LEO). 

28 Infrared - Space 1m IR LEO Concept for a space based 1m IR search telescope in low 

Earth orbit (LEO). 

29 Infrared - Space 0.5m IR L1/L2 Concept for a space based 0.5m IR search telescope at SunEarth Lagrange point (L1/L2). 

30 Infrared - Space 1m IR L1/L2 Concept for a space based 1m IR search telescope at SunEarth Lagrange point (L1/L2). 

31 Infrared - Space 0.5m IR Venuslike Orbit Concept for a space based 0.5m IR search telescope in 

Venus-like orbit (Heliocentric ~0.7AU from Sun). 

32 Infrared - Space 1m IR Venuslike Orbit Concept for a space based 1m IR search telescope in Venuslike orbit (Heliocentric ~0.7AU from Sun). 

* Italicized and shaded concepts were not considered for further analysis 

Table 4. Description of Data Management Alternatives Considered 

# Classification Concept Name Description 

36 Ops and Data 

Management 

Scale Existing 

Systems 

Expand Minor Planet Center (MPC) capability to support 

expected increases in NEO detection rates. 

37 Ops and Data 

Management 

Adopt Other 

Systems 

Adopt system like Futron's Space Launch & Satellite 

Database, Aerospace Corp.'s Space Systems Engineering 

Database or Analytical Graphics Inc.'s Satellite Database. 

38 Ops and Data 

Management 

New Central 

Repository 

Proposed framework for a US National Virtual Observatory 

(NVO). 

39 Ops and Data 

Management 

Back-up Facility Grow the MPC capability as the detection rate grows using 

the NVO as a backup archive. 
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5.6.1. Survey Alternatives Not Analyzed and Rationale and Why

Table 5 describes the search alternatives that the study considered, but did not analyze. 

The study assumed that the Spaceguard assets would continue to operate as a baseline 

capability at least until the proposed Survey could begin. It is assumed that the new 

Survey will build from this capability and that existing assets and processes will continue 

until replaced. While the Spaceguard goal is to detect objects 1 km or larger, some 

objects as small as 140 meters are detected under favorable viewing conditions. 

The detection of PHOs using ground-based infrared (IR) techniques is not considered a 

viable option because Earth’s atmosphere radiates or absorbs strongly in the IR region. 

Although IR characterization of bright PHOs is possible from a few select, high-altitude 

observatories (e.g., NASA’s Infrared Telescope Facility or Keck), even the 10-meter 

Keck telescope only can observe objects much brighter than a 1-km PHO at typical 

observing distances, and therefore is of little use in the detection of smaller objects. 

The returned signal strength of radar decreases as inverse fourth power of distance, and 

therefore has a range that is insufficient to observe most PHOs. A typical radar beam is 

only 1 arc minute in diameter, which does not provide the sky coverage necessary for 

detection. Therefore, radars are not a viable search system for PHOs, but may be 

desirable to rapidly improve the orbital estimates of some objects. 

Classified DoD space surveillance systems were evaluated for contributions that they 

might make to the search for PHOs. The capabilities and requirements for six current and 

proposed DoD systems were studied, and this report has concluded that their capabilities, 

operations, and data-storage procedures are largely incompatible with Survey goals. In 

addition, PHO survey systems are not expected to contribute materially to satisfying any 

DoD requirements evaluated unless Survey goals are compromised substantially. Further 

discussion of the capabilities that might be shared with DoD missions and systems are 

discussed in Appendix O.

5.6.2. Survey Alternatives Analyzed

The alternatives analyzed in this study were developed to a level of detail to cover 

architectural and technology options. These options included combinations of space- and 

ground-based detection methods in both the visible- and infrared-spectral regions. These 

concepts have not been as optimized as they would be in a detailed design study. 

Figure 11 provides a matrix of the multi-element alternatives evaluated. A more detailed 

description of these elements appears in Appendix E. The “baseline” multi-element 

architecture represents a collection of ground-based assets that are currently operational 

(i.e. Spaceguard), as well as non NASA-funded survey observatories designed to conduct 

surveys that are in development by agencies other than NASA and that are expected to 

become operational before 2020. This architecture provides a basic capability that is 

likely to be available regardless of a dedicated NASA survey, and thus will provide a 

baseline upon which NASA could build any additional capability needed to meet the 

Survey goals. Elements besides Spaceguard are the Pan-STARRS (PS4), being developed 

by the U.S. Air Force, and the Large Synoptic Survey Telescope (LSST) planned by the 

National Space Foundation. Both agencies have indicated a willingness to share time on 

the assets in exchange for partial funding of operations. 
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2006 Near-Earth Object Survey and Deflection Study 

40 DRAFT Pre-Decisional Material

Figure 9. Complete Survey Alternatives Trade Tree 

Detect/Track/Catalog Detect/Track/Catalog

Visible Visible Infrared Infrared Radar Radar

Ground Based Ground Based Space Based Space Based

Existing Existing

Proposed Proposed

LINEAR LINEAR

NEAT NEAT

Pan-STARRS 1 Pan-STARRS 1

Spacewatch Spacewatch

Spaceguard Spaceguard

Catalina Sky Survey Catalina Sky Survey

VISTA VISTA

New Asset New Asset

Shared LSST Shared LSST

DCT DCT

Shared PS4 Shared PS4

SST SST

Dedicated LSST Dedicated LSST

Dedicated PS4 Dedicated PS4Dedicated PS8 Dedicated PS8

Space Based Space Based

Existing Existing

Proposed Proposed

New Asset New Asset

Spitzer Spitzer

JWST JWST

WISE WISE

Existing Existing

Arecibo Arecibo

New Asset New Asset

Goldstone Goldstone

Bistatic 100m Bistatic 100m

Existing Existing

Proposed Proposed

New Asset New Asset

Hubble Hubble

Kepler Kepler

1m Vis LEO/L1/L2 1m Vis LEO/L1/L2

2m Vis LEO/L1/L2 2m Vis LEO/L1/L2 0.5m IR LEO 0.5m IR LEO

1m IR L1/L2 1m IR L1/L2

1m Vis Venus-like 1m Vis Venus-like

Dedicated PS16 Dedicated PS16

2m Vis Venus-like 2m Vis Venus-like

0.5 IR Venus-like 0.5 IR Venus-like

1m IR Venus-like 1m IR Venus-like

3m Vis LEO/L1/L2 3m Vis LEO/L1/L2

4-6m Segmented Vis 4-6m Segmented Vis

0.5m IR L1/L2 0.5m IR L1/L2

1m IR LEO 1m IR LEO

Ops./Data Mgt. Ops./Data Mgt.

Scale Existing Systems Scale Existing Systems

Adopt Other Similar 

Systems

Adopt Other Similar 

Systems

New Central Repository New Central Repository

Back-up Facility Back-up Facility

Key

= eliminated from further consideration

Detect/Track/Catalog Detect/Track/Catalog

Visible Visible Infrared Infrared Radar Radar

Ground Based Ground Based Space Based Space Based

Existing Existing

Proposed Proposed

LINEAR LINEAR

NEAT NEAT

Pan-STARRS 1 Pan-STARRS 1

Spacewatch Spacewatch

Spaceguard Spaceguard

Catalina Sky Survey Catalina Sky Survey

VISTA VISTA

New Asset New Asset

Shared LSST Shared LSST

DCT DCT

Shared PS4 Shared PS4

SST SST

Dedicated LSST Dedicated LSST

Dedicated PS4 Dedicated PS4Dedicated PS8 Dedicated PS8

Space Based Space Based

Existing Existing

Proposed Proposed

New Asset New Asset

Spitzer Spitzer

JWST JWST

WISE WISE

Existing Existing

Arecibo Arecibo

New Asset New Asset

Goldstone Goldstone

Bistatic 100m Bistatic 100m

Existing Existing

Proposed Proposed

New Asset New Asset

Hubble Hubble

Kepler Kepler

1m Vis LEO/L1/L2 1m Vis LEO/L1/L2

2m Vis LEO/L1/L2 2m Vis LEO/L1/L2 0.5m IR LEO 0.5m IR LEO

1m IR L1/L2 1m IR L1/L2

1m Vis Venus-like 1m Vis Venus-like

Dedicated PS16 Dedicated PS16

2m Vis Venus-like 2m Vis Venus-like

0.5 IR Venus-like 0.5 IR Venus-like

1m IR Venus-like 1m IR Venus-like

3m Vis LEO/L1/L2 3m Vis LEO/L1/L2

4-6m Segmented Vis 4-6m Segmented Vis

0.5m IR L1/L2 0.5m IR L1/L2

1m IR LEO 1m IR LEO

Ops./Data Mgt. Ops./Data Mgt.

Scale Existing Systems Scale Existing Systems

Adopt Other Similar 

Systems

Adopt Other Similar 

Systems

New Central Repository New Central Repository

Back-up Facility Back-up Facility

Key

= eliminated from further consideration
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2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 41 

Table 5. Rationale for Detection and Tracking Alternatives not Analyzed 

Concept Name Downselect Criteria 

LINEAR Cannot detect sufficient number of 140m NEOs. Included in 

Spaceguard 

LONEOS Cannot detect sufficient number of 140m NEOs. Included in 

Spaceguard 

NEAT Cannot detect sufficient number of 140m NEOs. Included in 

Spaceguard 

Pan-STARRS 1 Cannot detect sufficient number of 140m NEOs. 

Spacewatch Cannot detect sufficient number of 140m NEOs. Included in 

Spaceguard 

Catalina Sky Survey Cannot detect sufficient number of 140m NEOs. Included in 

Spaceguard 

VISTA Not a survey telescope. Shared time asset. Foreign owned. 

DCT Cannot detect significant number of 140m NEOs. Shared time 

asset. 

SST Classified project, information is unavailable. Highly 

constrained asset access. 

Hubble Not a survey telescope. Highly constrained asset access. 

Kepler Cannot detect significant number of 140m NEOs. Will only 

observe 0.25% of entire sky. Highly constrained asset access. 

1m Vis LEO/L1/L2 Poor performance shown in initial analysis. 

3m Vis LEO/L1/L2 Initial analysis shows no need for telescope as large. High 

cost/risk. 

4-6m Segmented Vis Not a survey telescope. Initial analysis shows no need for 

telescope as large. High cost/risk. 

Spitzer Not a survey telescope. Short lifetime. 

JWST Not a survey telescope. Highly constrained asset access. 

WISE Highly constrained asset access. Short life. 

0.5m IR LEO Earth’s thermal radiation 

1m IR LEO Earth’s thermal radiation 

Arecibo Not a survey telescope. Signal drop-off limits detection to 

within 0.3 AU 

Goldstone Not a survey telescope. Signal drop-off limits detection to 

within 0.1 AU 

Bistatic 100m Not a survey telescope. Signal drop-off limits detection to 

within 0.3 AU 
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2006 Near-Earth Object Survey and Deflection Study 

42 DRAFT Pre-Decisional Material

Figure 10. Pared Survey Alternatives Trade Tree 

Detect/Track/Catalog Detect/Track/Catalog

Visible Visible Infrared Infrared

Ground Based Ground Based Space Based Space Based

Existing Existing

Proposed Proposed

D-1: Spaceguard D-1: Spaceguard

New Asset New Asset

D-2: Shared LSST D-2: Shared LSST

D-3: Shared PS4 D-3: Shared PS4

D-4: Dedicated LSST D-4: Dedicated LSST

D-5: Dedicated PS4 D-5: Dedicated PS4D-6: Dedicated PS8 D-6: Dedicated PS8

Space Based Space Based

New Asset New Asset New Asset New Asset

D-8: 2m Vis LEO/L1/L2 D-8: 2m Vis LEO/L1/L2

D:12: 1m IR L1/L2 D:12: 1m IR L1/L2 D-9: 1m Vis Venus-like D-9: 1m Vis Venus-like

D-7: Dedicated PS16 D-7: Dedicated PS16

D-10: 2m Vis Venus-like D-10: 2m Vis Venus-like D-13: 0.5 IR Venus-like D-13: 0.5 IR Venus-like

D-14: 1m IR Venus-like D-14: 1m IR Venus-like

D-11: 0.5m IR L1/L2 D-11: 0.5m IR L1/L2

Ops./Data Mgt. Ops./Data Mgt.

D-15: Scale Existing 

Systems

D-15: Scale Existing 

Systems

D-16: Adopt Other Similar 

Systems

D-16: Adopt Other Similar 

Systems

D-17: New Central 

Repository

D-17: New Central 

Repository

D-18: Back-up Facility D-18: Back-up Facility

Detect/Track/Catalog Detect/Track/Catalog

Visible Visible Infrared Infrared

Ground Based Ground Based Space Based Space Based

Existing Existing

Proposed Proposed

D-1: Spaceguard D-1: Spaceguard

New Asset New Asset

D-2: Shared LSST D-2: Shared LSST

D-3: Shared PS4 D-3: Shared PS4

D-4: Dedicated LSST D-4: Dedicated LSST

D-5: Dedicated PS4 D-5: Dedicated PS4D-6: Dedicated PS8 D-6: Dedicated PS8

Space Based Space Based

New Asset New Asset New Asset New Asset

D-8: 2m Vis LEO/L1/L2 D-8: 2m Vis LEO/L1/L2

D:12: 1m IR L1/L2 D:12: 1m IR L1/L2 D-9: 1m Vis Venus-like D-9: 1m Vis Venus-like

D-7: Dedicated PS16 D-7: Dedicated PS16

D-10: 2m Vis Venus-like D-10: 2m Vis Venus-like D-13: 0.5 IR Venus-like D-13: 0.5 IR Venus-like

D-14: 1m IR Venus-like D-14: 1m IR Venus-like

D-11: 0.5m IR L1/L2 D-11: 0.5m IR L1/L2

Ops./Data Mgt. Ops./Data Mgt.

D-15: Scale Existing 

Systems

D-15: Scale Existing 

Systems

D-16: Adopt Other Similar 

Systems

D-16: Adopt Other Similar 

Systems

D-17: New Central 

Repository

D-17: New Central 

Repository

D-18: Back-up Facility D-18: Back-up Facility
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2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 43 # 

Year of First Observations All programs start 10/1/2007 

 Spaceguard 

 Shared PS4 

 Dedicated PS4 

 PS8 

 PS16 

 Shared LSST 

 Dedicated LSST 

 0.5m IR @ L1 

 1m IR @ L1 

 0.5m IR in Venus-like Orbit 

 1m IR in Venus-like Orbit 

 1m Vis in Venus-like Orbit 

 2m Vis in Venus-like Orbit 

D-20 Spaceguard+Shared PS4 1998 2010 D-22 Spaceguard + Dedicated PS4 2000 2012 D-23 Spaceguard + Dedicated PS8 2000 2012 D-24 Spaceguard + Dedicated PS16 2000 2012 D-19 Spaceguard + Shared LSST 2000 2014 D-21 Spaceguard + Dedicated LSST 2000 2016 D-25 Baseline (Shared PS4/LSST) 2000 2010 2014 D-27 Baseline + Dedicated PS8 2000 2010 2012 2014 D-28 Baseline + Dedicated PS16 2000 2010 2012 2014 D-36 Baseline + Dedicated LSST 

Ground-based only 

2000 2010 2014 2016 

D-32 Baseline + 0.5m IR @ L1 2000 2010 2014 2012 D-33 Baseline + 1m IR @ L1 2000 2010 2014 2014 D-34 Baseline + 0.5m IR in Venus-like 2000 2010 2014 2014 D-35 Baseline + 1m IR in Venus-like 2000 2010 2014 2014 D-30 Baseline + 1m Vis in Venus-like 2000 2010 2014 2014 D-31 Baseline + 2m Vis in Venus-like 

Ground + Space 2000 2010 2014 2014 

Figure 11. Multi-Element Detection, Tracking, and Cataloguing Systems Analyzed 
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2006 Near-Earth Object Survey and Deflection Study 

44 DRAFT Pre-Decisional Material 

5.7. Survey System Elements Schedules 

To assess whether the alternatives could meet the congressionally directed goal of 90% 

completeness by the end of 2020, schedules for each of the search alternatives were 

developed. Figure 12 displays the nominal development time for the detection, tracking, 

and cataloguing system concepts. This time does not include issues related to the 

development of shared systems, and assumes that all programs are fully funded during 

development and testing. 

0.0

1.5

4.5

6.0

6.0

5.5

6.5

5.0

6.0

5.0

6.0

7.0

7.0

6.0

7.5

0 1 2 3 4 5 6 7 8 9 10

Existing

Shared PS4

Dedicated PS4

PS 8

PS 16

Shared LSST

Dedicated LSST

0.5m IR L1/L2

1.0m IR L1/L2

0.5m IR Venus-Like Orbit

1.0m IR Venus-Like Orbit

2.0m Visible LEO

2.0m Visible L1/L2

1.0m Vis Venus-Like Orbit

2.0m Vis Venus-Like Orbit

Development Schedule (Years)

Figure 12. Detection and Tracking Nominal Development Schedules 

For space assets, transit and checkout time must be added to estimate initial operational 

capability (IOC). Missions traveling to a Venus-like heliocentric orbit used a Venus 

swingby to limit required launch performance. This yielded one launch opportunity about 

every 13 months and caused the 0.5-meter IR system to delay its launch by about 1 year. 

If a more capable launch system is assumed, the launch delay due to a Venus swingby 

could be eliminated. 

For missions stationed at Sun-Earth L1, a 3-month transit was assumed based on the 

Genesis, ACE, SOHO, and WMAP missions. For missions in a Venus-like heliocentric 

orbit, a 5-month transit was assumed (Magellan, Venera, and Pioneer Venus). For 

ground-based missions, checkout time was based on nominal program schedules with 

credit given if similar hardware had been calibrated previously. For all space missions, a 

checkout time of 2 months was assumed. The development of all search systems is 

assumed to start at the beginning of FY08. Initial operational dates are shown in Table 6 

and Table 7. 
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2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 45 

Table 6. Start Date for Ground-based Survey Elements 

(Months)

Asset Develop Check IOC* 

Existing Assets 0 0 1998 

Shared PS4 19 6 2010 

Dedicated PS4 56 4 2013 

PS 8 73 4 2014 

PS 16 73 4 2014 

Shared LSST 67 10 2014 

Dedicated LSST 80 7 2015 

 * Initial Operational Capability 

Table 7. Start Date for Space-based Survey Elements 

(Months) 

Asset Develop Launch Transit Check IOC* 

0.5m IR @ L1 61 0 3 2 2013 

1.0m IR @ L1 71 0 3 2 2014 

0.5m IR @ Venus 61 13 5 2 2013 

1.0m IR @ Venus 70 3 5 2 2014 

1.0m Vis @ Venus 72 0 3 2 2014 

2.0m Vis @ Venus 86 0 3 2 2016 

 * Initial Operational Capability 

5.8. Survey Performance Simulation Results 

The results presented in this section are based on the analyses described in Appendix 

Section H.2. Absolute performance is expected to be within 5% of the results shown, and 

within 1-2% for the concepts that achieve 90%. These results are expected to have a 

relative uncertainty of 1-2%. 

Table 8 shows the analysis results for the performance of various ground-based survey 

options acting alone and assuming no discoveries until the beginning of the surveys. For 

example, the first line of Table 8 shows that the currently operating ground-based 

Spaceguard observing program will catalog 14% of all PHOs with diameters larger than 

140 meters (D > 140 m) by the end of 2020. This system would take decades beyond 

2030 to achieve the goal of 90% completeness. On the second line of Table 8, note that 

the shared PS4 will reach 72% survey completeness for 140-meter size objects by the end 

of 2020. It will reach 69% completeness after operating for 10 years and 90% 

completeness after 2030. 

Table 9 shows the analysis results for the performance of the space-based systems acting 

alone. The first line of Table 9 shows that a 0.5-meter IR telescope operating at the SunEarth L1 point beginning in 2013 could catalog 85% of the PHO population (D > 140 m) 

by the end of 2020, 88% after 10 years (2023) and 90% shortly thereafter. 
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2006 Near-Earth Object Survey and Deflection Study 

46 DRAFT Pre-Decisional Material 

Table 8. Ground-based Survey Performance 

140 meter PHO Completion 

Survey Systems* by end of 2020 10 years Year for 90% 

Spaceguard 14% 8% >>2030 

PS4 (shared) 72% 69% >2030 

PS4 (dedicated) 72% 77% >2030 

PS8 (dedicated) 74% 81% >2030 

PS16 (dedicated) 77% 83% 2029 

LSST (shared) 75% 81% >2030 

LSST (dedicated) 85% 90% 2024 

 * Continued operation of Spaceguard assets add marginally to performance

Table 9. Space-based Survey Performance 

140 meter PHO Completion 

Survey Systems* by end of 2020 10 years Year 90% 

0.5m IR @ L1 85% 88% 2024 

1.0m IR @ L1 86% 91% 2022 

0.5m IR in Venus-like 89% 93% 2021 

1.0m IR in Venus-like 92% 95% 2020 

1.0m VIS in Venus-like 82% 88% 2025 

2.0m VIS in Venus-like 87% 94% 2022 

 Requirement is 90% completion by end of 2020 

Table 10 shows space-based alternatives used in conjunction with the baseline groundbased systems. The first line of Table 10 shows that these systems could reach the 90% 

goal by the end of 2020 if the Spaceguard telescopes operate, the shared PS4 begins 

operations in 2010, the shared LSST begins operations in 2014, and the space-based IR 

telescope at Sun-Earth L1 starts in 2013. 

Table 11 shows that the options that exceed congressional goals also provide other 

benefits. The middle column of the table shows that systems that operate in Venus-like 

orbits are more efficient at finding Aten and IEOs, a potentially underrepresented 

population of PHOs. The final column of the table shows that for combinations of visual 

and IR detectors, some systems will be able to estimate object sizes to better than 50% 

for more than 70% of the catalog by the end of 2020. 
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2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 47 

Table 10. Survey Performance of Combinations 

140 meter PHO Completion 

Survey Systems* by end of 2020 Year 90% 

Shared PS4 + Shared LSST (Baseline) 83% 2026 

Dedicated PS8 + Baseline 85% 2024 

Dedicated LSST + Baseline 90% 2020 

0.5m IR @ L1 + Baseline 91% 2020 

1.0m IR @ L1 + Baseline 91% 2020 

0.5m IR in Venus-like + Baseline 97% 2017 

1.0m IR in Venus-like + Baseline 97% 2017 

1.0m VIS in Venus-like + Baseline 93% 2019 

2.0m VIS in Venus-like + Baseline 95% 2018 

 * Requirement is 90% by the end of 2020 

Table 11. Additional Benefits of Space Systems 

Survey Systems 

Diameter 

to 90%* Size

Sees More 

Atens/IEOs 

0.5m IR @ L1 + Baseline 125 m 77% 9

1.0m IR @ L1 + Baseline 125 m 78% 9

1.0m IR in Venus-like 125 m ** 99

0.5m IR in Venus-like + Baseline 90 m - 99

1.0m IR in Venus-like + Baseline 80 m 76% 99

1.0m VIS in Venus-like + Baseline 125 m 78% 99

2.0m VIS in Venus-like + Baseline 110 m - 99

 * Diameter of PHOs catalogued to 90% by the end of 2020 

** Not explicitly evaluated, likely 70-78% 

Apart from setting the discovery floor for the next-generation search systems, the existing 

Spaceguard system does not materially contribute to future searches because any objects 

it finds would more be found more quickly by one of the next-generation search systems. 

Assuming that a single system can cover the richest areas of the sky in one search period 

(around 5 days), the addition of another identical system generally adds very little 

because it, too, cannot see fainter objects. Asteroids tend to cycle in and out of view on 

timescales that would allow either of the two identical telescopes to “discover” a given 

NEO if it were observable at all. However, a second system may enable an improvement 

in the orbit quality because the two acting together would enable more observations and 

consequently provide a more precise estimation. 

Figure 13 shows the survey completion as a function of time for several individual and 

combinations of systems. The results for the PHO completeness percentages on the 

following pages are estimated to be accurate to ±2% for results near 90%. Performance 

results for additional detection and tracking combinations are reported in Appendix I. 
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2006 Near-Earth Object Survey and Deflection Study 

48 DRAFT Pre-Decisional Material 

0%

10%

20%

30%

40%

50%

60%

70%

80%

90%

100%

2008 2012 2016 2020 2024 2028

Calendar Year

Percent D>140m PHO Population Catalogued

Existing Assets (in all options)

Shared LSST

Dedicated LSST

Shared PS4 + Shared LSST (Baseline)

Baseline + Dedicated LSST

Baseline + 0.5m IR @ L1

Baseline + 0.5m IR in Venus-like Orbit

GOAL

Figure 13. Survey Performance for Selected Alternative Systems 

Figure 14 shows the various schedule elements that contribute to the completion date of 

the ground-based element alternatives. 

2008 2012 2016 2020 2024 2028 2032

Existing Assets

Shared PS4

Dedicated PS4

PS 8

PS 16

Shared LSST

Dedicated LSST

Calandar Year

Development Time

Checkout Time

Time to 90% Complete

Figure 14. Constituents of Ground-Based Survey Schedules to 90% 

Figure 15 shows the schedule elements that contribute to the completion date of the 

space-based alternatives. To reduce launch requirements, a Venus flyby was used to 

reach the Venus-like heliocentric orbits. This approach limited launch opportunities to 

about one every 13 months. How these launch opportunities may affect completion dates 

is illustrated in Figure 16. 
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DRAFT Pre-Decisional Material 49 

2008 2012 2016 2020 2024 2028 2032

0.5m IR L1/L2

1.0m IR L1/L2

0.5m IR Venus-Like

Orbit

1.0m IR Venus-Like

Orbit

2.0m Visible LEO

2.0m Visible L1/L2

1.0m Vis Venus-Like

Orbit

2.0m Vis Venus-Like

Orbit

Calandar Year

Development Time

Time to Launch

Tranist Time

Checkout Time

Time to 90% Complete

Figure 15. Constituents of Space-Based Survey Schedules to 90% 

Since start dates and acquisition schedules cannot be predicted to within a year at this 

juncture, the completion dates for Venus-like orbits should be considered to vary by up to 

an additional year over the other alternatives. This is due to limited launch dates, unless a 

higher-performing launch vehicle is used to eliminate the Venus swingby requirement. 

2008 2012 2016 2020 2024 2028

0.5m IR Venus-Like

Orbit

1.0m IR Venus-Like

Orbit

1.0m Vis Venus-Like

Orbit

2.0m Vis Venus-Like

Orbit

Venus Flyby Launch

Constraint

Calandar Year

Development Time

Time to Launch

Tranist Time

Checkout Time

Time to 90% Complete

... ...

13 month slip in completion 

date if lauch window is missed

Figure 16. Possible Effect of Venus Flyby on 90% Completeness Dates 

Figure 17 shows how combinations of ground-based systems contribute to reaching the 

90% completeness goal. Note that existing assets in addition to Shared PS4, Shared LSST, 

and a Dedicated LSST reach the 90% survey goal by the end of 2020. 
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50 DRAFT Pre-Decisional Material 

2008 2012 2016 2020 2024 2028 2032

Existing Only (Spaceguard)

Shared PS4 + Existing

Dedicated PS4 + Existing

Dedicated PS8 + Existing

Dedicated PS16 + Existing

Shared LSST + Existing

Dedicated LSST + Existing

Shared PS4+Shared LSST + Existing*

Dedicated PS4 + Baseline*

Dedicated PS8 + Baseline*

Dedicated PS16 + Baseline*

Dedicated LSST + Baseline*

Calandar Year

1st Architecture Element

2nd Architecture Element

3rd Architecture Element

Last Element to 90% Complete

 *Baseline = Existing + 

 Shared PS4 + 

 Shared LSST 

Figure 17. Survey Performance of Ground-Based Combinations 

Figure 18 shows how combinations of ground-based and space-based systems contribute 

to reaching the 90% completeness goal. Note that except for the near-Earth visible 

alternatives, all of these combinations meet the goal by the end of 2020. 

2008 2012 2016 2020 2024 2028 2032

0.5m IR L1/L2 +

Baseline*

1.0m IR L1/L2 +

Baseline*

0.5m IR Venus +

Baseline*

1.0m IR Venus +

Baseline*

2.0m Visible LEO +

Baseline*

2.0m Visible L1/L2 +

Baseline*

1.0m Visible Venus +

Baseline*

2.0m Vis Venus +

Baseline*

Calandar Year

1st Architecture Element

2nd Architecture Element

3rd Architecture Element

Last Element to 90% Complete

 *Baseline = Existing + 

 Shared PS4 + 

 Shared LSST 

Figure 18. Survey Performance of Ground and Space Based Combinations 
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DRAFT Pre-Decisional Material 51 

5.9. Orbital Uncertainty 

After detection, the most important parameter affecting mitigation decisions is orbit 

uncertainty. If decision makers had perfect knowledge of an object’s orbit and a basic 

understanding of its size, mitigation decisions would be much clearer. Due to radar’s 

inability to provide active ranging at distances beyond 0.3 AU (45 million km), 

determining the orbits of most PHOs is limited to optical means. 

Orbit uncertainty is relatively small during observations because the measurements 

constrain the object’s position during this period. When the object’s position is predicted 

into the future, however, orbit uncertainties grow; the farther one predicts into the future, 

the greater the uncertainty. The rate of uncertainty depends largely on how long the 

object is observed. If the object were observed for only a short period, a few weeks or 

months, say, and if radar-tracking data were not available, the orbit uncertainty would 

grow rapidly over the prediction. Unless the time to impact was very short, the encounter 

error ellipse could be hundreds or thousands of times larger than the Earth, leading to a 

small impact probability. 

When an object has been observed for several orbital periods (1-10) or when radartracking data are available, the uncertainty in predicting the orbit is reduced substantially. 

Figure 19 from Reference [17] illustrates how prediction accuracy improves as the length 

of the interval of observation grows. 

0.001

0.01

0.1

1

10

100

1000

10000

0.001 0.01 0.1 1 10

Orbit Periods Observed

Orbit Error per year (Earth Radii)

Figure 19. Orbit Prediction Accuracy Improves with Length of Observation 

Currently, for a single apparition observed by an optical asset, meaningful predictions can 

be accurate for decades. At that point, the uncertainty region is generally stretched out 

more than ±0.1 AU and it is not clear that an encounter is real. If radar data are obtained, 

encounter predictions can extend to centuries. Once a full orbit is sampled (two or more 

apparitions), the meaningful predictions jump to many centuries on average, regardless of 

whether radar data are available. Once the survey is complete, almost every object will 

have multiple apparitions of data, and thus, for any random PHO, analysts could predict 

its location hundreds of years into the future. This would provide providing centuries of 
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2006 Near-Earth Object Survey and Deflection Study 

52 DRAFT Pre-Decisional Material 

warning if the object poses a threat, assuming close planetary encounters (rare for any 

given object on human timescales) do not occur in the interim. [8] 

After an object has been observed for several years, or if the object has approached close 

enough to the Earth to be observed with radar, the uncertainty in its orbit is reduced as 

described in Section 5.10.3. The uncertainty in the object’s orbit then grows at a much 

slower rate when its position is predicted ahead of time, and the error ellipse will be 

comparable to or smaller than the size of the Earth decades into the future. The slow error 

growth for these objects also will allow reliable encounter predictions to extend to more 

than a century. Thus, once the survey is complete, a 100-year encounter prediction will 

be as accurate as a 10-year prediction is today assuming close planetary encounters do 

not occur. [8] 

Much has been written on the non-gravitational forces affecting asteroid orbits. Most 

forces (excluding random forces such as outgassing of active bodies) can be modeled to 

some extent given enough observations. Even if these forces cannot be modeled, the 

effect of non-gravitational forces on objects that are 100 meters to 1 km in size does not 

materially affect impact prediction uncertainty within 50 years. [19] [20] [21] [22] 

The approximate values of radial and transversal accelerations, which affect bodies in the 

10-cm to 10-km size range and when solar gravity is scaled to unity, are shown in Table 

12. For comparison, typical gravitational perturbations by planets and big asteroids are 

GMplanet ~ 10-3 and GMasteroid < 10-9. Among the most difficult forces to predict is the 

Yarkovsky/YORP Effect, a force created by re-radiation of photons from the visible solar 

flux of a rotating object. As developed in Reference [19], the maximum effect of the 

Yarkovsky/YORP Effect on the semimajor axis of an asteroid 1-km in diameter is about 

10 meters per year, and therefore constitutes a second-order effect on impact prediction 

uncertainty. Additional uncertainty accumulates due to the uncertainty in the mass of 

other celestial objects, but this effect is generally small compared with those described in 

Table 12. 

Table 12. The Magnitude of Various Forces Acting on Asteroids 

IAU Symposium No. 229, 2005. Non-gravitational forces acting on small bodies.

Broz, Vokrouhlicky, Bottke, Nesvorny, Morbidelli and Capek

It is a derived requirement that any deflection approach have some type of effectiveness 

assessment (i.e. post-action orbit determination) because predicting the object’s new orbit 

using remote methods likely will require many years. This delay in assessing the 

effectiveness of the deflection attempt likely will not be acceptable unless the deflection 

attempt occurred many decades ahead of the predicted impact. These missions are similar 

to the in-situ orbit determination alternatives discussed in Section 5.10.4. 
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2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 53 

5.10. Performance of Orbit Determination Alternatives 

5.10.1. Precision Orbit Determination Using Survey Assets

Simulations were developed to compare the detection efficiency of various survey 

elements by themselves or in combination. While the detection efficiency of the various 

alternatives should be considered as a primary goal, a second goal is to determine which 

among the viable detection surveys is best for precisely determining orbits. 

To a large extent, PHO orbits are determined by optical angle data (time, right ascension, 

declination) taken over an interval of observation. The most accurate orbits are those that 

are based on the longest interval of observational data. Once a PHO has two or more 

observed returns to perihelion, its orbit is generally well known and in the absence of any 

close planetary flybys, is capable of being accurately extrapolated more than 100 years 

into the future. 

Figure 20 shows the distribution of orbit quality among cataloged objects in the 140- to 

180-meter diameter size range for a number of individual 10-year surveys. Each curve is 

normalized to the catalog size; thus the plot shows no information about the systems’ 

discovery completeness. The term, dP/P, is a measure of in-track orbit accuracy 

calculated as the ratio of orbital period uncertainty (dP) divided by the calculated orbit 

period. An uncertainty in orbital period can be related to the error ellipse of the object 

during an Earth encounter. 

Figure 20. Comparison of Precision Orbit Capabilities of Survey Alternatives 

Better
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2006 Near-Earth Object Survey and Deflection Study 

54 DRAFT Pre-Decisional Material 

Several conclusions can be drawn from Figure 20: 

• The flattening around dP/P=5 x 10-6 is due to the transition from a single observed 

return (i.e., apparition) to multi-apparition orbits. 

• Spaceguard can see very few objects at two apparitions and most have poor 

precision due to very short data arcs. Thus, for the next generation of search 

systems, the current Spaceguard Survey performs poorly at both discovery and for 

precisely determining orbits. 

• The Venus-trailing missions have catalog precisions comparable to those of the 

dedicated LSST system. 

• The two PS4 systems have relatively few high-precision orbits compared with the 

other systems, but the dedicated PS4 system also has a relatively low fraction of 

poorly defined orbits (dP/P > 10-3). 

• The 0.5-meter IR at Sun-Earth L1 system appears similar to LSST for highprecision orbits, but has a greater percentage of low-precision orbits. 

• The 1-meter aperture systems are expected to have the same orbit precision as 

those with smaller apertures due to the assumption of identical sensor arrays. 

For surveys with comparable discovery rates, Figure 20 may be used to distinguish the 

survey with the superior cataloging precision. It cannot be used to distinguish the survey 

with the highest cataloging performance. 

5.10.2. The Role of Precovery in Precision Orbit Determination

“Precovery” is used to describe “pre-discovery” observations of an object that may be 

found in older archived images. Once an object has been detected and an initial orbit has 

been determined, astronomers can project the motion of the object backward in time and 

search for it in archived images. Unless a sequence of several images was taken at that 

time, it is very unlikely that a faint, moving object would have been detected. 

To a great extent, the accuracy of an object’s orbit depends upon the time interval of the 

observations (see Section 5.9). Because precovery has the ability to lengthen the 

observational data interval of a recently discovered object, it often can be used in the 

orbit-determination process to dramatically improve accuracy. Precovery is particularly 

important when a future Earth impact cannot be immediately ruled out for a particular 

object whose orbit is relatively uncertain due to a short-observational interval. 

Figure 21 shows how precovery observations were used to rule out an Earth impact in 

2029 by the asteroid Apophis. The solid line in this figure shows the theoretical behavior 

expected for the 2029 impact. The probability of an impact decreases as more optical 

observations become available for use in the orbital estimates. If the precovery 

observations had not been found, the plotted circles show that orbital solutions generated 

from observations collected after the discovery would have followed the theoretical curve. 

Note that the impact probability would have been calculated to reach 12%. Additional 

observations would then have brought the impact probability to zero. The red-lined plot 

shows the evolution of the actual calculated impact probability for 2029. These 

calculations peaked at about 3% in late December 2004 before precovery observations 

collected 4 months earlier were identified in the data archives and the probability of 

impact was eliminated. 
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2006 Near-Earth Object Survey and Deflection Study 

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Figure 21. Value of Precovery Analysis 

Currently, most precovery observations are identified in the data archives by a group of 

sophisticated amateur astronomers. If future surveys maintain their images in archives, or 

at least archive the time and positional information for each object detected, amateur 

astronomers could continue their work in this area. However, the potential two orders-ofmagnitude increase in the detection rate may prove too rapid for this group unless 

personnel and funds are allocated to this task. 

5.10.3. Precision Orbit Determination Using Radar

Radar observations of NEOs have been taken for several decades. Compared with optical 

data, these measurements are very accurate (8 meters in range, 1 mm/s in range rate). In 

addition, they provide radial data (observer to NEO direction) that complement angular, 

plane-of-sky data. Therefore, the use of radar observations can determine PHO orbits that 

are up to two-to-three orders of magnitude more precise, especially for those objects that 

have short optical observational intervals. 

The value of the radar data is shown in Figure 22. In this figure, precision, given by the 

accuracy of the orbital period (dP), divided by the period (P), is plotted against the total 

interval of observation (arc length). This figure demonstrates the precision for the 

existing catalog of PHO orbits. The red dots show the results for orbits using only optical 

data. The blue stars show the improved orbital precision resulting when orbital solutions 

include both radar and optical data. This figure shows that the addition of radar data can 

make a difference of several orders of magnitude in accuracy over relatively short 

intervals of observation. If optical data are collected for more than about 25 years, radar 

and optical results converge to about the same accuracy. 
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2006 Near-Earth Object Survey and Deflection Study 

56 DRAFT Pre-Decisional Material 

Figure 22. Orbital Precision of PHOs Observed with and without Radar 

Figure 23 shows the simulated distribution of orbit quality for a typical survey with and 

without radar observations. The green curve shows optical-only orbits, e.g., ~75% of 

cataloged orbits having better than 10-5 precision. The blue curve reveals the orbit quality 

of objects that have been observed by radar, showing the substantial improvement for 

low-precision orbits and the more moderate improvement for orbits that already have 

high precision. Assuming an optimistic upper limit for ranging by existing radars of about 

10% of the cataloged objects in this size bin (15 – 20 per month) over a 10-year survey, 

the red curve shows the quality of the combined radar and non-radar catalogs. The 

modest improvement relative to the green curve is because only a small fraction of 

objects are radar ranged and because most objects (~75%) have been optically observed 

at multiple apparitions. 

Radar is able to rapidly generate precise orbits for recently discovered objects with short 

observational intervals, when those objects pass within 0.3 AU of the Earth. This 

information is useful in deciding whether an object is a short-term threat. For example, 

radar data may be important for scenarios like Apophis (see Section 5.10.2) where one 

may need to refine orbits relatively rapidly (less than 5 years) to determine if an object 

has a high probability of becoming a threat. In short, while radar does not provide a 

statistically impressive improvement in precisely determining orbits for the entire catalog, 

radar is important for a relatively few objects of high interest. 

Figure 24 shows an estimate of the percentage of asteroids that pass within radar range 

over a given time period. This corresponds approximately to the probability that any 

particular asteroid will come within range of the radars during that time. 
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2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 57 

Better

Figure 23. The Benefits of Radar Data in the Orbit Determination Process 

5.10.4. Precision Orbit Determination Using a Transponder In-Situ

If an accurate orbit cannot be determined by remote means in time to make mitigation 

decisions, a mission to the PHO may be required. Most in-situ orbit estimates are 

accurate to within 1 km, sufficient to predict a potential impact many centuries into the 

future, assuming no planetary encounters occur. Such an in-situ orbit-determination 

sensor may be coupled with a spacecraft designed to characterize the threat. [9] 

0

10

20

30

40

50

60

70

80

90

100

0 20 40 60 80 100 120

Years Observing

Percentage of PHAs Within Range (%)

0.1 AU Limit (Characterization)

0.3 AU Limit (Precision Orbit)

Figure 24. Percentage of Asteroids within Range of Radars 
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2006 Near-Earth Object Survey and Deflection Study 

58 DRAFT Pre-Decisional Material 

5.11. Detection and Tracking System Element Costs 

Figure 25 shows the costs of the detection and tracking elements through 2020. It 

excludes the costs of cataloguing. Additional cost data are presented in Appendix M. 

$1.5

$1.1

$0.82

$1.0

$0.80

$0.51

$0.14

$0.19

$0.09

 $0.05

$1.8

$0.0 $0.5 $1.0 $1.5 $2.0

Shared PS4

Dedicated PS4

Dedicated PS 8

Dedicated PS 16

Shared LSST

Dedicated LSST

0.5m IR @ L1

1.0m IR @ L1

0.5m IR Venus-Like Orbit

1.0m IR Venus-Like Orbit

1.0m Vis Venus-Like Orbit

2.0m Vis Venus-Like Orbit

$0.01

Figure 25. Survey System Elements Costs through 2020 

5.12. Object Characterization Trade Space 

Figure 26 illustrates the full object characterization trade space. Existing space assets that 

were considered likely will not be available when the Survey is operating; however, they 

were evaluated to understand the types of systems that may have utility. Systems that are 

similar to detection and tracking assets may have different filters, concepts of operations, 

and data processing which will require additional development. Other variations of theses 

systems could be considered depending on characterization requirements; however, these 

are considered representative of the range of systems likely to be the most useful. Five 

alternatives were in the initial trade space, but the study did not consider analyzing them 

because of the reasons listed in Table 13. 

Table 13. Characterization Alternatives Not Analyzed - Rationale 

Concept Name Rationale for Lack of Further Analysis 

SIM PlanetQuest Insufficient sensitivity, cannot characterize 140m NEOs. 

Kepler Insufficient sensitivity, cannot characterize 140m NEOs. 

Will only observe 0.25% of entire sky. 

JWST Unable to track NEOs (fast-moving objects). 

Dedicated JWST Unable to track NEOs (fast-moving objects). 

WISE Points away from Earth, takes 6 months to survey sky. 
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Characterization Characterization

Remote Sensing Remote Sensing In-Situ In-Situ

Ground Based Ground Based Space Based Space Based

Visible Visible Radar Radar

Existing Existing

Proposed Proposed

Gemini Gemini

DCT DCT

Existing Existing

New Asset New Asset

Arecibo Arecibo

Goldstone Goldstone

Bistatic 100m Bistatic 100m

Visible Visible Infrared Infrared

Existing Existing

Proposed Proposed

New Asset New Asset

Hubble Hubble

SIMSIM

Kepler Kepler

2m Vis LEO/L1/L2 w/ 

Filter Wheel

2m Vis LEO/L1/L2 w/ 

Filter Wheel

2m Vis Venus-like w/ 

Filter Wheel

2m Vis Venus-like w/ 

Filter Wheel

Flyby Flyby

Orbit Orbit

Existing Existing

Spitzer Spitzer

Proposed Proposed

New Asset New Asset

JWST JWST

WISE WISE

Multi-channel 0.5m IR 

L1/L2

Multi-channel 0.5m IR 

L1/L2

Multi-channel 0.5m IR 

Venus-like

Multi-channel 0.5m IR 

Venus-like

Flyby + Impactor Flyby + Impactor

Flyby Tour 

(Electric)

Flyby Tour 

(Electric)

Flyby Tour (Chemical)

Flyby Tour (Chemical)

Orbiter Orbiter

Orbiter + Lander Orbiter + Lander

Orbiter + Orbiter + 

Penetrator Penetrator

Multi-channel 1m IR 

L1/L2

Multi-channel 1m IR 

L1/L2

New Asset New Asset

Shared LSST Shared LSST

Dedicated LSST Dedicated LSST

Dedicated 

PS8/PS16

Dedicated 

PS8/PS16

Multi-channel 1m IR 

Venus-like

Multi-channel 1m IR 

Venus-like

Dedicated JWST Dedicated JWST

Infrared Infrared

Existing Existing

IRTF IRTF

New Asset New Asset

Dedicated IRTF Dedicated IRTF

Key

= eliminated from further consideration

Characterization Characterization

Remote Sensing Remote Sensing In-Situ In-Situ

Ground Based Ground Based Space Based Space Based

Visible Visible Radar Radar

Existing Existing

Proposed Proposed

Gemini Gemini

DCT DCT

Existing Existing

New Asset New Asset

Arecibo Arecibo

Goldstone Goldstone

Bistatic 100m Bistatic 100m

Visible Visible Infrared Infrared

Existing Existing

Proposed Proposed

New Asset New Asset

Hubble Hubble

SIMSIM

Kepler Kepler

2m Vis LEO/L1/L2 w/ 

Filter Wheel

2m Vis LEO/L1/L2 w/ 

Filter Wheel

2m Vis Venus-like w/ 

Filter Wheel

2m Vis Venus-like w/ 

Filter Wheel

Flyby Flyby

Orbit Orbit

Existing Existing

Spitzer Spitzer

Proposed Proposed

New Asset New Asset

JWST JWST

WISE WISE

Multi-channel 0.5m IR 

L1/L2

Multi-channel 0.5m IR 

L1/L2

Multi-channel 0.5m IR 

Venus-like

Multi-channel 0.5m IR 

Venus-like

Flyby + Impactor Flyby + Impactor

Flyby Tour 

(Electric)

Flyby Tour 

(Electric)

Flyby Tour (Chemical)

Flyby Tour (Chemical)

Orbiter Orbiter

Orbiter + Lander Orbiter + Lander

Orbiter + Orbiter + 

Penetrator Penetrator

Multi-channel 1m IR 

L1/L2

Multi-channel 1m IR 

L1/L2

New Asset New Asset

Shared LSST Shared LSST

Dedicated LSST Dedicated LSST

Dedicated 

PS8/PS16

Dedicated 

PS8/PS16

Multi-channel 1m IR 

Venus-like

Multi-channel 1m IR 

Venus-like

Dedicated JWST Dedicated JWST

Infrared Infrared

Existing Existing

IRTF IRTF

New Asset New Asset

Dedicated IRTF Dedicated IRTF

Key

= eliminated from further consideration

Figure 26. Complete Characterization Trade Tree 
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2006 Near-Earth Object Survey and Deflection Study 

60 DRAFT Pre-Decisional Material

5.13. Introduction to Characterization 

5.13.1. Object Characterization Parameters Useful for Deflection

Appendix Section J.1.1 describes the object characterization parameters useful for 

designing deflection missions and potential methods of obtaining this data using remote 

and in-situ means. These characteristics are tied to specific deflection alternatives in 

Section 6.9 and Section 6.10. 

5.13.2. The Role of Characterization Using Remote Sensing

The role of characterizing objects that do not pose a specific threat to Earth is to provide 

a predictive understanding of the PHO population as a whole. Remote-sensing methods 

will provide the first information on any newly identified PHO. These data are indirect; 

therefore, key physical properties (mass, size, composition, spin-rate) must be inferred 

from the observations. In contrast, a NEO that has been visited by a spacecraft becomes a 

“standard star.” The data from such a visit validates remote-sensing techniques and 

increases confidence in the models. In the event of “short warning” cases where schedule 

does not allow a dedicated visit to the PHO, this information would provide the basis for 

planning mitigation. These short warning cases become less probable as the Survey 

progresses and are improbable when the population is fully catalogued. 

The methods to remotely sense PHOs are the same as those used to study the much larger 

population of main-belt asteroids, the principal source of NEOs. Thus, known techniques 

developed over many years are available to carry out these measurements. More recently, 

adaptive optics and radar have been used to detect asteroidal binaries. This technique can 

provide unique information about asteroidal masses, just as it has provided unique data 

on the stellar masses of binary stars. Masses of binary systems may be determined more 

accurately by using Kepler’s law than by observational data alone. 

Some of these techniques can be used at many facilities around the world using 

telescopes of sufficient sensitivity. Other capabilities, such as radar and polarimetry, are 

available at only a few locations. More important, a cadre of trained observational 

astronomers needs to be available should characterization work be deemed necessary. 

Applying several of these remote techniques to a single NEO forms a hierarchy in the 

accuracy of the inferred values of the key characterization parameters. As shown in 

Figure 27, optical-intensity measurements of a newly discovered NEO enable an estimate 

of its mass to within a factor of about 50. If remotely sensed broadband colors are added, 

the mass estimate may be improved by a factor of eight. Adding spectroscopic 

observations to the mass estimates improve accuracy by a factor of five or six. 

Polarimetric observations can improve the accuracy by about a factor of three, and radar 

can improve the mass estimate by a factor of 2. Finally, very accurate mass 

measurements (to 1%) can be achieved only by visiting the object. 
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2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 61 

Figure 27. Distribution of Types of NEO Characterization Performed

A key issue with ground-based and non-dedicated space-based activities is access and 

contention for resources. Astronomical facilities and their focal-plane instruments are 

scheduled far in advance. On any given night, observations have been scheduled well in 

advance and some are tied to a particular date. Most PHOs will be discovered near 

opposition and are best observed while still near opposition. Therefore, capabilities for 

“on demand” access at major facilities must be in place to effectively use non-dedicated, 

ground-based remote-sensing techniques. In addition, the rate of discoveries may 

overwhelm current infrastructure without additional resources. The window of when 

PHOs may be observed is short; so dedicated facilities may be needed to gather most 

detailed characterization information for a significant fraction of new discoveries. 

5.13.3. The Role of In-Situ Characterization Missions

Remote characterization provides data about the basic characteristics of objects in the 

population; however, the resulting information requires a measure of inference to 

determine the actual characteristics of a specific body. While the accuracy of these 

inferences improves with a wider range of remote characterization methods, only in-situ 

encounters can provide the definitive observations needed to calibrate remote 

observations. A much larger number of NEOs must be visited to collect statistically 

significant characterization information about the object population (the precise number 

depends on the variation allowed); therefore, it is likely that in-situ visits will be limited 

to model verification as opposed to model validation. 

In-situ characterization of actual threats also may be critical for enabling many of the 

deflection options described in Chapter 6. It may eliminate the probability of impact for 

some threats by providing high-precision orbit determination. This will improve the 

reliability and effectiveness of any deflection campaign. 
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62 DRAFT Pre-Decisional Material

5.14. Object Characterization Strategy 

The characterization required to provide warning of threats and inform mitigation 

depends on several factors. One factor is the possible relationship between detection, 

tracking, and characterization elements. For example if a ground-based survey system 

(such as LSST) is built, its concept of operations will likely limit its ability to do followup characterization unless search goals are compromised. If a space-based survey system 

(such as an infrared system in a Venus-like heliocentric orbit) is built, options for 

characterization include upgrading the system with filters, building a second dedicated 

space-based characterization system based on the design of the first, or building a 

dedicated ground-based system. If detection and tracking assets are available after search 

goals are reached, some survey systems may be transitioned to contribute to 

characterization after detection operations are less operationally intense. 

Another factor is the relative importance of committing resources to remote 

characterization to prepare for a potential “short warning” scenario. If short warning 

scenarios (less than 15 years warning) must be addressed, a statistically valid plan for 

sampling the diversity of the PHO population must be accomplished before the detection 

and tracking system can retire the threats. To enable a short warning characterization 

capability, models of the PHO population must be created and will likely need to be 

verified by in-situ visits. Model verification missions may include visits to each primary 

asteroid type or to specific threats as they are identified. 

The most important factor in developing the characterization strategy is to tie its 

requirements to a specific mitigation strategy or selected deflection options. For example, 

if only standoff nuclear explosive alternatives are considered, little characterization 

beyond orbit and approximate mass is required. If a space tug becomes the highest 

priority, more information would be required and a higher investment in characterization 

is warranted. Due to the diversity of information required to support all deflection 

alternatives, a very robust program will be required to enable all deflection systems. 

Without the selection of specific search and mitigation strategies, a specific choice of a 

characterization program is premature. Therefore, this study analyzed a range of 

characterization capabilities listed in Table 14. For these options, Option 7 is similar to 

Option 6 in that it combines dedicated ground-based and space-based remote 

characterization with a number of in-situ orbiters. In Option 6, the chosen number of 

orbiters is eight, possibly one for each of the primary asteroid classes needed to calibrate 

remote characterization models. At least one representative of each asteroid type (which 

may number greater than eight) must be visited to contribute materially to the model 

verification purpose of in-situ visits. In Option 7, the strategy is to send characterization 

missions to credible threats. 

For example, a strategy may be to characterize one of the highest risk PHOs discovered 

during 5-year intervals. This approach would involve visits to representative threats to 

validate models. It also would provide in-situ orbital determination to verify or eliminate 

specific hazards at a routine and sustainable mission rate. The orbits of the targets for 

Option 7 are likely more difficult to reach than those that would likely be chosen for 

Option 6. The capabilities of these options are evaluated against the deflection 

alternatives in Section 6.10. 
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Table 14. Characterization Capability Options 

Option Descriptions (O1 = Option 1) 

Option 1 Use Existing Assets + Detection and Tracking Systems 

Option 2 O1 + Dedicated Ground Systems 

Option 3 O1 + Dedicated Space-Based Remote Sensing (L1/L2) 

Option 4 O1 + Dedicated Space-Based Remote Sensing (Venus-Like Orbit) 

Option 5 O1+ O2+ O3 + 2 Flyby Missions to 8 Objects 

Option 6 O1 + O2 + O3 + 8 Orbiter Missions 

Option 7 O1 + O2 + O3 + Orbiters at a Fixed Threshold Probability of Impact 

5.15. Characterization Option Architectures 

Table 15 provides possible timelines for candidate architectures representing each of the 

characterization capability options. Additional discussion of their costs is presented in 

Appendix Section M.4. 
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64 DRAFT Pre-Decisional Material

Table 15. Possible Characterization Option Architectures and Timelines Architecture Elements Option 1: Use Existing Assets + Detection and Tracking Systems 2007 2026 Existing Assets IOC Year (Life Cycle Cost FY06$B) 2007 (*) Option 2: Option 1 + Dedicated Ground Systems 2007 2026 Option 3: Option 1 + Dedicated Remote Sensing @ L1 2007 2023 Option 4: Option 1 + Dedicated Remote Sensing in Venus-like Orbit 2007 2024 Option 5: O1+ O2+ O3 + 2 Flyby Missions 2007 2023 2007 2031 Option 7: Option 6 at Fixed Impact Probability Threshold (every 5 yrs) 2007 2053 Option 6: O1+ O2+ O3 + 8 Orbiter/Lander Missions Start End Existing Radars Shared DCT Shared LSST 2007 (*) 2010 (*) 2014 (0.1) Existing Assets 2007 (*) Existing Radars Shared DCT Shared LSST 2007 (*) 2010 (*) 2014 (0.1) Dedicated IRTF 2014 (0.1) Existing Assets 2007 (*) Existing Radars Shared DCT Shared LSST 2007 (*) 2010 (*) 2014 (0.1) Dedicated IRTF 2014 (0.1) 0.5m IR L1/L2 2013 (0.9) IOC Year (Life Cycle Cost FY06$B) IOC Year (Life Cycle Cost FY06$B) IOC Year (Life Cycle Cost FY06$B) IOC Year (Life Cycle Cost FY06$B) IOC Year (Life Cycle Cost FY06$B) IOC Year (Life Cycle Cost FY06$B) Existing Assets 2007 (*) Existing Radars Shared DCT Shared LSST 2007 (*) 2010 (*) 2014 (0.1) Dedicated IRTF 2014 (0.1) 0.5m IR Venus 2013 (1.0) Existing Assets 2007 (*) Existing Radars Shared DCT Shared LSST 2007 (*) 2010 (*) 2014 (0.1) Dedicated IRTF 2014 (0.1) 0.5m IR L1/L2 2013 (0.9) Flyby Tour Electric 2013 (0.5) 2015 (0.5) Flyby Tour Electric Existing Assets 2007 (*) Existing Radars Shared DCT Shared LSST 2007 (*) 2010 (*) 2014 (0.1) Dedicated IRTF 2014 (0.1) 0.5m IR L1/L2 2013 (0.9) NEO Orbiter 2013 (0.7) 2015 (0.7) NEO Orbiter Mission Every 2yrs ... (0.7) Existing Assets 2007 (*) Existing Radars Shared DCT Shared LSST 2007 (*) 2010 (*) 2014 (0.1) Dedicated IRTF 2014 (0.1) 0.5m IR L1/L2 2013 (0.9) Lander Every 5yrs 2013 (1.0) Mission Every 5yrs ... (1.0)

* Costs less than $50M

LCC

$0.1B

FY06

$0.2B

FY06

$1.1B

FY06

$1.2B

FY06

$2.0B

FY06

$6.7B

FY06

$8.2

FY06
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5.16. Characterization Development Schedules 

Figure 28 displays the nominal development time for the characterization system 

alternatives. These schedules are not directly tied to any need date. 

Figure 28. Characterization Systems Development Time 

5.17. Characterization Elements Cost Estimates 

Figure 29 displays the LCC through 2020 in FY06$B for the characterization concepts. 

$0.3

$0.1

$0.3

$0.6

$1.7

$1.8

$1.8

$0.8

$0.7

$1.0

$1.0

$0.2

$0.3

$0.2

$0.4

$0.3

$0.3

$0.4

$0.5

$0.5

$0.8

$0.7

$1.0

$0.9

$0.1

$0.2

$0.4

$0.1

$0.1

$0.0 $0.5 $1.0 $1.5 $2.0

Shared DCT

Shared LSST

Dedicated PS4

Dedicated PS8

Dedicated PS16

Dedicated LSST

Dedicated IRTF

Bistatic Radar

2.0m Visible + Filters LEO

2.0m Visible + Filters L1/L2

2.0m Visible Venus-Like Orbit

0.5m Multi-Channel IR L1/L2

0.5m MC IR Venus-Like Orbit

1.0m Multi-Channel L1/L2

1.0m MC IR Venus-Like Orbit

Flyby Electric Propulsion

Flyby Chemical Propulsion

Flyby (Chemical) w/Impactor

Orbiter

Orbiter with Lander

Orbiter with Penetrator

Development, Launch, and Operations Costs ($B 2006)

Dev / Ops Cost to 90% Search

Date, or 2030 ($B2006)

Incremental Cost to

Characterization ($B2006)

Notes for Ground or Space Based Telescopes:

1. Characterization missions through CY2030

2. Mods to instruments funded by characterization

3. Sharing assets with detection will delay 90% goal

Filters

If elements very similar to 

those of the search system 

are acquired, costs for 

dedicated characterization 

“copies” will be less

Figure 29. Characterization Alternatives Life-cycle cost Results 
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2006 Near-Earth Object Survey and Deflection Study 

66 DRAFT Pre-Decisional Material

5.18. Performance and Cost 

Table 16 displays a summary of search alternatives and their life-cycle costs through 

2020. Figure 30 and Figure 31 include combined performance and life-cycle cost metrics. 

Figure 30 displays percentage completeness at the end of 2020 vs. the life-cycle cost of 

various combinations, all of which include Shared PS4 and Shared LSST as the baseline. 

Figure 31 displays the year of achieving 90% completeness vs. the life-cycle cost of the 

same systems. In each figure, related data are grouped and labeled by similar attributes. 

Table 16. Summary of Detection, Tracking, and Cataloguing Alternatives 

 

through 

2020 

Cost thru 

2020 

(FY06) 

Year 

90% 

Cost to 

90% 

(FY06) 

Continue Spaceguard (in all) 14% < $0.2B >>2030 - 

Shared PS4 and shared LSST* 83% $0.31B 2026 $0.52B 

Dedicated PS8 + Baseline 85% $0.41B 2024 $0.56B 

Dedicated LSST 85% $0.66B 2024 $0.87B 

Dedicated LSST + Baseline 90% $0.82B 2020 $0.82B 

0.5m IR @ L1 + Baseline 91% $1.1B 2020 $1.1B 

1.0m IR @ L1 + Baseline 91% $1.3B 2019 $1.3B 

0.5m IR @ Venus + Baseline 97% $1.1B 2018 $1.0B 

1.0m IR @ Venus + Baseline 97% $1.4B 2017 $1.3B 

1.0m VIS @ LEO/L1 + Baseline 93% $1.8B 2017 $1.7B 

2.0m VIS @ LEO/L1 + Baseline 95% $2.1B 2019 $2.0B 

 * Baseline = Existing + Shared PS4 + Shared LSST 

60%

65%

70%

75%

80%

85%

90%

95%

100%

$0.0 $0.5 $1.0 $1.5 $2.0 $2.5

Lifecycle Cost ($B 2006)

% PHOs Surveyed by End of CY 2020

Goal

Baseline = Shared PS4 + Shared LSST

Baseline + 

New Dedicated 

Ground Based

Baseline + Dedicated LSST

Baseline + Shared 

Ground Based

Baseline + Space Visible

0.5 m

1 m

2 m

Drop to Baseline if space-based mission fails (~10%)

Baseline + Space IR

Figure 30. Survey Combinations % Completeness vs. Cost Through 2020 
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2015

2020

2025

2030

2035

$0.0 $0.5 $1.0 $1.5 $2.0 $2.5

Lifecycle Cost ($B 2006)

Year of 90% PHO Completeness

Baseline = Shared PS4 + Shared LSST

Baseline + 

New Dedicated 

Ground Based

Baseline + Dedicated LSST

Baseline + Shared 

Ground Based

1 m

0.5 m

2 m

Baseline +

Space Visible

Goal

Baseline +

Space IR 

Existing Systems Never Reach 90% Goal

Figure 31. Survey Combinations Date of 90% vs. Life-cycle cost thru 90% 

Table 17 summarizes the description and total costs of the seven characterization 

capability options. 

Table 17. Summary of Characterization Capability Options 

 Description

Cost (FY06) 

Period

Option 1 Use existing assets plus detection and tracking 

system elements. No dedicated characterization.

$0.1B 

2007-2026 

Option 2 Develop dedicated ground system(s) to gather and 

analyze data. No in-situ missions.

$0.2B 

2007-2026 

Option 3 Develop dedicated space system(s) to gather and 

analyze data at Sun-Earth L1. No in-situ missions.

$1.1B 

2007-2023 

Option 4 Develop dedicated space system(s) to gather and 

analyze data near Venus. No in-situ missions.

$1.2B 

2007-2023 

Option 5 Add to Options 1-3 eight (8) visits to representative 

NEOs using fly-bys to calibrate models. 

$2.0B 

2007-2024 

Option 6 Add to Options 1-3 eight (8) visits to representative 

NEOs using rendezvous missions to test models. 

$6.7B 

2007-2031 

Option 7 Perform Option 6 at a fixed mission rate to 8 

potential threats at 5 years intervals 

$8.2B 

2007-2053 
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5.19. Program Elements 

Part of the congressional direction was for NASA to develop a program to carry out the 

recommended search and characterization. The following elements are likely to be 

required by any such program. 

Survey and Characterization Systems

• Detection and tracking system (ground and/or space based) 

• Orbit improvement/follow-up systems (ground and space) 

• Object characterization options (ground, space, and/or in-situ) 

Infrastructure for Space-Based Elements

• Communications system 

o No new communications systems are expected to be required, but detection 

and tracking bandwidth needs to be considered in the NASA communications 

architecture 

• Command and control system 

Data Analysis Infrastructure

• Data cataloguing and distribution system 

• Preliminary orbital determination and ephemeris generation 

• “Precovery” - analysis of archived data of previous observations 

• Data collection, validation, archiving 

o Requires upgrade to current Minor Planets Center or new infrastructure 

• Precision orbital determination and prediction 

• Earth-impact prediction and warning (alerts) 

o Requires upgrade to current JPL and NEODyS capabilities or new systems 

5.20. Other Factors 

5.20.1. Opportunity Science from the Survey and Characterization 

NEOs are primitive bodies, primarily asteroids that represent almost the full range of 

material contained in the main asteroid belt. The population also contains the nuclei of 

extinct comets, which likely still reflect the composition of all but the most volatile 

species, and in particular still contain a significant inventory of organic substances. 

The Decadal Survey of the Solar System summarizes the key science issues with respect 

to primitive bodies as follows: 

• Where in the solar system are the primitive bodies found, and what range of sizes, 

compositions, and other physical characteristics do they represent? 

• What processes led to the formation of these objects? 

• Since their formation, what processes have altered the primitive bodies? 

• How did primitive bodies make planets? 

• How have they affected the planets since the epoch of formation? 

Characterization will certainly provide new information on the sizes, compositions, and 

other physical characteristics of asteroids and comet nuclei. Information on the material 

of these objects also will provide data to understand alteration processes. 
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Most searches will substantially increase the identification of Kuiper Belt Objects 

(KBOs). For example, if 10% of the observing time on the Dedicated LSST is spent in 

KBO search mode, roughly 100,000 faint KBOs will be discovered. An expanded KBO 

database will allow the study of dynamical distributions, further resonances, the existence 

of a KBO demarcation beyond 50 AU, high-eccentricity/high-inclination orbits, size 

distributions, frequency of binary objects and collision rates, chemical compositions and 

the relationship of objects to dust disks around other stars. The survey also will provide a 

rich database of destinations for future space missions. 

Detection surveys such as Pan-STARRS and LSST provide unique solar-system science 

because they are designed to detect and perform follow-up studies of moving objects. 

Centaurs, Jupiter Family Comets, and certain extinct comets may be related through a 

common origin in the Kuiper Belt. Dedicated assets will assure that appropriate followup is carried out over the yearly timeframes that are required to produce orbits for the 

slower-moving objects found in the outer solar system. Thus, a collateral result of the 

search program could be both the delineation of the structure of the Kuiper Belt and the 

discovery of many new minor planets. 

It also is important to understand what a threat assessment and characterization effort will 

not accomplish. Characterization to inform deflection missions does not require sample 

return from either an asteroid or comet. Asteroid and comet sample-return missions are 

high priorities in the Decadal Survey, but are not included in the trade space of this study. 

As noted, the Survey program will identify many likely candidates for scientific visits 

and sample return. Remote characterization will allow the most interesting objects to be 

selected for scientific visits and will allow the instruments and experiments of these 

missions to be tailored in ways that otherwise would be impossible. NEOs are generally 

among the easiest asteroids to visit, and the design of a spacecraft to work in the 

relatively benign environment near 1 AU offers less cost and risk than a mission to the 

main belt (crudely the difference between the Near Shoemaker mission and the Dawn 

mission). A sample return mission to a NEO characterized for a deflection mission will 

carry substantially lower risk than a mission to an object about which less is known. 

5.20.2. Potential Benefits to Exploration

5.20.2.1.Near-Earth Object Resources 

This study has identified a loose connection between the goals of the Vision for Space 

Exploration and a program to survey the population of NEOs. There may come a time 

when Earth’s resources are insufficient or too costly to support the planet’s growing 

population. Exploring resources that exist on the Moon, other planets, or NEOs may 

allow further human expansion. 

The survey assets examined by this study will take 5-10 years to provide an extensive 

map of the orbits and sizes of NEOs to 140 meters in diameter, as well as information on 

thousands of smaller objects. If infrared survey assets are built, these assets could be 

turned to the job of characterizing the composition of these objects. In addition, this study 

has also identified several funded efforts to survey and characterize the NEO population, 

which likely will come about with minimal NASA contribution. 
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If asteroid or comet resources prove enabling, having a map of the location and 

distribution of these assets may prove valuable. An analogy might be the mapping of 

oases to facilitate transportation across the desert. Assuming that humankind will not be 

ready to exploit such a map of asteroid resources for at least 30-40 years, it is very likely 

that this map will be created as a direct product of otherwise funded scientific surveys. If 

these envisioned efforts do not produce the required information, it is expected that a 

limited expenditure of time and resources (less than 10 years and $1B) will be needed to 

produce a map of asteroid and comet resources. 

5.20.2.2.Human Visits to Asteroids 

It is possible that the systems used to return humans to the Moon could be used to visit a 

NEO. While NASA has no published plans or budget to pursue such a mission, the NEO 

survey and characterization program could be used to help select the destination for such 

a mission. A visit to a NEO could be used to demonstrate technologies for lunar missions, 

or as an interim goal between lunar and Mars missions. 

5.20.3. Potential Synergies with Department of Defense

Potential benefits to DoD and the potential of using DoD assets to augment the NEO 

survey are discussed in Appendix O. 

5.21. Findings 

• Combining optical ground-based systems currently under development with a 

ground-based asset dedicated to the survey would allow NASA to reach the 

congressional goal of 90% by the end of 2020. Life-cycle costs for the complete 

architecture, including data management and data analysis, are estimated to be 

$820M through the end of 2020. 

• Space-based infrared systems, combined with shared ground-based assets, could 

reduce the overall time to reach the 90% goal by up to 3 years, with life-cycle 

costs of $1.0-$1.3B through 90% completion. Space systems have additional risks 

and benefits over ground-based alternatives. These benefits include improved 

estimation of PHO size (via IR), completeness to 90% for smaller object 

diameters, and an improved understanding of the population of Atens and IEOs. 

• The requirement to detect 140-meter PHOs will be near the limiting magnitude of 

most systems considered and will require these systems to perform their own 

follow-up observations, unless dedicated assets are acquired. Depending on the 

follow-up characterization requirements, this may delay survey completeness. 

• Atens and IEOs may be under-represented by current population estimates and 

can best be viewed from assets in a Venus-like orbit. No other vantage point 

considered in this study offers the opportunity to observe these objects as fully as 

an orbit well inside that of Earth’s. 

• The number of objects to be detected does not principally affect the date of 

completion, however, assumptions about the orbital distribution of objects may. 

• Continued operation of the current Spaceguard Survey after the more advanced 

systems are running will not significantly improve the total discovery potential. 
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• Collection and storage of as much raw data as possible is important for precovery 

and follow up. 

• Existing data systems, infrastructure, and personnel will be challenged by the 

increased rate of data collection (up to 100 times more than today) and by the 

increase in the number of objects. A significantly more robust infrastructure will 

be necessary to meet requirements. All architectures presented include a robust 

data program ($150M through 2020), but lower cost options are available 

depending on further definition of requirements. 

• Unpaid astronomers accomplish a substantial percentage of precovery analysis. 

The ability of unpaid astronomers to maintain this role at the higher absolute 

magnitudes and at an increased rate of discovery is a significant unknown and 

may require incentives or substantial funding. 

• Many follow-up systems that may be used to characterize PHOs are highly 

subscribed. They may be able to carry out “on demand” observations, which are 

necessary for the time that many of these objects are visible. Either “on demand” 

time needs to be allocated or assets need to be built or partially dedicated to 

follow-up observations. 

• Once the survey is complete, almost every object (to about 99%) will have 

multiple apparitions of data and any random PHO will, as a result, have centuries 

of predictable encounters, providing centuries of warning if it becomes a hazard. 

• Radar systems can be used to rapidly refine tracking and object size for a few 

objects of potentially high interest. 

- Both the Arecibo and Goldstone radars are heavily oversubscribed. Only a 

small percentage of their time is available for asteroid radar. 

- Near-Earth asteroid observations have been cancelled due to insufficient and 

difficult-to-maintain support personnel. 

- Radars may require immediate and stable funding to be available for followup PHO observations. This is particularly true if radar is necessary for the 

relatively few high-interest objects for which only radar can provide highly 

precise orbit and characterization data. 

• The orbit and approximate mass of PHOs are necessary to determine if they are a 

threat. Additional requirements for characterization of hazardous objects may be 

derived from a specific mitigation strategy and the search system selected. 

• Systems operated by the DoD were evaluated by this study. They are not expected 

to contribute to the congressional goals due to differences in mission requirements, 

concepts of operations, and procedures for data storage. 

• While several countries have capable programs to study NEOs, none of these 

efforts has materially influenced the findings of the study team. 
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6 Analysis of Deflection Alternatives 

6.1. Introduction and Background 

If the capabilities of survey systems progress as expected, a PHO that poses a credible 

threat to Earth will be discovered some day. Sensors will indicate that the object could 

possibly strike the Earth on a specific date in the future, likely many decades or centuries 

in the future. Systems will continue to track the object as it proceeds toward Earth. 

The probability of impact likely will start at a low level, perhaps on the order of 0.01% or 

less, meaning there will be one actual impact for 10,000 similar warnings. As tracking 

data accumulates, the recalculated probability of impact may rise, and likely will remain 

under 1% for some time. The probability calculation is based on both the PHO’s 

estimated orbit and the orbit’s uncertainty at the time of the predicted encounter. If the 

encounter is many years or decades into the future, the orbit uncertainty may be large 

because it grows over the prediction period. Additional uncertainty may be introduced if 

the PHO is perturbed by solar pressure, the Yarkovsky Effect, or if it encounters other 

objects before impact. Orbit uncertainty is discussed further in Section 5.9. As the object 

continues to be tracked and the time to impact shortens, the uncertainty will decrease and 

the probability of impact will change, with most threats being eliminated decades before 

the possible impact. 

Based on this information, decisions will need to be made about how to respond to the 

threat and whether to launch a mission to reduce the impact probability to some 

acceptable level, a risk level set at 1 in 1 million in this study. An effort to mitigate the 

threat may require any or all of the following: time to understand the threatening object; 

time to debate and resolve political and policy issues; time to approve funding; time to 

design and build spacecraft; time to obtain the appropriate launch vehicle(s) and launch 

site(s); time to assess the results of the deflection attempt; and time to make additional 

attempts, if necessary. 

This study differentiates between mitigation in general and the congressional direction to 

study methods of diverting (or deflecting) an object. Mitigation involves all efforts to 

reduce the severity of a potential impact with Earth, while deflection specifically 

addresses diverting the object. Therefore, these terms are not used interchangeably and 

deflection is generally considered to be one of several possible mitigation options, 

A wide range of deflection options were considered and evaluated. Perhaps the simplest 

of these is the kinetic impactor, where a spacecraft is collided with the PHO to change its 

orbit so that the object misses Earth. Detonating a conventional or nuclear explosive as 

part of the deflection effort can increase a kinetic impactor’s effectiveness. These are 

impulsive techniques that act nearly instantaneously to change the velocity of the PHO. 

In addition, a number of “slow push” techniques such as a space tug also were considered. 
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6.2. Definitions 

6.2.1. Deflection Campaigns

A deflection campaign is defined as the combined series of characterization, deflection, 

and effectiveness assessments required to mitigate the threat. This may include multiple 

in-situ characterization missions, multiple deflection approaches, multiple launch systems, 

and multiple attempts per deflection approach. See Section 6.12 for additional discussion 

of this topic. 

6.2.2. Launch Energy and C3

A parameter that matches launch capability with a certain payload at a certain time (flight 

time) to intercept an asteroid is C3. It is equal to twice the specific (per unit mass) orbital 

energy, with units of km2

/s2. The implications of this parameter on the number of PHOs 

that can be intercepted are discussed further in Appendix L. 

6.2.3. Momentum Exchange Efficiency (beta, β)

Beta factor (β) is a measure of how efficiently momentum transfers after a collision. If 

the impactor passes through the object or causes material to separate from the object in 

the direction the impactor was traveling, the beta factor would be less than 1. If a purely 

plastic collision occurs, that is no ejecta is produced and the impactor is absorbed, the 

beta factor is equal to one. If the impact ejects material in the direction from which the 

impacting object came, the momentum of the ejecta adds to that imparted by the 

impactor. For moderate ejecta production the beta factor increases to around 10. Beta 

factor is assumed to be ~3 for scenarios in this study, but may vary considerably with 

specific asteroid type and composition. [34] 

6.2.4. Specific Impulse (Isp)

The specific impulse of a propulsion system is defined as the thrust produced per unit 

weight flow of propellant. It is a measure of propulsion system efficiency. Specific 

impulse is a useful value to compare the efficiency of rocket engines and is analogous to 

“miles per gallon” for cars. Specific impulse for space systems varies from around 200 

seconds for simple monopropellant systems, to more than 450 seconds for the most 

energetic hydrogen-oxygen systems and several thousand seconds for low-thrust nuclearion propulsion. The chemical engines used in this study assume a specific impulse of 325 

seconds (typical for storable propellants) and 8,700 seconds for the conceptual NEXIS 

ion-propulsion system. 

6.3. Derived Requirements 

Several requirements were explicit in the definition of this study. For example, Congress 

directed NASA to study alternatives capable of diverting an object on a collision course 

with Earth. The derived deflection requirements drawn from the congressional language 

are as follows. 
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6.3.1. Derived Deflection Distance Requirement

To assess the effectiveness of the deflection alternatives, the study needs to define 

deflection. In the context of this study, deflection means imparting sufficient ΔV to a 

PHO so that it will not strike Earth. Since there are uncertainties about an object’s orbit, 

the amount of ΔV imparted must have sufficient margin to ensure an acceptably low 

probability of impact. Different scenarios may require different deflection distances to 

achieve this margin. 

Based on the data in Figure 2, there is about a 1 in 1 million chance that an object 1 km or 

larger will impact the Earth in any year. This is a well-understood level of “background” 

risk for a well understood asteroid population. Therefore, this study chose this level to 

indicate a successful deflection attempt. For each scenario discussed in this study, the 

hypothetically threatening asteroid or comet was deflected a distance to reduce its 

probability of impact (at the predicted impact date) to 1 in 1 million. 

6.3.2. Derived Reliability Requirement

Deflection requirements will have an effect on derived reliability requirements as any 

unsuccessful mission will reduce the effectiveness of the overall deflection campaign. For 

example, if there is a 100% probability that a particular asteroid will strike Earth in 50 

years and the requirement is to reduce the probability of impact to 1 in 1 million, a 

deflection campaign must have less than a 1 in 1 million probability of failure. This may 

be compared with the historical failure rate of more than 10% of interplanetary missions. 

If warning time permits, deflection missions may be launched incrementally, but random 

and common-cause failures will continue to play a significant factor in campaign 

reliability. The impact of these stringent reliability requirements for mission success is 

developed further in Section 6.12. 

6.3.3. Derived Characterization Requirements

Congressional direction states that the basis for characterization missions is twofold 

• To provide “warning.” Therefore, this relates primarily to precisely determining 

orbits and estimating gross mass/size. 

• To inform “mitigation.” From this statement, other requirements can be derived. 

Some deflection alternatives require specific information about the hazardous object, 

including its shape, rotation rate/axis, multiple primary masses, and composition, while 

others require less. Some of this information may be gathered remotely, while other data 

require in-situ visits to the target. 

Deflection concepts were not developed to a sufficient degree to produce specific, 

numerical requirements for characterization. In addition, before a deflection approach is 

chosen, it may be premature to invest in a broad range of characterization capabilities that 

may or may not support the selected approach. This is particularly true because the 

average impact interval of 5000 years means that any preparations for mitigation likely 

will be obviated (perhaps hundreds of times) by the passage of time. For this study, the 

team developed a range of characterization capabilities to see if they meet or exceed the 

notional requirements. If in-situ characterization missions are required for an alternative, 

the cost of those missions is included in the total cost. 
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6.3.4. Post-Action Effectiveness Assessment Requirements

If a deflection attempt does not end in complete failure, the PHO’s orbit will be perturbed 

by some amount. After this attempt, the previous long-arc of observations of the object is 

no longer valid because the PHO will be in a different orbit. Without a radar opportunity 

or an in-situ aid to determine orbit, there may not be enough time using remote means to 

determine if the deflection attempt was sufficiently successful or if additional attempts 

will be required. 

NASA has derived a requirement that a post-action effectiveness assessment be 

performed to understand if further action is required. In most cases, this will take the 

form of a transponder mission (launched before, with, or after the deflection attempt and 

possibly part of a characterization mission). This requirement does not materially affect 

the outcome of this analysis, but does increase by a marginal amount the cost and 

complexity of all the deflection alternatives. 

6.4. Deflecting a Potentially Hazardous Object 

Deflecting a PHO requires that the trajectory of the oncoming object be modified so that 

it will, with an acceptable probability, miss the Earth. Figure 32 illustrates a case where 

an asteroid has a 1 in 100 probability of striking the Earth some years into the future. The 

objective of a deflection mission might be to apply enough velocity change, ΔV, to the 

PHO to reduce the probability of impact to 1 in 1 million or less. 

Figure 32. PHO Approach Uncertainties and Deflection Goal 

To reduce the probability of impact, the velocity of the PHO must be modified, with the 

amount and direction of the modification dependent on the specific scenario. As Figure 

33 shows, one likely approach would be to change the timing of the PHO’s arrival at the 

intercept point, effectively slowing down or speeding up the object so that it arrives at the 

impact point at a slightly different time and misses Earth. 

Driving a spacecraft into the PHO can change the PHO’s velocity, as can activating an 

explosive device on, in, or near the object, or applying a relatively small force to the body 

over a longer period. The latter are referred to as “slow push” methods. 
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Figure 33. Illustration of PHO Deflection 

Figure 34 shows how the ΔV required to deflect a hypothetical PHO changes as the 

object approaches Earth [23], and how the calculated probability of impact changes 

depending on how much ΔV is applied. The ΔV required to deflect the object also 

increases as the object gets closer. The cyclic variation is related to the period of the 

object’s orbit and reflects the point in the orbit where the ΔV is applied. This case is 

further developed in Section 6.13.4 using the example of the 200-meter asteroid, Athos. 

0

10

20

30

40

0 500 1000 1500 2000 2500 3000 3500 4000

Days to Impact

ΔV Required to Deflect (cm/sec)

Reduce risk to 1:1,000,000

Reduce risk to 1:1,000

Deflect to miss by 1 Earth radius

Figure 34. ΔV Required to Deflect a Hypothetical Asteroid 

6.5. Deflection Mission Timelines 

Table 18 shows a top-level timeline for a deflection mission. As the PHO progresses 

from detection to the predicted impact point, its orbit is refined. At this point, it is 

determined to pose a threat to Earth. If sufficient time exists, a characterization mission is 

launched to gain more knowledge about the object’s mass and physical properties and to 

refine its orbit. This could take place while development begins on a deflection system, 

but likely will not yield results for a few years due to transit time. 
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Once deflections mission are designed and funded, hardware is fabricated and launched. 

Depending on the alternative selected, the action time at the PHO could be instantaneous 

or could extend over months to years. An important task will be to assess the 

effectiveness of the action in lowering the probability of Earth impact. If the assessment 

shows that the effort did not fully accomplish the goal, a second series of missions would 

be initiated. 

Each step along this timeline will require decision making with many types of ambiguous, 

non-intuitive, and sometimes contradictory information. For example, funding may need 

to be approved well before it is certain that the PHO will actually strike Earth. It is likely 

that the greatest element of schedule uncertainty will be due to social rather than 

technical delays. [24] 

Table 18. Potentially Hazardous Object Mission Timeline 

Event Duration 

PHO detected, orbit refined Months to Years 

Remote characterization performed Days to Months 

In-situ characterization designed, launched 2-3 Years 

In-situ characterization performed Months to 2 Years 

Threat threshold exceeded 

Deflection action initiated Indeterminate 

Mission design Months to 1 Year 

Funding Approval Weeks to Months 

Hardware Fabrication and Test 1-3 Years 

Approval of Launch(es) Weeks to Months 

Deflection Launch and Transit Months to years 

Action Time at PHO Instant if Impulsive 

5-10 Years for Slow Push 

Assessment Instant (with transponder) 

Backup Action Initiated Indeterminate (see above) 

Predicted Impact 

 = Necessary event. = Optional event. 

6.6. Figures of Merit 

Figure 35 shows the figures of merit used to evaluate the deflection alternatives. The 

performance metrics focus on the ability of an alternative to deflect an asteroid through 

the transfer of momentum to the object. Performance metrics include the momentum 

imparted normalized to the launch mass, the time to deploy and perform the deflection, 

and the percentage of the PHO population against which the alternative is effective. 
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Performance

— Momentum 

imparted per unit 

mass launched

— Time required to 

deploy and 

perform asteroid 

deflection

— % of PHOs the 

approach is 

effective against

Cost

— ROM cost of 

development

— ROM cost per 

deflection 

campaign

Other Factors

— Benefits to 

Exploration*

— Science data 

provided*

— Benefits to 

DoD*

Effectiveness Affordability

— Development Risk

as measured by 

technical maturity

— Operations Risk

in deploying 

deflection method

Risk

Risk Rated Performance Funding Profile

ROM = Rough Order *Secondary FOM

of Magnitude

Figure 35. Deflection Figures of Merit 

6.7. Deflection Alternatives Trade Space 

6.7.1. Deflection Alternatives Trade Space

Figure 36 illustrates the deflection alternatives trade space considered in this study. 

Deflection Deflection

Nuclear Nuclear Non- Nuclear Non- Nuclear

Fast Reaction Fast Reaction

Explosives Explosives

Nuclear Standoff Nuclear Standoff

Nuclear Surface Nuclear Surface

Nuclear Subsurface Nuclear Subsurface

Fast Reaction Fast Reaction Slow Reaction Slow Reaction

Nuclear Electric 

Propulsion

Nuclear Electric 

Propulsion

Slow Reaction Slow Reaction

Explosives Explosives

Kinetic Energy Kinetic Energy

Conventional Standoff Conventional Standoff

Conventional Surface Conventional Surface

Conventional Subsurface Conventional Subsurface

Kinetic Impactor Kinetic Impactor

Mass Driver Mass Driver

Focused Solar Focused Solar

Yarkovsky Effect Yarkovsky Effect

Gravity Tractor Gravity Tractor

Space Tug Space Tug

Mass Driver Mass Driver

Focused Laser Focused Laser

Nuclear Nuclear Non- Nuclear Non- Nuclear

Fast Reaction Fast Reaction

Explosives Explosives

Nuclear Standoff Nuclear Standoff

Nuclear Surface Nuclear Surface

Nuclear Subsurface Nuclear Subsurface

Fast Reaction Fast Reaction Slow Reaction Slow Reaction

Nuclear Electric 

Propulsion

Nuclear Electric 

Propulsion

Slow Reaction Slow Reaction

Explosives Explosives

Kinetic Energy Kinetic Energy

Conventional Standoff Conventional Standoff

Conventional Surface Conventional Surface

Conventional Subsurface Conventional Subsurface

Kinetic Impactor Kinetic Impactor

Mass Driver Mass Driver

Focused Solar Focused Solar

Yarkovsky Effect Yarkovsky Effect

Gravity Tractor Gravity Tractor

Space Tug Space Tug

Gravity Tractor Mass Driver

Focused Laser Space Tug

Nuclear Nuclear Non- Nuclear Non- Nuclear

Fast Reaction Fast Reaction

Explosives Explosives

Nuclear Standoff Nuclear Standoff

Nuclear Surface Nuclear Surface

Nuclear Subsurface Nuclear Subsurface

Fast Reaction Fast Reaction Slow Reaction Slow Reaction

Nuclear Electric 

Propulsion

Nuclear Electric 

Propulsion

Slow Reaction Slow Reaction

Explosives Explosives

Kinetic Energy Kinetic Energy

Conventional Standoff Conventional Standoff

Conventional Surface Conventional Surface

Conventional Subsurface Conventional Subsurface

Kinetic Impactor Kinetic Impactor

Mass Driver Mass Driver

Focused Solar Focused Solar

Yarkovsky Effect Yarkovsky Effect

Gravity Tractor Gravity Tractor

Space Tug Space Tug

Mass Driver Mass Driver

Focused Laser Focused Laser

Nuclear Nuclear Non- Nuclear Non- Nuclear

Fast Reaction Fast Reaction

Explosives Explosives

Nuclear Standoff Nuclear Standoff

Nuclear Surface Nuclear Surface

Nuclear Subsurface Nuclear Subsurface

Fast Reaction Fast Reaction Slow Reaction Slow Reaction

Nuclear Electric 

Propulsion

Nuclear Electric 

Propulsion

Slow Reaction Slow Reaction

Explosives Explosives

Kinetic Energy Kinetic Energy

Conventional Standoff Conventional Standoff

Conventional Surface Conventional Surface

Conventional Subsurface Conventional Subsurface

Kinetic Impactor Kinetic Impactor

Mass Driver Mass Driver

Focused Solar Focused Solar

Yarkovsky Effect Yarkovsky Effect

Gravity Tractor Gravity Tractor

Space Tug Space Tug

Gravity Tractor Mass Driver

Focused Laser Space Tug

Figure 36. Deflection Alternatives Trade Tree 
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6.7.2. Deflection Alternatives Analyzed

A representative set of potential PHO deflection approaches was presented during the 

NEO public workshop described in Appendix D. NASA evaluated some of these 

alternatives in prior studies. [26] [27] As directed, this study has examined a number of 

techniques for deflecting a PHO that have been categorized as either “impulsive” or 

“slow push”. Table 19 provides an overview of the impulsive methods, considered in this 

study. Likewise, Table 20 shows the slow push techniques, where the velocity change 

results from the continuous application of a small force. Each of these concepts is 

developed further in Appendix G. 

Table 19. Impulsive Deflection Alternatives Considered 

Impulsive Technique Description 

Conventional Explosive 

(surface) 

Detonate on impact 

Conventional Explosive 

(subsurface) 

Drive explosive device into PHO, detonate 

Nuclear Explosive (standoff) Detonate on flyby via proximity fuse 

Nuclear Explosive (surface) Impact, detonate via contact fuse 

Nuclear Explosive (delayed) Land on surface, detonate at optimal time 

Nuclear Explosive (subsurface) Drive explosive device into PHO, detonate 

Kinetic Impact High velocity impact 

Table 20. Slow Push Deflection Alternatives Considered 

Slow Push Technique Description 

Focused Solar Use large mirror to focus solar energy on a 

spot, heat surface, “boil off” material 

Pulsed Laser Rendezvous, position spacecraft near PHO, 

focus laser on surface, material “boiled off” 

surface provides small force 

Mass Driver Rendezvous, land, attach, mine material, 

eject material from PHO at high velocity 

Gravity Tractor Rendezvous with PHO, fly in close 

proximity for extended period, gravitational 

attraction provides small force 

Asteroid Tug Rendezvous with PHO, attach to PHO, push 

Enhanced Yarkovsky Effect Change albedo of a rotating PHO; radiation 

from sun-heated material will provide small 

force as body rotates 
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6.8. Technology Readiness and Robustness 

6.8.1. Technology Readiness

Included in the discussion of each technique is an assessment of its current level of 

technology readiness. This study developed the assessments by examining the readiness 

of technologies needed to implement each technique - an approach similar to NASA’s 

system of Technology Readiness Levels (TRL). Reference [28] develops the concept of 

TRL in detail, and it can be summarized as follows: 

If the technology has flown and has been demonstrated in space, it has a high level of 

technology readiness. If it has not been used before and requires further development 

before it can be implemented, the technology readiness level is low or medium. 

Each of the concepts was developed using information from literature searches, inputs 

from the NEO public workshop, interviews with members of the community, and by 

performing any additional analysis required to bring each concept to a common level of 

detail at the element and system levels. The results for each of the individual technology 

risks were combined to produce one of the following scores for each technology area: 

Low Level of Technology Readiness — Technologies are notional or have evolved to 

the point where proof-of-principle experiments have been performed. Deflection options 

in this range might be viable, but system-level technical and experimental verification 

have not been completed and documented. 

Medium Level of Technology Readiness — Technologies involved have been validated 

at the component, breadboard, or prototype level in a laboratory experiment or a test in a 

relevant environment. Furthermore, analysis and testing have validated the basic concepts. 

High Level of Technology Readiness – A subsystem/system prototype has been 

demonstrated in the space environment or in an actual test of the complete system 

through flight operations. Elements of the concept have been demonstrated by testing 

and/or actual operations, or they are thought to be readily achievable by analog. 

6.8.2. Overall Effectiveness

Overall effectiveness is defined as the ability to apply the concept to the range of PHO 

threats, taking into account the range of orbit trajectories, ranges for PHO size and mass, 

and the variety of PHO types. Deflection alternatives will be evaluated as Very High 

Effectiveness, High Effectiveness, Medium Effectiveness, or Low Effectiveness 

according to their ability to address the different PHO threats. Effectiveness scores are 

not absolute measures; they are relative to other alternatives. 

6.8.3. Readiness and Effectiveness Summaries

Table 21 combines the technology readiness scores from Table 47 in Appendix Section 

H.4 with assessments of the effectiveness of each alternative that uses an impulsive 

technique. Likewise, the results in Table 48 in Appendix Section H.4 are used to create 

Table 22. The rationale for each readiness and effectiveness score is discussed in 

Appendix Sections G.1 and G.2 respectively and are summarized in Appendix H.4. 
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Table 21. Impulsive Alternatives Readiness and Effectiveness Summary 

“Impulsive” Concepts Readiness Effectiveness 

Conventional Explosive - Contact High Medium 

Conventional Explosive - Subsurface Medium Medium 

Kinetic Impact High High 

Nuclear Surface Contact High Very High 

Nuclear Standoff High Very High 

Nuclear Subsurface Medium Medium 

Nuclear Surface Delayed Medium High 

Table 22. “Slow Push” Alternatives Readiness and Effectiveness Summary 

“Slow Push” Concepts Readiness Effectiveness 

Enhanced Yarkovsky Low Low 

Focused Solar Low Medium 

Gravity Tractor Medium Medium 

Mass Driver Low Medium 

Pulsed Laser Low Medium 

Space Tug Low Medium 

6.9. Linkage of Characterization and Mitigation 

Two types of characterization information are necessary for mitigation and apply to 

different phases of an effort to mitigate the hazard. If development of a deflection option 

(or options) proceeds before an actual threat is identified, some understanding of the 

general population of targets is necessary. For example, if the distribution of PHO sizes 

were not well understood, it would be difficult to know if a deflection concept would 

perform sufficiently to eliminate likely threats. Likewise, if concepts, such as the space 

tug, are chosen for development, knowing the statistical distribution of rotation rates is a 

key development parameter. 

Second, characterization is needed when an actual threat is identified. For some 

alternatives, little characterization beyond mass and orbit will be required, although 

additional information may improve performance. For others, precursor missions will be 

required to assure the successful design and implementation of the mission. 

Remote characterization of a PHO is only useful for deflection in two instances. The first 

is to provide information after a threat has been identified, but before an in-situ 

characterization mission can be launched. The second is to provide the only available 

information to inform a deflection mission, particularly when insufficient time is 

available to visit the threat. As the survey catalogue is completed, the likelihood of a 

short warning scenario is reduced to less than 1% of all warnings. 

To inform such a short warning case, remote means must be developed to provide the 

required information. Models can be developed from remote observations of a large 

number of NEOs or from in-situ data. These models then may be used to predict the 
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characteristics of subsequent threats. This information may allow planners to begin 

designing a deflection mission without first launching an immediate in-situ mission. 

Table 23 and 

Table 24 show a qualitative assessment of the characterization requirements for 

impulsive and slow push techniques respectively. These would apply to any development 

and design programs, regardless of whether they were initiated ahead of an actual threat 

or with the discovery of one. In all cases, it is assumed that sufficient information is 

available about the PHO’s orbit to design a deflection mission. 

Table 23. Characterization Required for Impulsive Alternatives 

 Mass Spin Density Material 

Properties 

Size & 

Shape 

Surface 

Properties 

Conventional Expl. 

Surface - Contact Yes No Helpful Helpful Helpful Helpful 

Conventional Expl. 

Subsurface Yes No Helpful Helpful No No 

Kinetic Impactor Yes No Helpful Helpful Helpful No 

Nuclear (Contact) Yes No Helpful Helpful Helpful No 

Nuclear (Standoff) Yes No No No No No 

Nuclear Explosive 

(Sub-Surface) Yes No Helpful Helpful No No 

Nuclear Explosive 

(Surface Delayed) Yes Yes Helpful Helpful No Helpful 

Table 24. Characterization Required for Slow Push Alternatives 

 Mass Spin Density Material 

Properties 

Size & 

Shape 

Surface 

Properties 

Yarkovsky Yes Yes No No Yes Yes 

Focused Solar Yes Helpful No No No Yes 

Gravity Tractor Yes Yes No No Yes No 

Mass Driver Yes Yes Yes Yes Helpful Helpful 

Pulsed Laser Yes Helpful No No No Yes 

Space Tug Yes Yes No No Yes Yes 

For example, if there is a factor-of-two uncertainty over the object’s mass, a mission must 

be designed to deflect the heavier body. Similarly, the design of a slow push technique 

would be based on moving the largest mass anticipated and lengthening the mission 

duration by the scale of the uncertainty. In both cases, the weight of the payload launched 

from Earth and the overall cost of the campaign would increase. 

In all cases, information on material properties, size, shape, albedo, rotation rate, axis of 

rotation, etc., reduces the uncertainties associated with the deflection mission, increases 

the likelihood of success, and decreases the overall cost of the effort. 
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6.10. Characterization Options Capabilities Matrix 

The characterization options developed in Section 5.15 are connected to specific 

deflection alternatives in this section. If any the seven options are executed, Table 25 

describes the deflection options enabled by each if no other characterization is available 

before the deflection mission is initiated. These primarily correspond to a “short 

warning” scenario where a dedicated in-situ characterization mission is not possible, 

except in the case of Option 7 which assumes an in-situ visit. An in-situ mission before 

the start of a deflection program obviates much of the need for remote characterization if 

sufficient time is available. 

Table 25. Deflection Alternatives Enabled by Characterization Options 

Characterization Capability Options 

Deflection Alternative* 1 2 3 4 5 6 7 

Nuclear Subsurfacea Y E E E E E E 

Nuclear Surfaceb Y E E E E E E 

Nuclear Surface delayedc N N N N N N Y 

Nuclear Standoffd Y E E E E E E 

Kinetic Impacte Y E E E E E E 

Subsurface Explosivef Y E E E E E E 

Surface Explosiveg Y E E E E E E 

Space Tug – Non-rotatingh N N N N N N Y 

Space Tug – Rotatingi N N N N N N Y 

Gravity Tractorj Y E E E E E 

Life-cycle cost FY06$B 0.1 0.5 1-2 1-2 2-3 5-8 5-8 

 * rationale for scores provided below 

A column in Table 25 is marked “Y” only if the characterization option is both necessary 

and sufficient to enable an effective deflection. If it exceeds the requirements, it is 

marked “E”. If it is not sufficient, it is marked “N”. These scores do not reflect the 

effectiveness of the option, which may be enhanced with additional characterization. 

Rationale for Scores in Table 25: 

a. Nuclear subsurface is aided by information about the specific PHO’s surface and 

composition, but remote characterization methods likely are sufficient. 

b. Nuclear surface explosions could use remotely gathered data and effectiveness 

would be augmented by in-situ verification. 

c. A delayed surface blast cannot occur until the orbit is properly phased, and 

therefore will require knowledge about the PHO’s surface. 

d. A nuclear standoff explosion would not require any information other than the 

PHO’s mass and orbit, which can be estimated by all options. 

e. A kinetic impact concept would not require any information other than mass and 

orbit, which can be obtained by all options. If the object is highly porous, a kinetic 

impactor may be ineffective. 
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f. A conventional subsurface explosive is aided by information about a specific 

PHO’s surface and composition, but remote methods are likely sufficient. 

g. A conventional surface explosion could use remote data and effectiveness would 

be augmented by in-situ verification. 

h. A non-rotating space tug requires information about the surface and composition 

of the specific threat; remote characterization methods are likely insufficient. 

i. A rotating space tug requires information about the surface and composition of 

the specific threat; remote characterization methods are likely insufficient. 

j. A gravity tractor would be enabled with knowledge about the object’s shape and 

rotation, which could be obtained on station for most objects. Detection assets 

may be adequate depending on type and vantage point. 

6.11. Deflection Performance Analyses 

6.11.1. Deflection Performance Analysis Methodology

To compare the relative performance of the proposed asteroid deflection concepts, the 

change in PHO momentum, based on the maximum available launch mass, is calculated. 

Available momentum change for both the Delta IV Heavy (largest current launch 

capability) as well as the Ares V Cargo Launch Vehicle (proposed lunar cargo vehicle) 

will be calculated using the methods described in Sections 6.11 and H.3. 

6.11.2. Deflection Performance Analysis Assumptions

Several assumptions are inherent in the following performance summary. 

• Impulsive deflection concepts assume that the launch vehicle is used for direct 

intercept. For delayed surface methods, this is an optimistic assumption. 

• During targeting maneuvers, it is assumed that 1,400 kg of propellant will be 

consumed. This amount of propellant is an average, which was developed across a 

series of simulations, and will vary for systems with system mass and operations. 

• Impulsive concepts assume that the PHO will not experience large-scale 

fracturing. For kinetic impact concepts, collision elasticity was parameterized. 

Momentum efficiency was assumed to be perfectly plastic or to experience 

significant PHO ejecta. This assumption may not be valid when a PHO is a loose 

rubble pile. In this case, mass may be ejected in the direction of impact and the 

concepts may be less effective. 

• For standoff nuclear concepts, a 200-meter asteroid, with 50% porosity, was 

assumed. This is based on the results from Holsapple and Gennery. [29] [30] The 

standoff distance was chosen as 10% of the object’s diameter, and assumes that 

spin axis and shape can be determined en route. There is a strong correlation 

between standoff distance and effectiveness, and a non-optimal standoff distance 

could reduce effectiveness by one-to-two orders of magnitude. 
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• It was assumed that current nuclear explosive technology could produce a device 

that yields 1-2 kt TNT per kilogram of dry mass. When considering standard 

nuclear explosives, it was assumed that 1% of the yield are neutrons. 

• The gravity tractor standoff distance was assumed to be 1.5 times the radius of the 

object to be towed. Other analyses have considered a standoff distance as small as 

0.5 radii. This assumption would yield a performance improvement of a factor of 

three, but will not alter the findings of this report. 

6.11.3. Deflection Performance Analysis Results

Figure 37 and Figure 38 are graphical representations of the deflection capabilities. The 

system performance required to deflect any object on a given trajectory may be described 

as the velocity change necessary to change its path, multiplied by its mass. The “effective 

momentum change” performance parameter allows many different scenarios to be plotted 

simultaneously across a wide range of asteroid masses and required deflection velocities 

(ΔV). It is displayed logarithmically on the Y-axis of these figures. The logarithmic Xaxis represents launch performance to place the deflection payload on an intercept 

trajectory. The launch C3 corresponding to payload capabilities of the two launch 

systems considered are at the top of each figure. 

The lines to the right of each figure may be used to translate effective momentum change 

to the design parameters of PHO mass (and size) and deflection velocity. Lines of 

constant object mass (and size) spaced logarithmically run diagonally across vertical lines 

representing a logarithmic range of deflection ΔV. As an example, following the diagonal 

line representing a mass of 1010 kg (200 m) to its extreme lower left at the vertical 1 cm/s 

ΔV line, this corresponds to an effective momentum change of 108

 kg m/s on the far left. 

The lines plotted represent the performance of the deflection alternatives. If an alternative 

has a higher effective momentum change capability than is required, it is considered 

“feasible” for a single-launch deflection. Therefore, using the previous example of an 

effective momentum change of 108

 kg m/s and assuming that a Delta IV Heavy launch 

vehicle is available and that C3 = 25 km2

/s2 is required to intercept, all but the kinetic 

interceptor and the conventional explosives would meet performance requirements. 

Likewise, none of the slow push techniques could meet this hypothetical requirement. 

Figure 37 shows that impulsive techniques using proximal nuclear explosives generally 

provide greater potential for momentum transfer per kilogram of payload weight 

delivered to the threat than any other option considered. Standoff nuclear concepts have a 

generally lower risk of fragmenting a PHO than impulsive techniques involving direct 

contact, but also produce a lower effective momentum change than surface or subsurface 

nuclear explosives. Performance may vary significantly, depending on the type of nuclear 

explosive used and whether it is “off-the-shelf” as opposed to optimized for the PHO 

deflection mission. The performance of kinetic impactors is somewhat less robust than 

any of the nuclear explosions; however, their effectiveness depends strongly on the 

structure of the PHO. Kinetic impactors also may be significantly less effective for loose 

rubble piles. Conventional explosives have the lowest performance among the impulsive 

techniques due to their relatively low-energy density. 
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Figure 38 illustrates that slow push techniques may be useful for imparting momentum 

changes smaller than 109

 kg m/s. The asteroid tug appears to have significantly greater 

performance than the gravity tractor for a given launch mass, even accounting for pulsed 

operation on a rotating PHO. The disadvantage of the asteroid tug is the additional 

complexity required to anchor the tug to the NEO, particularly if the PHO structure has 

not been well characterized or the target is rotating very rapidly. 

These figures show that nuclear explosives and kinetic impactors generally provide 

greater potential for momentum transfer per kilogram of payload weight delivered to the 

NEO than other alternatives. Each concept is described in detail in Appendix G. 
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Effective Momentum Change (kg-m/s) 

Figure 37. Deflection Performance of Impulsive Alternatives 
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Effective Momentum Change (kg-m/s) 

Figure 38. Deflection Performance of Slow Push Alternatives 
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6.12. Deflection Mission Operations Risk 

6.12.1. Overview

The purpose of this section is to explore reliability requirements for deflection. For 

example, if the goal is to reduce the risk of impact from 1 in 1,000 to 1 in 1 million, the 

chance of deflection failure must be less than 1 in 1,000. This requirement is two orders 

of magnitude more stringent than the typical 1 in 10 chance of failure experienced by 

planetary missions. For higher probabilities of impact, higher reliabilities are required to 

meet the requirement, although an incremental mission approach may reduce the overall 

cost of successful deflection if sufficient time to deflect is available. 

An approach is developed to illustrate mission reliability tradeoffs. This is done by 

generating plots overall failure risks as a function of 

• Reliability of individual space missions (90-99%) 

• Architecture of the deflection “campaign” (number, type of missions) 

• Number of independently developed approaches 

• Number of attempts for each approach 

Also, it is considered that several ways may exist to increase a campaign’s reliability or 

chance of collateral damage: 

• Diverse deflection approaches to reduce risk of common-cause failures 

• Redundant missions for each approach to reduce risk of random failures 

• Precursor characterization missions, if time permits, to tune the deflection 

approach and increase reliability 

The first three of these are explored in a parametric manner. 

6.12.2. Notional Deflection “Campaign”

The campaign to deflect an asteroid will involve more than just a single intercept or 

rendezvous mission, as suggested by popular movies. If sufficient warning time is 

available, a series of precursor characterization missions, followed by a wave of one type 

of deflection mission. If the first wave fails, then a series of more robust interceptors 

might be launched in a second wave. 

Thus, depending on the warning time and the size of the asteroid, some number of 

redundant missions and diverse design approaches will be needed to eliminate the threat. 

Figure 39 illustrates one possible scenario. 
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Mission Failures = Random Failures + Common Failures

DECISION 

TO ACT

Second 

Deflection 

Approach

First 

Deflection 

Approach

Redundant?

Characterization 

Missions

Diverse 

Approaches

Redundant 

Missions

IMPACT

Figure 39. Illustration of Types of Deflection Mission Redundancy 

6.12.3. Parametric Reliability Calculations

Two basic failure modes need to be considered: 

• Independent random failures 

• Common-cause failures that affect all flights of a given approach 

In the following parametric reliability plots it is assumed that: 

• All space missions (characterization or deflection) have the same random and 

common-cause failure rates, except in the excursion where it is assumed that the 

common-cause failure rate for deflection is reduced if information is received 

from a successful precursor characterization mission 

• Common cause modes are independent among the different characterization and 

deflection approaches. 

6.12.4. Precursor Characterization Not Required

The risk of overall failure in this case is 

m Risk = Pn

Where 

m = # of independent approaches 

n = # of attempts per approach 

Pn = 1 – (1 – Fn)(1 – C) 

F = random failure rate 

C = common failure rate 
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The overall probability of failure is plotted in Figure 40 as a function of the failure rate 

for a single generic space mission, assuming that 80% of the failures are random and 20% 

are due to a common cause in the particular design. 

1.0E-11

1.0E-10

1.0E-09

1.0E-08

1.0E-07

1.0E-06

1.0E-05

1.0E-04

1.0E-03

1.0E-02

1.0E-01

0% 1% 2% 3% 4% 5% 6% 7% 8% 9% 10%

Space Mission Failure Rate

Overall Probability of Failure

 1 Approach, 1 Mission

 1 Approach, 2 Missions

 2 Approaches, 2 Missions

Assumptions

• 80% random and 20% common cause failures

• Characterization mission not required for successful 

deflection All space missions have the same reliability

Figure 40. Chance of Deflection Failure if Characterization is Not Required 

To achieve a hypothetically “required” failure rate of 0.001 or less, two or more 

independent interceptor designs and at least two redundant missions for each design are 

required. Just increasing the number of redundant missions with a single design will not 

reduce the overall risk below a certain level, regardless of the assumed space mission 

failure rate. 

6.12.5. Precursor Characterization Required

The risk of overall failure in this case is 

2 1 (1 ) m Risk = − − Pn

Where 

m = # of independent approaches 

n = # of attempts per approach 

Pn = 1 – (1 – Fn)(1 – C) 

F = random failure rate 

C = common failure rate 

The risk is plotted in Figure 41 as a function of the failure rate for a single space mission, 

assuming that 80% of the failures are random and 20% are due to a common cause in a 

particular design. The overall probability of failure is higher in this case because a 

successful outcome depends on having at least one successful characterization mission 

and one successful deflection mission. 
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1.0E-11

1.0E-10

1.0E-09

1.0E-08

1.0E-07

1.0E-06

1.0E-05

1.0E-04

1.0E-03

1.0E-02

1.0E-01

0% 1% 2% 3% 4% 5% 6% 7% 8% 9% 10%

Space Mission Failure Rate

Overall Probability of Failure

 1 Characterization, 1 Deflection, 1 Mission/Approach

 1 Characterization, 1 Deflection, 2 Missions/Approach

 2 Characterizations, 2 Deflections, 2 Missions/Approach

Assumptions

• 80% random and 20% common cause failures

• Characterization mission required for successful deflection

• All space missions have the same reliability

Figure 41. Chance of Deflection Failure when Characterization is Enabling 

Assuming the space mission failure rate does not exceed 8%, two or more independent 

interceptor designs and at least two redundant missions for each design would be 

adequate to meet the hypothetical “requirement” of 0.001 or less. 

6.12.6. Characterization Reduces Common-Cause Failure Rate

The risk of overall failure in this case is 

( )( ) (1 )( ) 'm

n

m

n

m

n

m Risk = Pn P + − P P

Where 

m = # of independent approaches 

n = # of attempts per approach 

Pn = 1 – (1 – Fn)(1 – C) 

F = random failure rate 

C = common failure rate 

Pn' = 1 – (1 – Fn

)(1 – rC) 

r = reduction factor (0 – 1) for common failure rate for successful deflection 

characterization 

In this case the overall probability of failure is lower because characterization improves 

the reliability of deflection (over the baseline reliability), but is not required for deflection, 

as shown in Figure 42. Unless the space mission failure rate can be reduced to less than 

3%, two or more independent interceptor designs and at least two redundant missions for 

each design would still be necessary to meet the hypothetical 0.001 “requirement.” 
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1.0E-11

1.0E-10

1.0E-09

1.0E-08

1.0E-07

1.0E-06

1.0E-05

1.0E-04

1.0E-03

1.0E-02

1.0E-01

0% 1% 2% 3% 4% 5% 6% 7% 8% 9% 10%

Space Mission Failure Rate

Overall Probability of Failure

 1 Characterization, 1 Deflection, 1 Mission/Approach

 1 Characterization, 1 Deflection, 2 Missions/Approach

 2 Characterizations, 2 Deflections, 2 Missions/Approach

Assumptions

• 80% random and 20% common cause failures

• Characterization improves probability of successful deflection

• All space missions have the same reliability

Figure 42. Chance of Deflection Failure if Characterization is Enhancing 

6.12.7. Summary

Linking the reliability of the deflection method to the probability of impact reduction 

may seem unrelated. However, as an example, assume that the impact cannot be 

calculated to better than 1 in 10 in time for deflection; therefore, there are nine “near 

misses” for each “impact” predicted. Assume, too, that the probability rate must be 

reduced to 1 in 1,000. With this allowable threshold, for every 1,000 warnings, one 

impact will occur (likely an unacceptable level, but chosen to illustrate the example). 

If no deflection missions are attempted, 100 impacts and 900 “near misses” will occur for 

the 1,000 warnings. If deflection campaigns are very unreliable (10% reliability, for 

example), 90 impacts will occur, leaving a risk of 9 in 100, well above the acceptable 

threshold in this example. If deflection is 99% effective, only one of the 100 actual 

impacts will occur and the required risk of 1 in 1,000 will be achieved. Finally, if no 

probability of impact is acceptable, then the deflection campaigns will need to be 100% 

reliable or some impacts will occur. 

This analysis emphasizes that the reliability requirements for asteroid deflection are much 

higher than the current success rates for interplanetary missions. To achieve levels of 

confidence in overall mission success commensurate with significant reductions in 

impact probability (e.g. three orders of magnitude), diverse design approaches and 

multiple missions of each design will be required. If in-situ characterization is required 

for deflection, then reliability requirements are even more stringent. 
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6.13. Possible Scenarios – Application of the Alternatives 

This section provides several illustrations of how the alternatives might be applied to 

hypothetical deflection scenarios drawn primarily from Reference [31]. The inclusion of 

actual objects in these scenarios does not indicate any increase or decrease in the 

understanding of the hazard they pose; instead they were chosen both because they are 

publicly known and are representative of classes of potential threats. 

The scenarios include missions to deflect: 

A. The 330-meter asteroid, Apophis, before its close approach to Earth in 2029 (a 

possible keyhole, see Section 5.2.3). 

This scenario was divided into two design points: 

A1. For the first, a relatively large momentum change is required to deflect the 

object with the required certainty. Apophis must be deflected by at least 1 

Earth radius or about 6,400 km to achieve a probability of collision of less 

than 10-6. 

A2. For the second, very accurate information about the object’s orbit is 

assumed and the impetus necessary to divert the asteroid with certainty is 

therefore substantially reduced. Apophis must be deflected by at least 5 km 

to achieve a probability of collision of less than 10-6. 

B. Apophis after the close approach and before the 2036 Earth encounter, assuming a 

collision is predicted 

C. The 500-meter asteroid (VD17) that could be a threat in the year 2102 

D. A hypothetical 200-meter asteroid, representative of 100-meter-class asteroids 

E. A hypothetical asteroid larger than 1 km in diameter 

F. A hypothetical long-period comet with a very short time (9-24 months) between 

detection and possible impact 

Figure 37 and Figure 38 of Section 6.11 summarize the momentum capability of each 

alternative, with increasing launch performance. The approximate performance 

requirements for each of the scenarios overlaid on these to produce Figure 43 for the 

impulsive techniques and Figure 44 for the slow push methods. 

An explanation of how to read these figures is presented in Section 6.11. A detailed 

analysis of how different deflection techniques would perform against these scenarios is 

presented in the following sections. 
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2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 95 

Effective Momentum Change (kg-m/s) 

Figure 43. Momentum Capability of Impulsive Alternatives Applied to Scenarios 
DRAFT Pre-Decisional Material

2006 Near-Earth Object Survey and Deflection Study 

96 DRAFT Pre-Decisional Material

Effective Momentum Change (kg-m/s) 

Figure 44. Momentum Capability of Slow Push Alternatives Applied to Scenarios 
DRAFT Pre-Decisional Material

2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 97 

6.13.1. Scenario - Apophis (Deflect before 2029)

Asteroid 99942, also known as Apophis (2004 MN4), is estimated to be about 320 meters 

in diameter, with a mass of 4.6x1010 kg. Uncertainty about its diameter is currently a 

factor of two, which means that the mass could vary by a factor of 16 (5.8x109

 - 

3.7x1011 kg). The equivalent impact energy is proportional to the mass. Specific 

information on its shape and rotation currently are not available. Table 26 describes this 

scenario further. 

Table 26. Apophis before 2029 Scenario Description 

Scenario Apophis (before 2029) 

Predicted Frequency Frequency of keyholes is undetermined 

Time to Act 22 years 

Action Begins 6 years prior to impact 

Diameter of Threat 320 m 

Mass of Threat 4.6x1010 kg 

ΔV Design Point 1 5.000 mm/s (DP1) 

ΔV Design Point 2 0.026 mm/s (DP2) 

Δ Momentum DP1 2.3 x 108

 kg m/s 

Δ Momentum DP2 1.2 x 106

 kg m/s 

Unique Features • Keyhole scenario complicates decision 

to deflect in 2029 

Apophis is currently predicted to have a close approach to Earth in 2029, passing within 

30,000 km, with a subsequent 2.2x10-5 probability of impact on April 13, 2036. The 

probability of impact in 2036 will be strongly influenced by the precise location of the 

close approach in 2029. If it should pass within a 600-meter-wide “keyhole” in 2029 (see 

Section 5.2.3), the likelihood of impact in 2036 will be much higher. [32] 

One approach for avoiding a threat in 2036 is to deflect Apophis so that it is guaranteed 

to miss the keyhole in 2029. An advantage of this approach is that the asteroid requires 

only a very relatively small change in the velocity to miss the keyhole, as shown in 

Figure 45. Assuming optical and radar observations are taken in 2013, 2020, and 2021, it 

is anticipated that one could achieve a tracking accuracy of 5 km. [16] 

To take advantage of either opportunity, acquisition of a deflection system must be 

started years in advance to account for vehicle development and transit time to the 

asteroid. Consequently, such a program may need to begin with incomplete information. 

An in-situ characterization mission may provide better tracking accuracy early on, 

allowing for a less costly deflection mission or elimination of the threat entirely. Figure 

45 shows that the ΔV grows substantially as the time to close approach decreases, which 

is typical of deflection scenarios. 
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2006 Near-Earth Object Survey and Deflection Study 

98 DRAFT Pre-Decisional Material 

0.0001

0.001

0.01

0.1

1

10

0 2000 4000 6000 8000

Days to Keyhole

ΔV Required to Deflect (cm/sec)

Deflect to miss by 2 Earth radii

Deflect to miss by 1 Earth radius

Deflect to miss by 0.1 Earth radii

Deflect to miss by 5 km

Reduce risk to ~1e-6

Design Point 2

Design Point 1

Figure 45. ΔV to Avoid the Keyhole in 2029 

Two design points will be examined for this case: 

1. Assuming an action time of 10 years before the asteroid arrives at the keyhole, a 

change in momentum of roughly 2.3 x 108

 kg m/s would be required to alter the 

asteroid’s velocity by 5 mm/s to reduce the impact probability to the required 1 in 

1 million. In this case, it is assumed that uncertainty about the orbit is sufficient to 

move the orbit one Earth radius at the keyhole. This is a conservative projection 

as it assumes little improvement in orbital accuracy with additional observations. 

2. If the error ellipse is refined by additional observations and assuming an action 

time of 6 years before the asteroid arrives at the keyhole and the availability of 

improved tracking accuracy, a change in momentum of roughly 1.2 x 106

 kg m/s 

(velocity change = 0.026 mm/s) would be required to deflect Apophis, causing the 

refined error ellipse to miss the keyhole. This assumes that very accurate orbit 

information was gathered during the previous approaches (by radar) and that an 

orbit accuracy of 5 kilometers is achievable. This would still require that 

deflection preparations be initiated well before the final orbit determination is 

accomplished. 

An on-board propulsion system that can produce a change in velocity of 1 km/s is 

necessary to accomplish an impact with the asteroid. Assuming a liquid propulsion 

system with a specific impulse of 325 seconds, an estimated 1900 kg of fuel will be 

required, leaving a vehicle dry mass of 5200 kg. Of the vehicle dry mass, roughly 57% of 

the vehicle will consist of structures and navigational systems, leaving a possible payload 

mass of 2,200 kg. Reference [32] designed a similar mission and proposed that an 

observer spacecraft be used as a second deflector; however, this would require additional 

launch capability. 
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2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 99 

A performance index (P), defined as the momentum change delivered by the alternative, 

divided by the momentum change required by the scenario, will be used to help compare 

the various techniques. 

required

delivered

M

M P = (1)

A performance index of 1.0 (or greater) indicates that the technique can deliver the 

required impulse (or more) nominally required by the scenario. Values above 1.0 indicate 

the amount of margin available. A performance index less than 1.0 indicates that 

insufficient momentum change is provided with a single launch. 

Table 27 and Table 28 show the performance results of the concepts for the two design 

points in this scenario and the number of launches required for each launch system. 

Table 27. Apophis Keyhole – Deflection Performance Design Point 1 

Performance 

Index (P) 

Launches 

Required 

Launch Vehicle Î

Concept 

Delta 

IV H 

Ares 

V 

Delta 

IV H 

Ares 

V 

Nuclear Subsurface1 1793 17075 1 1 

Nuclear Surface1 897 8539 1 1 

Nuclear Standoff - Neutron1 52 244 1 1 

Nuclear Standoff – X-ray1 19 87 1 1 

Nuclear Standoff - Standard1 14 70 1 1 

Kinetic Impact, 50 km/s, β=101 9.6 80 1 1 

Kinetic Impact, 10 km/s, β=101 1.9 16 1 1 

Kinetic Impact, 50 km/s, β=11 1.0 8.0 2 1 

Kinetic Impact, 10 km/s, β=11 0.2 1.6 6 1 

Space Tug – Non-rotating2 0.5 7.4 2 1 

Space Tug - Rotating2 0.2 2.4 6 1 

Gravity Tractor2 0.0 0.3 29 3 

Subsurface Explosive 1 0.0 0.1 76 8 

Surface Explosive 1 0.0 0.1 152 16 

1

 Assumed to require C3=25 for an intercept trajectory 

2

 Assumed to require launch C3=0 for a rendezvous using electric propulsion 

If conventional explosives are sent on an intercept trajectory using a Delta IV Heavy 

with C3 = 25 km2

/s2, a maximum of 3800 kg of explosives would arrive at the asteroid. 

The resulting explosion would be unable to deflect the asteroid for Design Point 1, but 

would be enough for Design Point 2. 

Maximum intercept velocity is more beneficial for the kinetic impact than for any other 

concept. Therefore, the launch date is extremely important to allow for the maximum 

amount of mass to impact the asteroid. If launched at the optimum time, the relative 

velocity at impact can be as high as 15 km/s. Results presented are parameterized for 10 
DRAFT Pre-Decisional Material

2006 Near-Earth Object Survey and Deflection Study 

100 DRAFT Pre-Decisional Material 

km/s and 50 km/s to allow for advances in propulsion systems or the use of swing by 

trajectories. Assuming that Apophis is relatively dense and that material will be ejected 

during impact, a beta factor can be assumed to be roughly three. [34] As discussed 

earlier, a kinetic deflection is dependent on the beta factor (β). 

Table 28. Apophis Keyhole – Deflection Performance for Design Point 2 

Performance 

Index (P) 

Launches 

Required 

Launch Vehicle Î

Concept 

Delta 

IV H 

Ares 

V 

Delta 

IV H 

Ares 

V 

Nuclear Subsurface1 343699 3272761 1 1 

Nuclear Surface1 171951 1636583 1 1 

Nuclear Standoff - Neutron1 10000 46667 1 1 

Nuclear Standoff – X-ray1 3667 16667 1 1 

Nuclear Standoff - Standard1 2667 13333 1 1 

Kinetic Impact, 50 km/s, β=101 1835 15346 1 1 

Kinetic Impact, 10 km/s, β=101 367 3069 1 1 

Kinetic Impact, 50 km/s, β=11 183 1534 1 1 

Kinetic Impact, 10 km/s, β=11 36 307 1 1 

Space Tug – Non-rotating2 101 1419 1 1 

Space Tug - Rotating2 32 452 1 1 

Gravity Tractor2 6.8 65 1 1 

Subsurface Explosive 1 2.5 24 1 1 

Surface Explosive 1 1.3 12 1 1 

1

 Assumed to require C3=25 for an intercept trajectory 

2

 Assumed to require launch C3=0 for a rendezvous using electric propulsion 

Nuclear explosives have the ability to transfer much more energy for a given mass than 

any other option. Using a Delta IV Heavy with a C3 of 25 km2

/s2 and assuming the 

spacecraft defined above, the maximum mass of a nuclear explosive that could be 

transported to the asteroid is 2,200 kg and would provide a change in momentum of 2.0 x 

1011 kg m/s if detonated on the surface. This would change the asteroid’s velocity by 3 

m/s. If the device is placed beneath the surface, it will impart twice as much momentum 

to the asteroid, doubling both the performance index and velocity change. 

Slow push techniques create a change in momentum by incrementally moving the 

threatening object off its orbit. For the approaches considered here, thrust is produced by 

either the gravitational attraction between bodies or by the propulsion system on the 

spacecraft. Both of these techniques require a great deal of mass to provide the necessary 

thrust. For this reason, it is assumed that the launch vehicle will transport the spacecraft 

to a lower-energy Earth orbit and that the spacecraft’s propulsion system will be used to 

intercept the asteroid. 

A Delta IV Heavy Launch Vehicle has the ability to lift roughly 25,000 kg to a LEO orbit 

if launched at optimum conditions. A 3-year transit period was assumed. The fuel 

required by the propulsion system is calculated by assuming constant thrust over the full 

transit period. 
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2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 101 

The gravity tractor has been suggested to deliver the small momentum changes required 

for the keyhole deflection of Apophis. For this concept, the most mass at the PHO 

provides the most momentum change; therefore, an ion engine would be used to propel 

the vehicle. To provide enough thrust to reach the asteroid in a short period of time and 

rendezvous with the asteroid, eight NEXIS thrusters, similar to those used on the Jupiter 

Icy Moons Orbiter (JIMO), is used as the tractor’s propulsion system. The propellant 

mass used for transit, rendezvous, and hovering would be roughly 5000 kg, allowing for a 

mass of 20,000 kg to produce the gravitational force on the asteroid. 

Different deflection opportunities would change the available lead-time for the gravity 

tractor. Assuming a few years transit time for the low-thrust system, the gravity tractor 

may need to be launched in advance of the 2021 observation opportunity. Here, a 3-year 

transit time and a 6-year action time are assumed. If a mass of 20,000 kg were applied for 

6 years against Apophis, it would impart a change in momentum of 4.8 x 107

 kg m/s. This 

would change the asteroid velocity by 1 mm/s (0.001 m/s), insufficient to meet the 2019 

opportunity (Design Point 1), but sufficient to avoid the keyhole, assuming improved 

tracking accuracy for the 2023 opportunity (Design Point 2). To achieve a change of one 

Earth radius in the short time span, the tractor’s mass would need to be about 150,000 kg. 

The space tug is another slow push technique, and its use assumes that Apophis is 

structurally able to support attachment. The vehicle’s design could be similar to JIMO’s. 

The JIMO vehicle used eight NEXIS ion thrusters to propel an 18,000-kg spacecraft. If a 

Delta IV Heavy were used to launch the spacecraft, only 7,000 kg of fuel would be 

required, with 4,000 kg required for the 3-year transit period and rendezvous/docking, 

leaving 3,000 kg of fuel to produce a change in Apophis’s momentum. If the thrusters 

perform at full power, the active push time will be approximately 2.25 years, producing a 

change in momentum of 2.3 x 108

 kg m/s. It is estimated that this would be sufficient for 

both the 2019 and the 2023 opportunities. A factor-of-eight increase in the mass of the 

PHO would increase the push time for the tractor by more than 14 years, exceeding the 

time available for this design. Building a system that is somewhat larger than JIMO using 

on-orbit assembly techniques could possibly solve the problem. 

Although relatively little momentum change is required to prevent Apophis from passing 

through the keyhole, many of the deflection techniques are simply too massive for current 

lift capabilities for the first design point. If the extremely limited requirements of Design 

Point 2 are considered, all concepts could hypothetically deflect the threat. A mission to 

prevent Apophis from passing through the 2026 keyhole appears to be possible with 

current technologies. 

6.13.2. Scenario - Apophis (Deflect after 2029)

If Apophis passes within the keyhole in 2029, a mission to deflect before the 2036 

conjunction may be required. If unprepared and deflection is required, taking action 

within this 7-year timeframe will be a challenge from several perspectives. First, decision 

makers must authorize the mission and funding must be allocated. Second, spacecraft and 

launch vehicles (likely multiple spacecraft and multiple launch vehicles to assure 

success) must be designed and built. Previous studies have shown that 10 years is an 

ambitious schedule for a deflection of this type, [23] but less than 7 years will remain if 

the 2029 keyhole event occurs. Table 29 contains a summary of this scenario. 
DRAFT Pre-Decisional Material

2006 Near-Earth Object Survey and Deflection Study 

102 DRAFT Pre-Decisional Material 

Figure 46 shows the velocity required to deflect asteroid Apophis after the 2029 close 

approach. The plot begins just after the 2029 close approach, approximately 2,600 days 

before the 2036 encounter. This analysis shows that the ΔV required to deflect Apophis 

in 2034 is approximately 4 cm/s in 2034, which corresponds to a momentum change 

requirement of 1.8 x 109

 kg m/s. The ΔV is less than 2 cm/s in 2029 after the keyhole. 

0

5

10

15

20

0 500 1000 1500 2000 2500

Days to Impact

ΔV Required to Deflect (cm/sec)

Miss by 1 Earth radius

Reduce risk to 1:1,000,000

Reduce risk to 1:1,000

Deflect to skim the Earth

Design Point

Figure 46. ΔV required to Deflect Asteroid Apophis after 2029 Keyhole 

The performance required to deflect Apophis in 2034 is shown in Table 30. 

Table 29. Apophis after 2029 Scenario Description 

Scenario Apophis (after 2029) 

Impact Frequency ~10,000 years 

Time to Act 7 years 

Action Begins 2 years prior to impact 

Diameter of Threat 320 m 

Mass of Threat 4.6x1010 kg 

ΔV 4 cm/s 

Δ Momentum 1.8 x 109

 kg m/s 

Unique Features • Short warning 

• Comparison to keyhole 
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2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 103 

Table 30. Asteroid Apophis 2034 - Deflection Performance of Alternatives 

Performance 

Index (P) 

Launches 

Required* 

Launch Vehicle Î

Concept 

Delta 

IV H 

Ares 

V 

Delta 

IV H 

Ares 

V 

Nuclear Subsurface1 224 2134 1 1 

Nuclear Surface1 112 1067 1 1 

Nuclear Standoff - Neutron1 6.5 30 1 1 

Nuclear Standoff – X-ray1 2.4 11 1 1 

Nuclear Standoff - Standard1 1.7 8.7 1 1 

Kinetic Impact, 50 km/s, β=101 1.2 10.0 1 1 

Kinetic Impact, 10 km/s, β=101 0.2 2.0 5 1 

Kinetic Impact, 50 km/s, β=11 0.1 1.0 9 1 

Kinetic Impact, 10 km/s, β=11 0.0 0.2 42 5 

Space Tug – Non-rotating2 0.1 0.9 16 2 

Space Tug - Rotating2 0.0 0.3 48 4 

Gravity Tractor2 0.0 0.0 227 24 

Subsurface Explosive 1 0.0 0.0 605 64 

Surface Explosive 1 0.0 0.0 1209 127 

1

 Assumed to require C3=25 for an intercept trajectory 

2

 Assumed to require C3=0 for a rendezvous using electric propulsion 

* About half the number of launches is required if deflected in 2030 

6.13.3. Scenario - Asteroid VD17

Asteroid VD17 is estimated to be 580 meters in diameter with a mass of 2.6x1011 kg. [35] 

A close approach with Earth is predicted for May 4, 2102 and the current probability of 

impact is 1.6x10-5. No details are currently available on the object’s rotation rate or 

material properties. Table 31 contains a summary of the VD17 scenario and Figure 47 

shows the velocity increment that may be required. 

Table 31. VD17 Scenario Description 

Scenario VD17 

Impact Frequency ~100,000 years 

Time to Act > 90 years 

Action Begins 15 years prior to impact 

Diameter of Threat 580 m 

Mass of Threat 2.6x1011 kg 

ΔV 1 cm/s 

Δ Momentum 2.6 x 109

 kg m/s 

Unique Features • Long warning 

• Moderate mass 
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2006 Near-Earth Object Survey and Deflection Study 

104 DRAFT Pre-Decisional Material 

0.01

0.1

1

10

100

0 20 40 60 80 100

Years to Impact

ΔV Required to Deflect (cm/sec)

Reduce risk to 1:1,000,000

Miss by 1 Earth radius

Design Point

Figure 47. ΔV Required to Deflect Asteroid VD17 

To deflect VD17, a change in velocity of 1 cm/s must be imparted to the asteroid roughly 

15 years or more before the 2102 impact date, requiring a change in momentum of 2.6 x 

109

 kg m/s. Table 32 shows a summary of the performance parameters for this scenario. 

Table 32. Asteroid VD17 - Deflection Performance of Alternatives 

Performance 

Index (P) 

Launches 

Required 

Launch Vehicle Î

Concept 

Delta 

IV H 

Ares 

V 

Delta 

IV H 

Ares 

V 

Nuclear Subsurface1 159 1511 1 1 

Nuclear Surface1 79 755 1 1 

Nuclear Standoff - Neutron1 4.6 22 1 1 

Nuclear Standoff – X-ray1 1.7 7.7 1 1 

Nuclear Standoff - Standard1 1.2 6.2 1 1 

Kinetic Impact, 50 km/s, β=101 0.8 7.1 2 1 

Kinetic Impact, 10 km/s, β=101 0.2 1.4 6 1 

Kinetic Impact, 50 km/s, β=11 0.1 0.7 12 2 

Kinetic Impact, 10 km/s, β=11 0.0 0.1 60 8 

Space Tug – Non-rotating2 0.0 0.7 22 2 

Space Tug - Rotating2 0.0 0.2 68 5 

Gravity Tractor2 0.0 0.0 320 34 

Subsurface Explosive 1 0.0 0.0 855 90 

Surface Explosive 1 0.0 0.0 1709 180 

1

 Assumed to require C3=25 for an intercept trajectory 

2

 C3=25. Rendezvous using electric propulsion likely unrealistic 

* C3=25 may be optimistic by a factor of 2-4 due to launch constraints 
DRAFT Pre-Decisional Material

2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 105 

Preliminary estimates show that if launched optimally, the necessary C3 to intercept 

VD17 may be as low as C3 = 25 km2

/s2. However this is dependent on a very short 

launch window. If launched at a later time, the C3 needed to intercept VD17 likely will 

be between 50 and 100 km2

/s2, reducing deflection performance considerably. 

This example assumes that the launch delivers a C3 of 25 km2

/s2 (a potentially optimistic 

assumption), with a transit time of 2 years. If a Delta IV Heavy is used, a maximum 

payload of 6,000 kg can be sent to this C3 and 25,000 kg can be sent to a low-Earth orbit. 

The Ares V launch vehicle is assumed to lift 38,000 kg to a C3 of 25 km2

/s2 and 

130,000 kg to low-Earth orbit. 

Kinetic impactors use the dry mass of the vehicle to produce the change in momentum. 

To maximize the mass, the 6,000 kg Delta IV Heavy payload will consist of a 1,700 kg 

mass, and 4000 kg total dry mass will produce the change in momentum. Due the high 

eccentricity of VD17’s orbit, the relative velocity of the impact will be on the order of 20 

km/s. If an Ares V launch vehicle is used, an interceptor dry mass of 28,000 kg is 

possible. Using the same beta factor and relative velocity, the momentum imparted by 

this impact would be on the order of 1.68 x 109

 kg m/s. Two to eight missions would be 

required to impart the required momentum depending on the value of β. 

Nuclear detonations can impart the most change in momentum for a given payload. The 

Delta IV Heavy can deliver a 1700 kg nuclear explosive payload, and detonation on the 

asteroid’s surface would change the asteroid’s momentum by 1 x 1011 kg m/s — well 

exceeding the requirement. Surface detonation may fragment the asteroid. Consequently, 

a standoff detonation may be required to limit this possibility, reducing performance. 

For this scenario, the options appear more limited than for Apophis. Nuclear and kinetic 

energy impactors can meet mission requirements; in some cases, more than one launch of 

the Ares V may be required. This scenario appears to exceed the capability of most slow 

push techniques and conventional explosives. 

6.13.4. Scenario – Small Asteroid with Satellite

Athos is a hypothetical small asteroid in a low-inclination prograde orbit. [37] It has an 

orbital period of 1.3 years. Its orbit is just elliptical enough to cross Earth’s path. At the 

time of discovery, very little is known about its specific dimensions, rotation, orientation, 

or material properties. It is thought to be an S-type (silicaceous) asteroid about 200 

meters in diameter, with a mass of 1.1 x 1010 kg. A small satellite about one-third the size 

of Athos is part of the system. Table contains a summary of this scenario. 

This class of asteroids represents the lower bound of those considered in this study. As 

Figure 48 shows, short warning times translate into relatively large ΔVs; a spacecraft that 

intercepts the asteroid 2 years before predicted impact would need to impart about 15 

cm/s to reduce the probability of a collision to less than 1 in 1 million. The same vehicle 

intercepting the asteroid 10 years before impact would need to impart only 4 cm/s to 

achieve the same result. 
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2006 Near-Earth Object Survey and Deflection Study 

106 DRAFT Pre-Decisional Material 

Table 33. 200 Meter Class Asteroid Scenario Description 

Scenario 200 m class Asteroid 

Impact Frequency ~5,000 years 

Time to Act 20 years 

Action Begins 10 years prior to impact 

Diameter of Threat 200 m 

Mass of Threat 1.1x1010 kg 

ΔV 4 cm/s 

Δ Momentum 4.4 x 108

 kg m/s 

Unique Features • Moderate warning 

• Small mass 

• Launch constraints 

Figure 48 shows the ΔV required to deflect Athos. Reference [23] provides details of the 

design of a stand-off nuclear deflection mission for Athos and illustrates some of the 

mission complexities. 

0

10

20

30

40

0 500 1000 1500 2000 2500 3000 3500 4000

Days to Impact

ΔV Required to Deflect (cm/sec)

Reduce risk to 1:1,000,000

Reduce risk to 1:1,000

Deflect to miss by 1 Earth radius

Design Point

Figure 48. ΔV Required to Deflect a Hypothetical 200 m Asteroid 

Many asteroids smaller than 180 meters have very short rotational periods - on the order 

of a few hours or minutes, as shown in Appendix Section K.5. Some of these do not have 

a principal axis and are tumbling. Objects of this size present considerable challenges. 
DRAFT Pre-Decisional Material

2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 107 

Figure 49. Launch Opportunities for Optimistic Launch Constraints 

Figure 50. Opportunities for More Realistic Launch Constraints 

For example, Reference [23] showed that the number of launch opportunities available 

depends strongly on the constraints imposed. Figure 49 and Figure 50 illustrate this point 

by relating launch date to time-of-flight for this example, with the shaded areas 

corresponding to feasible launch combinations. The results in Figure 49 constrain the C3 

at Earth to less than 49 km2

/s2 and the equivalent velocity at the PHO to less than 10 

km/s. Figure 50 uses 5 km/s for the velocity constraints and includes a requirement that 

the spacecraft strike the PHO within 15o of its head or tail, assuring that the interceptor is 

directly in front of or behind the object when the explosive detonates. Note the variation 

in the time of flight and the great reduction in launch opportunities as constraints are 

added. 

Since the asteroid is predicted to hit Earth 11 years after detection, a mission must be 

designed and launched in a short time. To reach the preferred trajectories, the launch 

would need to occur either 3 or 6.5 years after the asteroid’s detection. Once launched, 

travel time to Athos is about 210 days. This example assumes that the C3 imparted by the 

launch vehicle will be no larger than 25 km2

/s2. The goal is to impart a change in velocity 

of 5 cm/s and a change in momentum of roughly 5.5 x 108

 kg m/s. 

Figure 51 shows the spacecraft designed for this mission. The spacecraft’s total mass at 

launch is 6,000 kg. The 1,600 kg nuclear explosive used to illustrate this case is a 

specially designed device, with a high neutron yield. The design, based on current 

technology, includes a cruise stage. It is designed to observe the detonation and then 

report back to Earth. 
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2006 Near-Earth Object Survey and Deflection Study 

108 DRAFT Pre-Decisional Material 

Figure 51. Athos Deflection Vehicle 

Table 34 provides a summary of the performance analyses for this concept. 

Table 34. 200 Meter Asteroid - Deflection Performance of Alternatives 

Performance 

Index (P) 

Launches 

Required 

Launch Vehicle Î

Concept 

Delta 

IV H 

Ares 

V 

Delta 

IV H 

Ares 

V 

Nuclear Subsurface1 937 8926 1 1 

Nuclear Surface1 469 4463 1 1 

Nuclear Standoff - Neutron1 27 127 1 1 

Nuclear Standoff – X-ray1 10 45 1 1 

Nuclear Standoff - Standard1 7.3 36 1 1 

Kinetic Impact, 50 km/s, β=101 5.0 42 1 1 

Kinetic Impact, 10 km/s, β=101 1.0 8.4 1 1 

Kinetic Impact, 50 km/s, β=11 0.5 4.2 2 1 

Kinetic Impact, 10 km/s, β=11 0.1 0.8 10 2 

Space Tug – Non-rotating2 0.3 3.9 4 1 

Space Tug - Rotating2 0.1 1.2 12 1 

Gravity Tractor2 0.0 0.2 55 6 

Subsurface Explosive 1 0.0 0.1 145 16 

Surface Explosive 1 0.0 0.0 290 31 

1

 Assumed to require C3=25 for an intercept trajectory 

2

 Assumed to require C3=0 for a rendezvous using electric propulsion 
DRAFT Pre-Decisional Material

2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 109 

If a 1,600 kg standard nuclear explosive detonated at the best possible distance from the 

asteroid’s surface, it would yield a momentum change of 1.5 x 1010 kg m/s, which would 

be enough to successfully deflect this asteroid. Surface and subsurface detonations would 

produce a momentum change of 1.5 x 1012 kg m/s and 3.0 x 1012 kg m/s, respectively. 

These detonations would be much more likely to fragment the asteroid than a standoff 

detonation. 

Reference [23] also raises the issue of reliability. The reference notes that the deflection 

mission demands a higher probability of mission success than does a typical space 

mission. The initial 1-in-100 probability of Athos colliding into Earth will likely need to 

be reduced many orders of magnitude. This issue is considered further in Section 6.12. 

It is likely that scientists will know very little about the asteroid when the mission 

launches; so the beta factor used to calculate the impulse given by a kinetic impactor is 

assumed to be 1.5. With this assumption, a 3,725 kg spacecraft impacting the asteroid’s 

surface at a velocity of 10 km/s will change the momentum by 5.6 x 107

 kg m/s. This 

calculation provides a good lower bound for the impact’s outcome; the beta factor may be 

much higher. After the first launch, a flyby observer will provide more detailed 

information about the asteroid’s condition and tell mission planners whether the mission 

succeeded. 

The asteroid’s companion body or moon complicates the scenario. At some threshold 

value of ΔV, the companion (which could be on an Earth intercept trajectory as well) may 

separate from the primary body. Currently, the effects of such an impact on different-size 

bodies in a very low-gravity field are not well understood. It also is possible that if 

multiple deflection missions were mounted to ensure a sufficient margin and if one of the 

initial attempts successfully deflected the primary body but left the secondary intact, one 

of the trailing missions could be targeted at the companion. 

A slow push technique might offer an advantage in these cases. The rate of impulse that 

the technique imparted would be very small, so small, in fact, that the orbit of the 

companion body would simply move along with the parent. The specifics of such cases 

are beyond the scope of this study. 

Using the largest launch vehicle expected to be available at the time - the Ares V - many 

of the alternatives can address this scenario. 

6.13.5. Scenario – 1 km-Class Asteroid 

Aramis is a hypothetical large asteroid several km in size. [37] A largely C-type 

(carbonaceous) assemblage of solid, gravitationally bound rocks, Aramis is a classic 

“rubble pile.” Aramis has an orbital period of 6.5 years and its orbit is elliptical. It is 1.8 x 

1.2 x 0.8 km and has a mass of 1.2 x 1012 kg. The inherent material of the rubble pile has 

a density of 2.0 gm/cm3

, while the aggregate body has an overall density of about 1.3 

gm/cm3

. Table 35 summarizes this scenario. 

Figure 52 shows the ΔV required to deflect Aramis as a function of time. If action is 

taken 15 or more years before it encounters Earth, the minimum ΔV required may be less 

than 1 cm/s. As time to impact decreases, the ΔV increases markedly. 
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2006 Near-Earth Object Survey and Deflection Study 

110 DRAFT Pre-Decisional Material 

Table 35. 1-km Class Asteroid Scenario Description 

Scenario 1-km Class Asteroid 

Time to Act 17 years 

Impact Frequency ~1,000,000 years 

Action Begins 6 years prior to impact 

Diameter of Threat 1000 m 

Mass of Threat 1.2x1012 kg 

ΔV 2 cm/s 

Δ Momentum 2.4 x 1010 kg m/s 

Unique Features • Rubble pile 

• Large mass 

• Short warning 

In this hypothetical case, Aramis was detected in February of 2006 and is predicted to 

strike Earth in May 2033. This scenario offers time to advance required technologies and 

to send precursor missions to better characterize the asteroid. This scenario also allows a 

more detailed look at how different missions react to a “rubble pile” asteroid. 

Aramis has an elliptical orbit, with a period of 6.5 years. As Figure 52 shows, a ΔV of 2 

cm/s will deflect the threatening asteroid 6 years ahead of impact. To obtain the 

necessary change in velocity, the deflection mission must impart a change in momentum 

of 2.4 x 1010 kg m/s by 2027. 

Aramis requires a C3 of more than 50 km2

/s2 to intercept the asteroid - a relatively large 

launch energy requirement. The Ares V is expected to be ready by 2020 allowing 7 years 

to fulfill the mission needs. If launched at the optimum time with a C3 of 60 km2

/s2, the 

transit time to Aramis is about 5 years. 

 

0

5

10

15

20

0 2000 4000 6000 8000 10000

Days to impact

ΔV Required to Deflect (cm/sec)

Reduce risk to 1:1,000,000

Reduce risk to 1:1,000

Deflect to miss by 1 Earth radius

Design Point

Figure 52. Velocity to Deflect a Hypothetical 1-km Asteroid 
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2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 111 

Table 36 summarizes the performance of the deflection alternatives for the 1-km asteroid 

Aramis which approaches Earth from a highly elliptical orbit. 

Table 36. 1-km Asteroid - Deflection Performance of Alternatives 

Performance 

Index (P) 

Launches 

Required 

Launch Vehicle Î

Concept 

Delta 

IV H 

Ares 

V 

Delta 

IV H 

Ares 

V 

Nuclear Subsurface1 8.1 92 1 1 

Nuclear Surface1 4.1 46 1 1 

Nuclear Standoff - Neutron1 0.3 1.3 3 1 

Nuclear Standoff – X-ray1 0.1 0.7 10 2 

Nuclear Standoff - Standard1 0.1 0.5 15 2 

Kinetic Impact, 50 km/s, β=101 0.1 0.4 20 3 

Kinetic Impact, 10 km/s, β=101 0.0 0.1 100 12 

Kinetic Impact, 50 km/s, β=11 0.0 0.0 200 23 

Kinetic Impact, 10 km/s, β=11 0.0 0.0 1000 115 

Space Tug – Non-rotating*,2 0.0 0.0 497 24 

Space Tug – Rotating*,2 0.0 0.0 1561 75 

Gravity Tractor2 0.0 0.0 6009 492 

Subsurface Explosive 1 0.0 0.0 16667 1478 

Surface Explosive 1 0.0 0.0 33334 2956 

1

 Assumed to require C3=60 for an intercept trajectory 

2

 Assumed to require C3=25 for a rendezvous using electric propulsion 

* Space Tug may have significant issues attaching to a rubble pile of any size 

If a kinetic impact mission is launched, the entire dry mass of the vehicle will be used to 

change the asteroid’s momentum. Because the asteroid is a “rubble pile,” one can assume 

that the beta factor for the impact is 1.2. Additional information on the asteroid’s surface 

strength would be needed to ensure that the interceptor did not punch through the 

asteroid. For this analysis, it was assumed that the interceptor would stay imbedded in the 

asteroid; otherwise, performance would be reduced. 

For their size, nuclear explosives release a relatively large amount of energy. Since the 

asteroid is a rubble pile, imparting this large amount of energy on or below the surface 

may fragment the asteroid, see Section 6.14. 

The Spaceguard Survey continues to reduce the unwarned threat of a 1-km asteroid. If a 

credible 1-km threat were discovered, this scenario shows that multiple launches of the 

most energetic alternatives would be needed to apply the required impulse. Lowerperforming kinetic impact options, the slow push techniques, and conventional explosives 

are probably insufficient to address this hypothetical case. 

6.13.6. Scenario – 1-km diameter Long Period Comet 

Long-period comets represent a difficult mitigation challenge. Likely, they will be 

detected only months, not years, before impact. Furthermore, accurately determining their 
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2006 Near-Earth Object Survey and Deflection Study 

112 DRAFT Pre-Decisional Material 

orbits may be difficult due to out-gassing as they approach the Sun and their higher 

relative velocity. Finally, these objects can be very large, with diameters of several 

kilometers or more. The impact velocity would be several times that of an asteroid (55-70 

km/s compared with 15-20 km/s for an asteroid), meaning that the energy delivered at 

impact would likely be an event with global consequences. Table 37 contains a summary 

of this scenario. 

Table 37. 1-km Comet Scenario Description 

Scenario 1 km Comet 

Time to Act < 2 years 

Impact Frequency >> 1,000,000 years 

Action Begins 1 year prior to impact 

Diameter of Threat 1 km 

Mass of Threat 1x1012 kg 

ΔV 500 cm/s 

Δ Momentum 5 x 1012 kg m/s 

Unique Features • Very short warning 

• Moderate mass 

• Cometary orbit 

Recent calculations have estimated comets to make up less than 1% of the total threat; 

however, due to their size, they offer an even more extreme example of a low-probability, 

high-consequence event. [6] Until they are activated by interaction with the solar wind, 

comets are many magnitudes fainter than comparable asteroids. If they are to be 

discovered before reaching Jupiter’s distance from the Sun, a survey might be required to 

see objects seven magnitudes (more than a factor of 10) fainter than analyzed for the 

Survey in this study. Currently, astronomers discover them only “a few months to a year” 

before passing by Earth. [38] 

To discover incoming km-scale comets before they reach Jupiter’s distance, about thirty 

10-meter aperture telescopes on Earth or about fifty to one hundred 2.5-meter spherically 

distributed telescopes would be required at Jupiter’s distance. Although such a program 

does not appear practical, the study assumes that it could provide up to 30 months of 

warning. [39] This is very little time to mount a deflection campaign. 

The hypothetical object Porthos [37] is a long-period comet in a high-inclination orbit. It 

is detected on February 22, 2013 and will strike on October 18, 2015 if not deflected. A 

classical “dirty snowball,” it rotates once every 4 days. Porthos is roughly one-eighth the 

size of Halley’s nucleus in each axis. Porthos is 2x1x1 km and has a mass of 1x1012 kg (a 

very small comet). 

Because of the short time between discovery and predicted impact, scientists will know 

little about its physical characteristics and potential impact point on Earth. The comet’s 

non-gravitational forces will add to these uncertainties, and any program to address this 

scenario must take such uncertainties into account. The probability of a real impact may 

be 1 in 1,000 or less due to the comet’s high relative velocity, orbit, and relatively short 

observational interval. 
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2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 113 

Figure 53 shows the ΔV that might be required for such an object. Note that the required 

ΔV close to impact is orders of magnitude higher than for an object detected years before 

impact. Should a deflection effort be launched within a year (400 days) of the possible 

impact, a ΔV of more than 5.0 m/s to lower the impact probability to less than 1 in 1 

million would be required. [34] Also note that due to the uncertainty in the orbit, a miss 

distance of many Earth radii may be required. 

0

500

1000

1500

2000

2500

3000

0 200 400 600 800 1000

Days to impact

ΔV Required to Deflect (cm/sec)

Reduce risk to 1:1,000,000

Reduce risk to 1:1,000

Deflect to miss by 1 Earth radius

Design Point

Figure 53. ΔV to Deflect a Hypothetical 1 km Long Period Comet 

A comet as small as Porthos (~1-km radius) appears to be rare. If the observed frequency 

of short-period (SP), dynamically new (DN), and long-period (LP) comets shown in 

Figure 54 are typical [40], Porthos would be among the smallest. 

Figure 54. Frequency of Observed Comets by Size 

While comets are thought to vary in mass from 1010 to 1018 kg, most are expected to fall 

in the range of 1011-1014 kg. [41] At 1012 kg, Porthos is among the least massive comets 

by several orders of magnitude. However, the likelihood of a possible comet impact is 

highly unlikely, even compared with the threat posed by asteroids. 
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2006 Near-Earth Object Survey and Deflection Study 

114 DRAFT Pre-Decisional Material 

Table 38 shows analysis for the performance of the deflection alternatives for a 1-km 

long-period comet such as Porthos. 

Table 38. Long Period Comet - Deflection Performance of Alternatives 

Performance 

Index (P) 

Launches 

Required 

Launch Vehicle Î

Concept 

Delta 

IV H 

Ares 

V 

Delta 

IV H 

Ares 

V 

Nuclear Subsurface1 0.0 0.4 26 3 

Nuclear Surface1 0.0 0.2 52 5 

Nuclear Standoff - Neutron1 0.0 0.0 625 157 

Nuclear Standoff – X-ray1 0.0 0.0 2084 313 

Nuclear Standoff - Standard1 0.0 0.0 3125 417 

Kinetic Impact, 50 km/s, β=101 0.0 0.0 4167 479 

Kinetic Impact, 10 km/s, β=101 0.0 0.0 20834 2393 

Kinetic Impact, 50 km/s, β=11 0.0 0.0 41667 4785 

Kinetic Impact, 10 km/s, β=11 0.0 0.0 208334 23921 

Space Tug – Non-rotating2 0.0 0.0 103506 4935 

Space Tug - Rotating2 0.0 0.0 325171 15503 

Gravity Tractor2 0.0 0.0 1251833 102375 

Subsurface Explosive 1 0.0 0.0 1643008 307837 

Surface Explosive 1 0.0 0.0 3285152 615673 

1

 All assumed to require C3=60 for an intercept trajectory, which is very low 

2

 Rendezvous to this orbit is likely infeasible. Assumed C3=25, which is very low. 

In addition to deflection issues related to the mass of most comets, their high-energy 

trajectories pose a different, yet still difficult problem. Figure 55 shows the C3 required 

to arrive at Porthos and then to deflect it. The calculations include the time of flight 

between launch and arrival. This launch energy is in addition to the ΔV required to 

prevent an impact. 

Figure 53 shows the relatively large ΔV that is required, while Figure 55 shows that a 

timely interception is even more problematic. Each curve represents a different warning 

time, with the X-axis values showing the corresponding time of flight, assuming a 

mission can be launched. The required values of C3 are well beyond the capability of 

current or expected launch vehicles (0-100 km2

/s2 for current and planned launch 

vehicles). The assumed launch C3 of 60 km2

/s2 (for intercept trajectories) is, therefore, 

considered unrealistically optimistic. However, any larger values of C3 would make the 

results in Table 38 that much less meaningful. 

It can be concluded from Figure 55 that very early identification of a long-period comet 

threat is necessary to have any chance to successfully deflect it. 
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2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 115 

0

20000

40000

60000

80000

100000

0 50 100 150 200 250 300

Time of Flight (days)

C3 (km /s )

270 days before impact

210 days before impact

150 days before impact

90 days before impact

30 days before impact

2 2

Comet is detected ...

Figure 55. C3 required to Reach Porthos as a Function of Time of Flight. 

6.14. Fragmentation 

A mitigation approach not specifically analyzed in this report is fragmentation. If the 

threatening object can be broken into smaller pieces (for example, less than 10 meters in 

diameter), then these pieces will disintegrate in the Earth’s atmosphere if they are not 

sufficiently diverted to miss Earth entirely. 

The primary challenge with fragmentation is how to design the breakup so that most of 

the resulting fragments are sufficiently small. This may require extensive knowledge of 

the object’s material properties, even more than what is required for many deflection 

approaches. A second challenge, also potentially requiring a higher level of knowledge, is 

whether the fragments will be destroyed after passing through the atmosphere. While this 

is likely the case for non-metallic objects, it probably is not the case for iron fragments. 

The literature suggest that the fragmentation approach is best suited for cases where the 

lead time to impact is very long, on the order of decades, and where the PHO is not very 

large (< 1 km) [67]. The modest size and long lead time provide margin for any slowmoving debris to travel away from the initial intercept path and pose less of a risk of 

Earth contact. 

Any of the impulsive deflection approaches designed to alter the trajectory of an object 

also may fragment it. An impulsive delivery of acceleration conveys energy in a shock 

wave. This energy contributes to the resulting change in the object’s kinetic energy. It 

also is the source of other changes, including heating, rotational kinetic energy (if the 

action line is not through the center of mass), spalling, as well as possible fracturing and 

dispersion of the fragments to escape velocity. 
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2006 Near-Earth Object Survey and Deflection Study 

116 DRAFT Pre-Decisional Material 

In the process of reducing the probability of Earth impact, the production of fragments 

may not be deleterious. For the deflection of an intact object versus one that has been 

fragmented, the center of mass of the system would undergo the same change in velocity. 

If the velocity imparted to the system is adequate in magnitude and direction, the 

fragmented body will achieve a miss trajectory just as an intact object would. [34] [67] 

A subtle consequence of deliberate or unintended fragmentation is the fact that a second 

attempt might have to act on and interact with a debris cloud. Finally, while 

fragmentation is a possible mitigation strategy, it is not included in the congressional 

direction to “divert” threats and requires analyses beyond the scope of this study. 

6.15. Deflection Systems Schedule Analysis 

Figure 56 shows the estimated development schedules for the deflection alternatives 

considered. The estimates should be considered “rough order of magnitude.” 

A deflection campaign may consist of much more than just the development schedule, as 

shown in Table 18. In addition to developing one or more unique systems and launch 

vehicles to meet mission reliability requirements, in-situ characterization missions are 

required in some cases. To some extent, these missions may proceed in parallel with the 

development of the deflection systems. In addition, launch constraints may make it 

difficult to target many PHOs. Combined with transit time, these constraints could add 

several years to the interception time. Finally, decisions regarding deflection could be 

complex and controversial, particularly if nuclear materials or explosives are used. These 

factors must be considered in the overall schedule. 

9

9

6

9

9

10

10

0 1 2 3 4 5 6 7 8 9 10 11 12

Conventional

Standoff Explosive

Conventional

Surface Explosive

Kinetic Energy

Impactor

Nuclear Standoff

Explosive

Nuclear Surface

Explosive

Nuclear Electric

Space Tug

Nuclear Electric

Gravity Tractor

Development Schedule (Years)

Figure 56. Deflection System Estimated Development Schedules 
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2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 117 

6.16. Deflection System Cost Analysis 

Figure 57 displays the life-cycle cost (LCC) through 2030 for the deflection concepts 

while Figure 58 displays the recurring costs of subsequent missions. Appendix M 

contains more detailed cost estimates. Note that the nuclear systems do not include the 

cost of developing nuclear explosives for PHO deflection, and additional study as well as 

a specific deployment strategy and/or scenario will be required to understand the total 

costs of the alternatives. Costs are presented as a rough order of magnitude (ROM). 

$2.5

$2.5

$1.0

$1.0

$2.5

$2.0

$10.0

$10.0

$0.0 $2.0 $4.0 $6.0 $8.0 $10.0 $12.0

Conventional Subsurface

Conventional Surface - Delayed

Kinetic Impactor (beta=1, DV=10km/s)

Kinetic Impactor (beta=10, DV=50km/s)

Nuclear Standoff (Standard)

Nuclear Surface

Nuclear Electric Space Tug - Rotating

Nuclear Electric Gravity Tractor

ROM Development + Demonstration Cost ($B FY06)

Notes 

- Delta IV H launch costs

- Does not include cost of 

 nuclear explosives

Figure 57. Deflection Concepts Development and First Mission LCC 

$1.0

$1.0

$0.5

$0.5

$1.0

$1.0

$4.0

$4.5

$0.0 $1.0 $2.0 $3.0 $4.0 $5.0

Conventional Subsurface

Conventional Surface - Delayed

Kinetic Impactor (beta=1, DV=10km/s)

Kinetic Impactor (beta=10, DV=50km/s)

Nuclear Standoff (Standard)

Nuclear Surface

Nuclear Electric Space Tug - Rotating

Nuclear Electric Gravity Tractor

ROM Recurring Cost per Deflection Attempt ($B FY06)

Notes 

- Delta IV H launch costs

- Does not include cost of 

 nuclear explosives

Figure 58. Deflection Mission Recurring Cost 
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2006 Near-Earth Object Survey and Deflection Study 

118 DRAFT Pre-Decisional Material 

6.17. Performance and Effectiveness vs. Cost 

Figure 59 plots the range of performance vs. life-cycle cost for the alternatives. The 

impulsive techniques have a clear cost-performance advantage over the slow push 

alternatives. Nuclear options cost more than other impulsive methods, but provide several 

orders of magnitude more performance. This plot does not consider characterization 

required or the chance of mission success due to unique target characteristics. 

1.0E+05

1.0E+06

1.0E+07

1.0E+08

1.0E+09

1.0E+10

1.0E+11

1.0E+12

1.0E+13

$0 $2 $4 $6 $8 $10 $12

 ROM Development Plus Demonstration Cost ($B 2006)

Performance - Momentum Imparted (kg m/s)

Conventional

Kinetic Explosives

Impactor

V=10km/s

β=1

Nuclear

Standoff 

Nuclear

Surface

Nuclear Electric

 Space Tug

Nuclear Electric 

Gravity Tractor

Delta IV H C3=60

Ares V C3=0

Kinetic

Impactor

V=50km/s

β=10

Figure 59. Deflection Performance vs. ROM Development Cost 

6.18. Other Factors 

Other factors are discussed in the following appendices. 

• Appendix P. Potential Synergies of Deflection with Exploration 

• Appendix Q. Potential Synergies of Deflection with the DoD 

• Appendix R. Role of Nuclear Explosives Technology 

6.19. Findings 

• Nuclear standoff explosives are an effective deflection option for many threat 

scenarios. They minimize the possibility of fracturing the target asteroid and 

require only basic information on the target for mission design. Other techniques 

involving nuclear explosives may be more effective, but may be less reliable. 

• Kinetic impactors are the most mature approach and could be used in some 

scenarios, especially for threats consisting of a single, small, solid mass. 

• Slow push techniques are the most expensive, and their ability to divert an object 

is very limited unless one assumes very long action times. 
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2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 119 

• For deflection requiring physical interaction with the surface, the diversity in 

NEO surface properties, material types, and internal configurations must be 

considered. These attributes will make many possible deflection scenarios a very 

complex operation compared with almost any historical space mission. 

• The time to impact is an important determining factor in deflection decisions. As 

the survey finds potential (but uncertain) threats over the next half century, 

questions of “short warning” (5, 10, or 20 year threats) may become germane. 

After the population is fully catalogued (25 or 50 years in the future), most orbit 

impact predictions will become more certain 100 years into the future. 

• Launch vehicles and interplanetary spacecraft fail at relatively high rates (2-5% 

for launches, 10%+ for spacecraft). Deflection options may not perform as desired 

and unexpected failures can happen. Planners must consider these issues when 

designing a deflection campaign. It is likely that several spacecraft, several launch 

vehicles, several launch sites, etc. will be required to ensure mission success. 

• Direct launch options may be limited. Each potential threat will be unique and the 

launch windows available and the launch vehicles required will depend strongly 

on the relative orbits of the PHO and the Earth. In many cases (as high as 70%), 

there may be insufficient launch capability to reach the PHO at a point where it 

can be deflected. Swingby trajectories or assembly of modular propulsion systems 

may be the only way to reach many threats. 

• Survey systems that can warn of a comet threat 2 years into the future are likely 

technologically infeasible. In addition, these objects are probably more massive 

than hazardous asteroids and too distant for current or planned launch systems to 

deflect. It may be easier to fragment comets than asteroids, but this complex 

analysis is beyond the scope of this study. 
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2006 Near-Earth Object Survey and Deflection Study 

120 DRAFT Pre-Decisional Material 

7 Integrated Analyses 

7.1. Warnings 

7.1.1. The Torino Scale

Figure 60 is a representation of the Torino Scale [42] [43], a method for categorizing the 

impact hazard associated with PHOs. It is intended as a tool for astronomers and the 

public to assess the seriousness of collision predictions. The scale combines probability 

statistics and known kinetic damage potentials into a single threat value. The Palermo 

Technical Impact Hazard Scale [44] is a similar, but more complex, and therefore not 

generally used for communicating with the public. 

Figure 60. The Torino Impact Hazard Scale 

In 1999, asteroid 1999 AN10 was the first to be rated at a non-zero value on the Torino 

Scale (TS). It was rated TS 1 due to a calculated probability of about 1:500,000 that the 

asteroid would collide in 2044, about the same as the average annual “background” risk 

for an untracked 1-km PHO. Subsequent tracking of the object later narrowed the range 

of possible future orbits and eliminated the TS 1 rating. Since the advent of automated 

impact-monitoring systems, such as the NASA/JPL Sentry and the University of Pisa’s 

CLOMON2 in 2002, all PHO orbits are now systematically checked for potential future 

collisions. TS 1-level predictions have become fairly common, occurring at the rate of 

about five per year. In total, almost 20 non-zero Torino Scale warnings have been issued 

since the beginning of 2002. 

The highest Torino Scale rating to date occurred for the 2029 close approach of asteroid 

Apophis, one of the objects used as an example in this report. In late 2004, Apophis 

became the first object to reach TS 2, and for a few days, it reached TS 4. The rating was 
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2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 121 

downgraded to zero when a marginal detection was eventually located within precovery 

(Section 5.10.2) data. Apophis is now expected to pass close to the Earth in 2029, with no 

probability of impact. 

Prior to Apophis, no NEO ever had been given a Torino value higher than TS 1. For a 

short while in February 2006, asteroid 2004 VD17’s predicted 2102 encounter with Earth 

was rated TS 2, making it the second asteroid to exceed TS 1. The encounter was 

subsequently downgraded to TS 1, and, as of August 2006, it remains the only object 

with a TS 1 rating. 

7.1.2. Warnings Generated by the Survey

In Reference [24], a policy framework for the PHO hazard is developed, and in particular, 

the potential socio-economic reactions to the “cosmic impact hazard” are considered. The 

reaction of society to an increased number of warnings generated by the next-generation 

NEO Survey is one of the problems examined. The approximate number of warnings 

expected over the course of the SDT survey can be predicted by extrapolating the 

collision predictions already detected by the NASA/JPL Sentry system over the course of 

its first 4.5 years in operation. [45] The analysis shows that the SDT survey can be 

expected to increase the number of TS 1 warnings by more than an order of magnitude, to 

nearly 1,000. TS 2, 3, and 4 warnings will similarly increase to about 15, 50, and 5, 

respectively. There are fewer TS 2 warnings than TS 3 because TS 2 warnings are for 

larger objects, which are less common than the small objects that might be rated TS 3 or 

TS 4. This extrapolation also predicts a 1% chance of a TS 6 warning and a 0.7% chance 

of a TS 7 warning during the survey. 

Using these results, and assuming that a survey detects 90% of the objects 140 meters or 

larger by the end of 2020, the expected warnings can be broken down further by the year 

during which the warning is likely to be issued. Figure 61 shows a summary of expected 

warnings per year and by Torino Scale. This figure shows that when the PanSTARRS 4 

(PS4) survey system is deployed by the USAF, it will generate Torino Scale warnings at 

a rate several orders of magnitude greater than today in each class, independent of a 

NASA-sponsored survey. 

These results do not indicate any increased risk of PHO impact. There is about a 2% 

chance of an actual impact of a PHO larger than 140 meters in the next 100 years. The 

survey program likely would predict such an event in advance. However, the transient 

reaction of society to 98% of the other warnings, which will eventually be withdrawn, is 

a matter of speculation and is possibly worth more consideration by public policy 

analysts. This topic and the potential economic costs are discussed further in 

Reference.[24] 
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2006 Near-Earth Object Survey and Deflection Study 

122 DRAFT Pre-Decisional Material 

Torino Scale = 1 : Warnings/year - PS4 + LSST + 0.5m IR@L1

0

20

40

60

80

100

120

140

160

180

2010 2012 2014 2016 2018 2020

Number of Torino Scale 1 Warnings

Torino Scale = 2-4 : Warnings/Year

0

2

4

6

8

10

12

2010 2012 2014 2016 2018 2020

Number of Warnings

TS=2

TS=3

TS=4

Survey Assets

PS4 in 2010

0.5m IR @ L1 in 2012

LSST in 2014

~900 total Warnings

Figure 61. Representative Warnings for a Survey Completed by 2020 

7.1.3. The Role, Effect, and Effectiveness of Warning Systems

The purpose of an NEO surveillance system is to generate warnings that enable a 

mitigation response. On the other hand, there may be social costs associated with 

warnings that are later withdrawn. These costs are determined by the frequency of “false” 

warnings, the warning horizon (time to possible impact), how long a warning remains in 

effect, and, least certain of all, how society reacts to them. 

Reference [24] also points out that warnings have a social component, in addition to a 

technical one. The public receives many warnings on many topics in many forms, and the 

format is often as important as the message. This is particularly true if recipients are 

unable to personally respond to the threat. 

7.2. Framework for Action Thresholds 

7.2.1. Phases of Discovery

The distribution of warning times for new discoveries and the fact that two or more 

decade’s worth of observations (sometimes much less) will yield very accurate 

predictions for more than 50 years (see Section 5.9) into the future provide a framework 

for discussing action thresholds. This is used as an overlay to probability of impact. The 

search and deflection problem for any newly discovered threat can be seen as one of 

several time-phased possibilities that depend on time of observation and predicted time to 

impact. This is particularly true if surveys and cataloguing are assumed to continue 

beyond 90% completion. 
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2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 123 

A possible framework for a newly discovered object is outlined in Table 39. 

Table 39. Phases after an Object is Discovered 

Phase Discussion of Characteristics 

Phase 1 

1-4 days after 

detection of an 

object

• Propagation of the initial orbit has 10-100’s of Earth 

radii (ER) of error per year 

• Except for very short warning impacts, all warnings 

carry a low calculated impact probability and then are 

only good for 1-10 years in the future 

Phase 2 

1-4 months after 

detection of an 

object

• During this time precovery is used and additional 

observations are sought 

• Impacts 1-10 years are << 1% but the calculated 

probability of impact for some rises for a time and then 

the risk is eliminated quickly in almost all cases 

• Impacts 10-40 years in the future are < 1%. The 

probability of impact rises for some and then is 

eliminated for most 

• Warnings for impacts 40-100 years can be made but are 

of very low probability 

Phase 3 

1-4 years after 

detection, about 

1 orbital period 

for most objects

• Precovery has been used to the maximum extent 

possible 

• There is a 25% chance of radar observing the object 

during the first 4 years, rapidly improving the orbit by a 

factor of 1000 

• 1-10 year warning threats have either passed, or have 

been fully eliminated or identified as threats 

• 10-40 year warned threats are well characterized with 

few “probabilistic threats” remaining (threats with a 

probability of impact) 

• Additional 30-100 year warning threats are added, and 

many more are eliminated 
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7.2.2. The Two Epochs of the NEO Problem

Taking an even broader view of these phases, two “epochs” of the NEO hazard become 

apparent. During the first epoch, the system is in flux. This period of “discovery and orbit 

refinement” lasts about 20-40 years as new objects are discovered and their orbits are 

refined. Short-duration (1-20 years) impact warnings are possible, but represent a small 

fraction of the total warnings. Changes in warning classifications are made, warnings are 

added, and then most are eliminated. 

In the second epoch, the detection system is phased out and tracking remains the system’s 

only function. More than 99% of objects 140 meters and larger (and many smaller 

objects) have been detected. Their orbits that can be accurately propagated, assuming 

tracking capabilities continue to be supported. Average warning times increase from 

decades to hundreds of years. Short-duration warnings are limited to the threat of comets, 

which are largely outside our current capabilities to detect with warnings exceeding 1-2 

years. Comet threats are expected to be less than 1% of the already low-probability total 

threat and are outside the scope of this study. To gain statistically relevant benefits from 

the Survey, modest tracking systems must be maintained for centuries as a consequence 

of monitoring a 5,000-year average hazard. 

It is likely that in conjunction with or following the proposed survey, many systems will 

be brought on line for scientific or other purposes. It is expected that these will identify 

and track threats below the level of potential hazard within one or two generations. To 

detect a high percentage of objects to this limit, the detection systems likely will need to 

be space-based unless unexpected advances in ground-based telescope technologies are 

made (see Appendix Section N.1). Some of the space-based IR alternatives evaluated for 

the survey can detect 90% of objects down to 80 meters. 

Details of the epochs are described further in Table 40. 

Table 40. Phases after the Survey is Complete 

Epoch Discussion of Characteristics 

Epoch A 

20-40 years 

after survey is 

initiated 

• Radar has provided a precision orbit for 75% of warned 

threats, but most objects have enough observations to 

make radar less useful 

• No threats have been known for less than 10 years 

• Newly identified threats are uncommon 

• If any objects have high impact probabilities, this has been 

known for more than 10 years 

• 10-40 year warned threats are either eliminated or their 

probability of impact is very well understood 

• Few 30-100+ year “probabilistic” threats remaining 

Epoch B 

40+ years after 

survey is 

complete 

• A century of very accurate propagation is possible 

• Many centuries of warning are likely 

• New threats are very uncommon 

• Precision tracking by radar is almost completely obviated 
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7.2.3. Amortized Cost of Unnecessary Threat Response Missions

Figure 62 shows the amortized (without inflation) cost of unnecessary characterization 

and deflection campaigns. The figure determines the cost over a range of rates of 

characterization and deflection per actual impact. The lowest cost alternative assumes one 

of each type of mission per impact, while the highest cost alternative assumes 1,000 

characterizations and 100 deflection campaigns per impact. This translates to one 

characterization mission every 5 years (999 unnecessary missions) and one deflection 

mission every 50 years (99 unnecessary missions) for objects 140 meters and larger. This 

figure shows that even very high rates of unnecessary missions have relatively lowamortized costs; therefore, investing in assets to reduce the rate of unnecessary missions 

is likely not warranted. This figure also suggests that investment in deflection prior to an 

actual threat may be unwarranted unless protecting against “short warning” scenarios. 

0

50

100

150

200

250

1-1 10-1 100-1 100-10 1000-10 1000-100

Characterization-Deflection Campaings per Potential Impact 

Amortized Yearly Cost (FY06$M)

Assumes

- Characterization missions cost $100M

- Deflection campaigns cost $10B

- Impacts occur on average 5000 years

- All costs amortized, actual mission

 occurrences will be $100M/$10B spikes

If threats to 50m are included, the rate 

rises to 1 every 500 years (or so) and the 

costs shown increase 10x

e.g., 100 characterizations and

10 deflections per potential impact

Figure 62. Amortized Cost of Unnecessary Threat Response Missions 
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126 DRAFT Pre-Decisional Material 

7.3. Survey and Deflection Strategy Options 

Figure 63 shows a decision tree for selected strategy options for both parts of this study. 

In Step 1, a detection and tracking architecture is selected. Ground-based, space-based, 

and combinations of ground- and space-based systems can meet the congressional goal of 

detecting 90% of PHOs 140 meters and larger by 2020. In Step 2, it is expected that 

search systems can provide a portion of the orbit and mass characterization required to 

assess the threat, but options for enhancing this capability including radar are presented. 

In Step 3, mitigation strategy options (including perhaps deflection demonstrations) are 

chosen informing the choice of further characterization in Step 4 (if required). 

Infrared in Venus-like orbit + Baseline 97% $1.1B

Infrared @ L1 + Baseline 91% $1.1B

Infrared in Venus-like orbit 89% $0.84B

Dedicated ground-based 85% $0.66B

Dedicated ground-based + Baseline 90% $0.82B

Shared ground-based (Baseline)* 83% $0.31B

Option FY06 $ through 2020 LCC thru 2020*

Infrared in Venus-like orbit + Baseline 97% $1.1B

Infrared @ L1 + Baseline 91% $1.1B

Infrared in Venus-like orbit 89% $0.84B

Dedicated ground-based 85% $0.66B

Dedicated ground-based + Baseline 90% $0.82B

Shared ground-based (Baseline)* 83% $0.31B

Option FY06 $ through 2020 LCC thru 2020*

2018 $1.0B

2020 $1.1B

2021 $0.86B

2020 $0.82B

2024 $0.87B

2026 $0.52B

Year 90% LCC to 90%*

2018 $1.0B

2020 $1.1B

2021 $0.86B

2020 $0.82B

2024 $0.87B

2026 $0.52B

Year 90% LCC to 90%*

1. Select Detection and Tracking Architecture other similar options exist

~$150M cataloguing costs in all options 1-2% uncertainty in relative performance

2. Add Orbit and Mass Characterization 

Dedicated space $1-$1.5B

Radar for some threats < $100M

Dedicated ground $100-$300M

Augment search systems $100M-$400M

Use search systems $0B

Options FY06 $ LCC thru 2020*

Dedicated space $1-$1.5B

Radar for some threats < $100M

Dedicated ground $100-$300M

Augment search systems $100M-$400M

Use search systems $0B

Options FY06 $ LCC thru 2020*

3. Develop a Mitigation Strategy

Cooperate with international efforts $50M+

Develop nuclear explosives technologies ?

Perform deflection demonstration (by 2014) $100M+

Coordinate deflection technologies w/ ESMD $0+

No dedicated preparation $0+

Options FY06 $ LCC thru 2020

Cooperate with international efforts $50M+

Develop nuclear explosives technologies ?

Perform deflection demonstration (by 2014) $100M+

Coordinate deflection technologies w/ ESMD $0+

No dedicated preparation $0+

Options FY06 $ LCC thru 2020

4. Develop Characterization Further if Needed for the Mitigation Strategy ($0-$8B)

 * LCC = Life-cycle cost 

Figure 63. Decision Tree for Survey and Deflection Strategy Options 
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8 Summary 

In the 2005 Budget Authorization Act, the U.S. Congress directed the NASA 

Administrator to provide an analysis of alternatives to detect, track, catalogue, and 

characterize potentially hazardous near-Earth objects (NEO). Congress required that the 

Administrator submit a program by December 28, 2006 to survey 90% of the potentially 

hazardous objects measuring at least 140 meters in diameter by the end of 2020. In 

addition, the legislation required the Administrator to submit an analysis of alternatives 

that NASA could employ to divert an object on a likely collision course with Earth. 

A study team, led by the Office of Program Analysis and Evaluation (PA&E), derived 

requirements and figures of merit from the Act, and used these factors to evaluate the 

alternatives. The team developed a range of options from public and private sources and 

then analyzed their capabilities, levels of performance, life-cycle costs, schedules, and 

development and operations risks. This document presents the detailed results of these 

analyses. A summary report was submitted to Congress in December of 2006. 

8.1. Survey Analysis of Alternatives 

Detection and tracking alternatives identified by the study team included optical systems 

located on the ground and optical and infrared assets located in space. For ground-based 

alternatives, the study team considered sharing planned observatories such as 

PanSTARRS 4 (PS4), funded by the U.S. Air Force, and the Large Synoptic Survey 

Telescope (LSST), partially funded by the National Science Foundation. The team also 

considered new NASA-funded facilities that would be dedicated to the search for 

hazardous objects and would be based on these planned observatories. Although cost 

margin was applied to alternatives that leveraged planned assets, programs that rely on 

these projects may carry additional cost and schedule risk. Specific results include: 

• An architecture, which combines the sharing of the planned PS4 and LSST 

systems with a second, dedicated NASA-funded LSST, was the only groundbased alternative able to meet the congressional goal of identifying 90% of the 

hazardous objects by 2020. This combination is estimated to have a life-cycle cost 

of $820M ($FY06). 

• A shared PS4, a shared LSST, and a dedicated NASA-funded PS8 were able to 

catalog 90% of hazardous objects by 2024, with a life-cycle cost of $560M. 

• A dedicated, NASA-funded observatory based on LSST’s design was also able to 

catalog 90% of potentially hazardous objects in 2024 without the contributions of 

other programs. Its estimated life-cycle cost is $870M. 

Space-based search alternatives were located in low-Earth orbit, at Sun-Earth Lagrange 

points, and in heliocentric Venus-like orbits. Only an infrared system operating in a 

Venus-like orbit was able meet the congressional goals without the contribution of shared 

ground-based assets. All space-based alternatives were able to meet the goals when 

combined with a ground-based baseline of a shared PS4 and a shared LSST. 
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A space mission failure could delay achieving the 90% goal by up to 5 years, after which 

the catalog could be completed with shared ground-based assets. Infrared systems 

operating in space could provide more accurate size estimates of up to 80% of objects in 

the catalog. Observatories located in a Venus-like orbit are the most efficient at finding 

objects inside Earth’s orbit, a potentially underestimated population. Additionally, by the 

end of 2020, infrared systems in Venus-like orbits can find 90% of the objects measuring 

over about 80 meters, exceeding the 140-meter requirement. Finally, space-based systems 

have much less uncertainty in the date of reaching 90% due to their superior sensitivity. 

Selected space-based alternatives include: 

• A 0.5-meter infrared system operating in a Venus-like heliocentric orbit 

completes 89% of the survey by 2020 which is within the uncertainty of the 

analysis. This system has a life-cycle cost of $840M ($FY06). 

• A similar 0.5-meter infrared system operating in a Venus-like orbit and working 

in concert with a shared PS4 and a shared LSST completes 90% of the survey in 

2018, with a life-cycle cost of $1B through 2018. 

• A 0.5-meter infrared system operating at Sun-Earth L1 in conjunction with the 

baseline finishes 90% of the survey in 2020. Its life-cycle cost is $1.1B. 

Infrared systems with a 1.0-meter aperture complete the survey about 1 year earlier than 

the 0.5-meter alternatives described above, and have life-cycle costs about $300M higher. 

Optical systems with 1.0-meter and 2.0-meter apertures in Venus-like orbits, combined 

with the baseline ground-based systems, completed the survey by 2017 and 2019 

respectively, with life-cycle costs in excess of $1.7B. The visible system with a 2.0-meter 

aperture progressed more slowly than the 1.0-meter system due to differences in 

development time. Acquisition of new systems was assumed to start October 1, 2007, and 

delays in funding will affect the ability of some alternatives to meet the 90% 

completeness goal by the end of 2020. Table 41 and Figure 64 summarize the costperformance variation of a range of detection options. 

Table 41. Summary of Detection, Tracking, and Cataloguing Alternatives 

 

through 

2020 

Cost thru 

2020 

(FY06) 

Year 

90% 

Cost to 

90% 

(FY06) 

Continue Spaceguard (in all) 14% < $0.2B >>2030 - 

Shared PS4 and LSST* 83% $0.31B 2026 $0.52B 

Shared PS8 + Baseline 85% $0.41B 2025 $0.56B 

Dedicated LSST 85% $0.66B 2024 $0.87B 

Dedicated LSST + Baseline 90% $0.82B 2020 $0.87B 

0.5m IR @ L1 + Baseline 91% $1.1B 2020 $1.1B 

1.0m IR @ L1 + Baseline 91% $1.3B 2019 $1.3B 

0.5m IR @ Venus + Baseline 97% $1.1B 2018 $1.0B 

1.0m IR @ Venus + Baseline 97% $1.4B 2017 $1.3B 

1.0m VIS @ LEO/L1 + Baseline 93% $1.8B 2017 $1.7B 

2.0m VIS @ LEO/L1 + Baseline 95% $2.1B 2019 $2.0B 

 * Baseline = Shared PS4 + Shared LSST 
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2006 Near-Earth Object Survey and Deflection Study 

130 DRAFT Pre-Decisional Material 

2015

2020

2025

2030

2035

$0.0 $0.5 $1.0 $1.5 $2.0 $2.5

 Lifecycle Cost ($B 2006)

Year of 90% PHO Completeness

Baseline = Shared PS4 + Shared LSST

Baseline + 

New Dedicated 

Ground Based

Baseline + Dedicated LSST

Baseline + Shared 

Ground Based

1 m

0.5 m

2 m

Baseline +

Space Visible

Goal

Baseline +

Space IR 

2015

2020

2025

2030

2035

$0.0 $0.5 $1.0 $1.5 $2.0 $2.5

 Lifecycle Cost ($B 2006)

Year of 90% PHO Completeness

Baseline = Shared PS4 + Shared LSST

Baseline + 

New Dedicated 

Ground Based

Baseline + Dedicated LSST

Baseline + Shared 

Ground Based

1 m

0.5 m

2 m

Baseline +

Space Visible

Goal

Baseline +

Space IR 

Existing Systems Never Reach 90% Goal

Figure 64. Year of Reaching 90% Completeness vs. Life-Cycle Cost 

Congress provided two objectives for characterizing potentially hazardous objects. The 

first objective, to “assess the threat,” requires analysts to determine the orbit and 

approximate the mass of each hazard. Detection and tracking systems with judicious 

follow-up are all able to provide warning, and some are able to provide very good size 

and mass estimates. Systems operating in the visible spectrum are limited by a factor of 

two for size estimates, resulting in a factor-of-eight uncertainty in mass. Infrared systems 

provide data for much more accurate size estimates. 

If detection systems must characterize the catalog, the time to complete the survey to a 

90% completion level will be extended. Furthermore, the costs of these systems may 

increase $100M-$400M to accommodate filters and additional data processing. In 

addition, the smallest and faintest objects may remain visible to sensors only for a few 

days or weeks. Therefore, if characterization is required and it is not performed by 

detection systems, either formal relationships with extant observatories for “on demand” 

access must be negotiated or new dedicated characterization facilities will be needed. 

Radar may quickly and precisely characterize and determine the orbit of about 10-25% of 

the objects of interest within 5 years of their detection. While the number of objects 

observed by radar increases with time, the relative value of radar to precisely determine 

the orbits of the full catalog declines over the same period. Orbits determined from 

optical data alone will nearly match the accuracy of radar-improved orbits after decades 

of observation. Therefore, the utility of radar is limited to a relatively few “short 

warning” cases that may be of very high interest during the survey. Up to $100M in 

funding (not included in detection and tracking life-cycle costs) may be required to 

maintain radar capability through 2020, as NASA and National Science Foundation 

funding for existing radars is currently in flux. 
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The second objective of characterization is to “inform mitigation.” Depending on the 

mitigation strategy selected, this objective may require information beyond the size and 

orbit of potential threats. This information may include the structure, porosity, rotation 

rate, material composition, and surface features of the threats. The deflection alternatives 

considered are sensitive to the maximum mass that needs to be deflected, but some 

alternatives are orders of magnitude less sensitive than others. 

Characterization by remote sensing provides some information about the diversity of 

objects in the population. From this information, analysts build models that can be used 

to infer a limited number of characteristics of a particular object. Only in-situ encounters 

can provide the definitive observations necessary to calibrate the remote observations. 

More importantly, only in-situ visits can obtain the information needed by some of the 

deflection alternatives to mitigate a specific threat. For credible threats with sufficient 

warning, it is expected that in-situ characterization will always be performed to both 

confirm the probability of impact (with a transponder, for example) and to characterize 

the potential threat if deflection is necessary. 

This study has determined that it is premature to set specific characterization 

requirements to enable mitigation until a mitigation strategy has been determined; 

therefore, the study has developed characterization options that provide a range of 

capabilities. These options included the use of detection and tracking assets, dedicated 

ground and space systems for remote observation, and in-situ missions to inform 

mitigation of threats with sufficiently high impact probabilities. These options have lifecycle costs ranging from $50M-$8B ($FY06) over several decades, and are summarized 

in Table 42. 

Table 42. Summary of Characterization Capability Options 

 Description

Cost (FY06) 

Period

Option 1 Use existing assets plus detection and tracking 

system elements. No dedicated characterization.

$0.1B 

2007-2026 

Option 2 Develop dedicated ground system(s) to gather and 

analyze data. No in-situ missions.

$0.2B 

2007-2026 

Option 3 Develop dedicated space system(s) to gather and 

analyze data at Sun-Earth L1. No in-situ missions.

$1.1B 

2007-2023 

Option 4 Develop dedicated space system(s) to gather and 

analyze data near Venus. No in-situ missions.

$1.2B 

2007-2023 

Option 5 Add to Options 1-3 eight (8) visits to representative 

NEOs using fly-bys to calibrate models. 

$2.0B 

2007-2024 

Option 6 Add to Options 1-3 eight (8) visits to representative 

NEOs using rendezvous missions to test models. 

$6.7B 

2007-2031 

Option 7 Perform Option 6 at a fixed mission rate to 8 

potential threats at 5 years intervals 

$8.2B 

2007-2053 
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132 DRAFT Pre-Decisional Material 

It is expected that during the 5-10 years of a survey, a total of 500,000 objects will be 

discovered by more than 2 million individual observations. About 21,000 of these objects 

will measure 140 meters or larger and be tracked as potentially hazardous. Although this 

study uses an estimate of the population of potentially hazardous objects based on 

statistical projections, the actual number of objects will not affect the date of reaching the 

90% goal as long as the objects are approximately distributed in orbits as predicted. 

This volume of observations will require a data-processing capability that is 100 times 

more capable than current cataloguing systems. After objects are detected, the system 

must be able to obtain follow-up observations, store and distribute collected data, and 

analyze these data for observed but previously undetected objects. Currently, 

uncompensated or under-funded analysts perform many of these functions. Such an 

approach likely will not remain viable. Finally, either the NASA Survey or otherwise 

funded activities, such as PS4 and LSST, are expected to produce impact warnings at a 

rate that is 40 times greater than what is experienced today. This much higher rate of 

warnings will start as early as 2010. 

8.2. Deflection Analysis of Alternatives 

The study considered a wide range of techniques to divert a threatening object. These 

alternatives were broadly classified as “impulsive” if they acted nearly instantaneously, 

or “slow push” if they acted over an extended period of time. Launch, orbit transfer, 

technology development, and object characterization requirements were developed for 

each of these alternatives. They were applied to a set of five scenarios representing the 

likely range of threats over million-year timescales. 

The use of nuclear explosives was found to be the most effective alternative in the near 

term. While an explosion on or below the surface of a threatening object is 10-100 times 

more effective than a detonation above the surface, the standoff detonation would be less 

likely to fragment the target. Nuclear options require the least amount of detailed 

information about the threatening object. A nuclear standoff mission could be designed 

knowing only the orbit and approximate mass of the threat, and most impulsive missions 

could be carried out incrementally to reach the required amount of deflection. Additional 

information about the object’s mass and physical properties would perhaps increase the 

effectiveness, but likely would not be required to accomplish the goal. The study 

examined conventional explosives, but found they were ineffective against most threats. 

Kinetic impact alternatives are the most effective non-nuclear option, transferring 10-100 

times less momentum than nuclear options for a fixed launch mass. Impact velocities, 

varying from 10-50 km/s, produced a factor-of-three variation in deflection performance. 

In addition, kinetic impacts also are sensitive to the porosity, elasticity, and composition 

of the target and may require larger performance margins if these characteristics are not 

well determined. 

Slow push techniques analyzed in this study included a gravity tractor, which would alter 

the course of an object using the gravitational attraction of a massive spacecraft, and a 

space tug, which would attach to an object and move it using high-efficiency propulsion 

systems. An attached space tug has generally 10-100 times more performance than the 

gravity tractor, but it requires more detailed characterization data and more robust 
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guidance and control and surface attachment technologies. Slow push techniques were 

determined to be useful in relatively rare cases (fewer than 1% of expected threat 

scenarios), but these techniques could be effective in instances where small increments of 

velocity (less than 1 mm/s) could be applied to relatively small objects (less than 200 

meters in diameter) over many decades. 

Figure 65 shows the span of deflection performance and costs for the alternatives 

analyzed. Deflection performance is expressed as momentum imparted to the target and 

covers the range of likely threats over a million-year timescales. 

1.0E+05

1.0E+06

1.0E+07

1.0E+08

1.0E+09

1.0E+10

1.0E+11

1.0E+12

1.0E+13

$0 $2 $4 $6 $8 $10 $12

 ROM Development Plus Demonstration Cost ($B 2006)

Performance - Momentum Imparted (kg m/s)

Conventional

Kinetic Explosives

Impactor

V=10km/s

β=1

Nuclear

Standoff 

Nuclear

Surface

Nuclear Electric

 Space Tug

Nuclear Electric 

Gravity Tractor

Delta IV H C3=60

Ares V C3=0

Kinetic

Impactor

V=50km/s

β=10

Figure 65. Deflection Performance vs. ROM Development Cost 

The level of risk reduction required of a deflection campaign needs to be clearly 

understood, as it has a first-order impact on cost and complexity. While this report uses a 

goal of reducing the probability of impact to 1 in 1 million, this is not a nationally or 

internationally accepted threshold. Additionally, when designing the deflection campaign, 

planners must take into account that launch vehicles and interplanetary spacecraft fail at 

relatively high rates (2-5% for launches; 10+% for spacecraft) and that deflection 

approaches may not perform as designed. Planning for many flights of multiple 

spacecraft designs launched from several different launch vehicles may be necessary to 

achieve the reduction in impact probability projected to be required. 
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2006 Near-Earth Object Survey and Deflection Study 

134 DRAFT Pre-Decisional Material 

8.3. Summary of Findings 

• Combining optical ground-based observatories currently under development with 

a dedicated ground-based asset can reach the congressional goal by the end of 

2020. Life-cycle cost for this architecture, including a robust data-management 

and data-analysis infrastructure, is estimated to be $820M through 2020. 

• Space-based infrared systems, combined with shared ground-based assets, could 

reduce the overall time to reach the 90% goal by up to 3 years, with life-cycle 

costs of $1.0-$1.3B through 90% completion. Space systems have additional 

benefits and risks over ground-based alternatives, and are generally more capable 

(sensitive) than ground based alternatives. 

• Radar systems cannot contribute to the search for potentially hazardous objects, 

but may be used to rapidly refine tracking and to determine object sizes for a few 

objects of potentially high interest. Existing radar systems are oversubscribed, and 

funding to operate these systems may be in flux. A budget for radar is not 

included in the detection and tracking life-cycle costs. 

• Determining an object’s mass and orbit are required to determine whether it 

represents a threat and to inform deflection alternatives. Beyond these parameters, 

characterization requirements and capabilities are tied directly to the mitigation 

strategy selected. Life-cycle costs for the characterization options vary by billions 

of dollars depending on the mitigation strategy pursued. 

• While several countries have capable programs to study near-Earth objects, none 

of these efforts has materially influenced the results of the study team. 

• Nuclear standoff explosions are assessed to be 10-100 times more effective than 

the non-nuclear alternatives analyzed in this study. Other techniques involving 

nuclear explosives may be more effective, but they run an increased risk of 

fracturing the target. They also carry higher development and operations risks. 

• Kinetic impactors are the most mature approach and could be used in some 

scenarios, especially for objects that consist of a relatively small, solid body. 

• Slow push deflection techniques are the most expensive, and their ability to both 

travel to and divert a threatening object is limited unless mission durations of 

many decades are available. 

• Deflection campaigns may need to be 100-1,000 times more reliable than current 

space missions to meet mitigation requirements. 

• Many potentially hazardous objects (30-80%) are in orbits that are beyond the 

capability of current or planned launch systems. Therefore, if these objects need 

to be deflected, swingby trajectories or on-orbit assembly of modular propulsion 

systems may be required to augment launch vehicle performance. 
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9 Definition of Terms 

Acronym Description 

ACE Advanced Composition Explorer 

APL Applied Physics Laboratory 

AU Astronomical Unit 

D Diameter 

DCT Discovery Channel Telescope 

DoD Department of Defense 

FOM Figure of Merit 

FY Fiscal Year 

HQ Headquarters 

IEO Interior Earth Object 

IOC Initial Operational Capability 

IR Infrared 

IRTF InfraRed Telescope Facility 

Isp Specific Impulse 

JHU Johns Hopkins University 

JPL Jet Propulsion Laboratory 

JSC Johnson Space Center 

JWST James Webb Space Telescope 

KBO Kuiper Belt Objects 

L1 First Sun-Earth Lagrange Point 

LCC Life-Cycle Cost 

LEO Low-Earth Orbit 

LINEAR Lincoln Near Earth Asteroid Research 

LONEOS Lowell Observatory Near-Earth-Object Search 

LSST Large Synoptic Survey Telescope 

MOID Minimal Orbital Intersection Distance 

MPC Minor Planet Center 

NASA National Aeronautics and Space Administration 

NEAT Near-Earth Asteroid Tracking 

NEO Near-Earth Object 

NEOO Near-Earth Object Office 

NVO National Virtual Observatory 

PA&E Office of Program Analysis and Evaluation 

Pan STARRS Panoramic Survey Telescope & Rapid Response System 

PHO Potentially Hazardous Object 

PS Pan STARRS 

RFI Request for Information 

SDT Science Definition Team 

SIM Space Interferometry Mission 
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Acronym Description 

SMD Science Mission Directorate 

SOHO Solar and Heliospheric Observer 

SST Space Surveillance Telescope 

TS Torino Scale 

Vis Visible 

VISTA Visible and Infrared Survey Telescope 

WISE Wide-field Infrared Survey Explorer 

WMAP Wilkinson Microwave Anisotropy Probe 
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10 References 

[1] Public Law 109-155. http://frwebgate.access.gpo.gov/cgibin/getdoc.cgi?dbname=109_cong_public_laws&docid=f:publ155.109

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Objects. White Paper. NASA NEO Workshop, Vail, Colorado. June 2006. 

[3] Study to Determine the Feasibility of Extending the Search for Near-Earth 

Objects to Smaller Limiting Diameters. Report of the Near-Earth Object 

Science Definition Team. Prepared at the Request of NASA Office of Space 

Science, Solar System Exploration Division. August 22, 2003. 

[4] Bottke, W.F., Morbidelli, A., Jedicke, R., Petit, J.M., Levison, H.F., Michel, P., 

Metcalfe, T. S. Debiased Orbital and Absolute Magnitude Distribution of the 

Near-Earth Objects. 2002. Icarus, 156, 399. 

[5] Michel, P., Zappala, V., Cellino, A., Tanga, P. Estimated abundance of Atens 

and asteroids evolving on orbits between Earth and Sun. ICARUS, 2000, 

Vol.143, No.2, pp.421-424 

[6] The Spaceguard Survey: Report of the NASA International NEO Detection 

Workshop, Jet Propulsion Laboratory, Pasadena. Morrison, D., ed. 1992. 

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11 Appendices 

Appendix A. George E. Brown Jr. Near-Earth Object Survey Act ...................... 149

Appendix B. Study Organization and Process ........................................................ 151

B.1. Organizational Structure ............................................................................. 151 

Appendix C. Derived Requirements ........................................................................ 153

Appendix D. Public Workshop ................................................................................. 155

Appendix E. Descriptions of Survey Alternatives................................................... 158

E.1. Detailed Descriptions of Survey Alternatives............................................. 158 

E.1.1. Spaceguard [D-1]................................................................................ 158 

E.1.2. Shared LSST [D-2] ............................................................................. 159 

E.1.3. Shared Pan-STARRS 4 (PS4) [D-3]................................................... 160 

E.1.4. Dedicated LSST [D-4] ........................................................................ 161 

E.1.5. Dedicated Pan-STARRS 4 (PS4) [D-5].............................................. 163 

E.1.6. Dedicated Pan-STARRS 8 (PS8) [D-6].............................................. 164 

E.1.7. Dedicated Pan-STARRS 16 (PS16) [D-7].......................................... 164 

E.1.8. 2-meter Visible (Vis) LEO/L1/L2 [D-8]............................................. 165 

E.1.9. 1-meter Visible (Vis) Venus-like [D-9].............................................. 166 

E.1.10. 2-meter Visible (Vis) Venus-like [D-10]............................................ 167 

E.1.11. 0.5-meter Infrared (IR) L1/L2 [D-11]................................................. 168 

E.1.12. 1-meter Infrared (IR) L1/L2 [D-12].................................................... 170 

E.1.13. 0.5-meter IR Venus-like [D-13].......................................................... 171 

E.1.14. 1-meter Infrared (IR) Venus-like [D-14] ............................................ 172 

E.2. Precision Orbit Determination Alternatives................................................ 173 

E.2.1. Radar................................................................................................... 173 

E.3. Data Management System Alternatives...................................................... 174 

E.3.1. Scale Existing Data Management Systems [D-15]............................. 174 

E.3.2. Adopt Similar Systems [D-16]............................................................ 175 

E.3.3. New Central Repository [D-17].......................................................... 175 

E.3.4. Backup Facilities [D-18]..................................................................... 176 

Appendix F. Descriptions of Characterization Alternatives.................................. 178

F.1. Object Characterization Alternatives.......................................................... 178 

F.1.1. Remote Characterization – Optical and Near Infrared (to 1um)......... 178 

F.1.2. Remote Characterization – Infrared.................................................... 179 

F.1.3. Remote Characterization – Radar ....................................................... 179 

F.1.4. In-Situ Characterization – Flyby......................................................... 181 

F.1.5. In-Situ Characterization – Rendezvous .............................................. 183 

F.1.6. Ground-Based Characterization Concept Sheets................................ 185 

F.1.7. Space-based Characterization Concept Sheets ................................... 188 

F.1.8. In-Situ Characterization Concept Sheets ............................................ 192 

Appendix G. Description of Deflection Alternatives .............................................. 195

G.1. Deflection Alternatives Using Impulsive Techniques ................................ 195 

G.1.1. Conventional Explosive – Contact...................................................... 195 
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G.1.2. Conventional Explosives - Subsurface ............................................... 197 

G.1.3. Kinetic Impactor ................................................................................. 198 

G.1.4. Nuclear Explosive – Surface Contact ................................................. 201 

G.1.5. Nuclear Explosive — Standoff ........................................................... 203 

G.1.6. Nuclear Explosive - Subsurface.......................................................... 206 

G.2. Deflection Alternatives Utilizing Slow Push.............................................. 207 

G.2.1. Enhanced Yarkovsky .......................................................................... 207 

G.2.2. Focused Solar...................................................................................... 209 

G.2.3. Gravity Tractor.................................................................................... 210 

G.2.4. Mass Driver......................................................................................... 213 

G.2.5. Pulsed Laser........................................................................................ 215 

G.2.6. Space Tug............................................................................................ 216 

Appendix H. Analysis Methodologies ...................................................................... 220

H.1. Analysis of Alternatives Methodology ....................................................... 220 

H.2. Survey Performance Analysis..................................................................... 221 

H.2.1. Performance Simulations.................................................................... 221 

H.2.2. Survey Performance Simulation Verification..................................... 223 

H.3. Deflection Analysis Methodology .............................................................. 223 

H.3.1. The Asteroid Deflection Formula ....................................................... 223 

H.3.2. Effective Momentum Change............................................................. 225 

H.4. Technology Readiness Assessment of Deflection Alternatives.................. 225 

Appendix I. Additional Survey Performance Results................................................ 227

Appendix J.Current Assets and Activities ................................................................. 230

J.1. International Activities................................................................................ 230 

J.1.1. United Nations .................................................................................... 230 

J.1.2. European Space Agency (ESA).......................................................... 230 

J.1.3. Italy – Italian Space Agency (ASI)..................................................... 230 

J.1.4. Japan - Japan’s Aerospace Exploration Agency (JAXA)................... 231 

J.1.5. Germany - German Aerospace Center (DLR) .................................... 231 

J.1.6. Canada................................................................................................. 232 

J.1.7. United Kingdom (UK) ........................................................................ 232 

J.2. United States Activities............................................................................... 232 

J.2.1. NASA.................................................................................................. 232 

J.2.2. Jet Propulsion Laboratory (JPL) NEO Program Office...................... 233 

J.2.3. Harvard/Smithsonian Minor Planet Center......................................... 233 

J.2.4. Spaceguard Survey Assets and Capabilities ....................................... 234 

J.2.5. Current Cataloguing and Data Management Infrastructure................ 235 

Appendix K. Object Characteristics Useful for Mitigation ................................... 239

K.1. Mass of Potentially Hazardous Asteroids................................................... 239 

K.2. Orbits of Hazardous Asteroids.................................................................... 240 

K.2.1. Potentially Hazardous Objects Orbit Periods ..................................... 240 

K.2.2. Orbit Inclination.................................................................................. 240 

K.2.3. Potentially Hazardous Objects Orbit Eccentricities............................ 241 

K.3. Density of Potentially Hazardous Asteroids............................................... 242 

K.4. Taxonomic Classes and Asteroid Types..................................................... 243 

K.5. Porosity and Efficient Transfer of Momentum........................................... 244 
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K.6. Rotation Rates of Near-Earth Asteroids ..................................................... 244 

Appendix L. Launch Capability for Characterization and Mitigation ................ 247

L.1. Launch Capability (C3) of Current and Planned Launch Systems ............. 247 

L.2. Launch Capability Required ....................................................................... 247 

L.3. Launch Energy (C3) for Varied and Fixed Launch or Arrival Dates ......... 248 

L.4. Launch Capability Required to Intercept Comets....................................... 251 

L.5. Launch Capability Summary and Conclusions........................................... 251 

Appendix M. Additional Cost Results...................................................................... 252

M.1. Detection and Tracking Survey Element Costs .......................................... 252 

M.2. Characterization Element Results Costs ..................................................... 253 

M.3. Detailed Deflection Element Cost Results.................................................. 254 

M.4. Characterization Architecture Cost Results................................................ 255 

Appendix N. Roles of Advanced Systems and Technologies.................................. 256

N.1. Detection, Tracking, Cataloguing, Characterization Systems .................... 256 

N.1.1. Advanced Follow-up Systems and Technologies............................... 256 

N.1.2. Advanced Survey Systems and Technologies .................................... 258 

N.2. In-Situ Mitigation........................................................................................ 260 

N.2.1. Technologies....................................................................................... 260 

N.2.2. Suggested Studies ............................................................................... 260 

Appendix O. Potential Synergy of Survey Systems with DoD............................... 262

O.1. Contributions of DoD Systems to PHO Survey.......................................... 262 

O.1.1. System A............................................................................................. 262 

O.1.2. System B ............................................................................................. 262 

O.1.3. System C ............................................................................................. 263 

O.1.4. System D............................................................................................. 263 

O.1.5. System E ............................................................................................. 263 

O.1.6. System F.............................................................................................. 263 

O.1.7. Summary............................................................................................. 263 

O.2. Possible Contributions of NEO Survey Systems to Other Missions .......... 264 

O.3. Sharing of Archived Data ........................................................................... 264 

Appendix P. Potential Synergies of Deflection with Exploration.......................... 265

Appendix Q. Potential Synergies of Deflection with the DoD................................ 268

Q.1. Overview..................................................................................................... 268 

Q.2. Launch Systems .......................................................................................... 268 

Q.3. Deflection Campaign Operations................................................................ 269 

Q.4. Guidance Technology ................................................................................. 269 

Q.5. Industrial Base ............................................................................................ 270 

Q.6. Conclusion .................................................................................................. 271 

Appendix R. Role of Nuclear Explosives Technology ............................................ 272

R.1. Nuclear Explosive Design........................................................................... 272 
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Appendix A. George E. Brown Jr. Near-Earth Object Survey Act 

Public Law No: 109-155. SEC 321 

(a) Short Title.--This section may be cited as the ``George E. Brown, Jr. Near-Earth 

Object Survey Act''. 

(b) Findings.--The Congress makes the following findings: 

(1) Near-Earth objects pose a serious and credible threat to humankind, as many 

scientists believe that a major asteroid or comet was responsible for the mass 

extinction of the majority of the Earth's species, including the dinosaurs, nearly 

65,000,000 years ago. 

(2) Similar objects have struck the Earth or passed through the Earth's atmosphere 

several times in the Earth's history and pose a similar threat in the future. 

(3) Several such near-Earth objects have only been discovered within days of the 

objects' closest approach to Earth, and recent discoveries of such large objects 

indicate that many large near-Earth objects remain undiscovered. 

(4) The efforts taken to date by NASA for detecting and characterizing the hazards of 

near-Earth objects are not sufficient to fully determine the threat posed by such 

objects to cause widespread destruction and loss of life. 

(c) Definitions.--For purposes of this section the term “near-Earth object” means an 

asteroid or comet with a perihelion distance of less than 1.3 Astronomical Units from the 

Sun. 

(d) Near-Earth Object Survey 

(1) Survey program.--The Administrator shall plan, develop, and implement a NearEarth Object Survey program to detect, track, catalogue, and characterize the 

physical characteristics of near-Earth objects equal to or greater than 140 meters 

in diameter to assess the threat of such near-Earth objects to the Earth. It shall be 

the goal of the Survey program to achieve 90 percent completion of it’s nearEarth object catalogue (based on statistically predicted populations of near-Earth 

objects) within 15 years after the date of enactment of this Act. 

(2) Amendments.--Section 102 of the National Aeronautics and Space Act of 1958 

(42 U.S.C. 2451) is amended-- 

(A) by redesignating subsection (g) as subsection (h); 

(B) by inserting after subsection (f) the following new subsection: 

“(g) The Congress declares that the general welfare and security of the 

United States require that the unique competence of the National 

Aeronautics and Space Administration be directed to detecting, tracking, 

cataloguing, and characterizing near-Earth asteroids and comets to provide 
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warning and mitigation of the potential hazard of such near-Earth objects 

to the Earth.”; and 

(C) in subsection (h), as so redesignated by subparagraph (A) of this 

paragraph, by striking ‘and (f)’ and inserting ‘(f), and (g)’. 

(3) Fifth-year report.--The Administrator shall transmit to the Congress, not later than 

February 28 of the fifth year after the date of enactment of this Act, a report that 

provides the following: 

(A) A summary of all activities taken pursuant to paragraph (1) since the date 

of enactment of this Act. 

(B) A summary of expenditures for all activities pursuant to paragraph (1) 

since the date of enactment of this Act. 

(4) Initial report.--The Administrator shall transmit to Congress not later than 1 year 

after the date of enactment of this Act an initial report that provides the following: 

(A)An analysis of possible alternatives that NASA may employ to carry out 

the Survey program, including ground-based and space-based alternatives 

with technical descriptions. 

(B) A recommended option and proposed budget to carry out the Survey 

program pursuant to the recommended option. 

(C) Analysis of possible alternatives that NASA could employ to divert an 

object on a likely collision course with Earth. 
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Appendix B. Study Organization and Process 

B.1. Organizational Structure 

Core Team

Study Lead

Bill Claybaugh

Study Coordinator

Dan Mulville

Study Executive

Marcus Shaw

Characterization 

Working Group

Detection / Tracking 

Working Group

Deflection 

Working Group

Tom Coonce NASA PA&E

Jim Puhek USAF

Rod Liesveld NASA PA&E

Lindley Johnson NASA SMD

Tom Morgan NASA SMD

Minoo Dastoor NASA ESMD

Robert Wegeng NASA ESMD

Don Avery NASA LaRC

Ed Barker NASA JSC

Consultants to Core Team

Bill Ailor Aerospace

Jason Derleth JPL

Rob Gold JHU/APL

Vern Weyers NASA (Ret)

Larry Ross NASA (Ret)

Frank Spurlock NASA (Ret)

Don Yeomans JPL

Barker Ed NASA JSC

Don Yeomans JPL

Scott Stuart MIT/LL

Steve Chesley JPL

Paul Chodas JPL

Tim Spahr SAO CfA MPC

Al Harris SSI

Dan Heimerdinger Consultant

Dave Lawrie Aerospace

Morgan Tom NASA SMD

Rob Gold JHU/APL

Richard Binzel MIT/LL

Andy Rivkin JHU/APL

Larry Lemke NASA ARC

Ed Lu NASA JSC

Tom Jones LM/NASA (ret)

Dave Lynch Aerospace

Alan Tokunaga U of Hawaii

Vern Weyers NASA (Ret)

Larry Ross NASA (Ret)

Frank Spurlock NASA (Ret)

Marty Winkler Retired Ind.

Bill Ailor Aerospace

Robert Wegeng NASA ESMD

Eric Asphaug UC Santa Cruz

Robert Hanrahan NNSA

Analysis Support: Aerospace Corporation, JPL, NASA Langley, MIT / Lincoln Labs

Figure 66. NEO Study Organizational Structure 

Risk Assessment 

Development / Operations

Concept and System 

Development

Synthesis

Schedule / 

Timeline

Development / 

Operations Risk

Affordability Effectiveness

Performance/Utility 

Assessment 

Steering Group 

Core Team + 

Consultants

Cost 

Assessment

Analysis of Alternatives

Synergy with 

Exploration

Policy, International, 

Other Issues

Analysis Team

• JPL

• Aerospace

• NASA Langley

• MIT / Lincoln Labs

Figure 67. Analysis of Alternatives Organizational Structure 
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152 DRAFT Pre-Decisional Material 

April-June

Study Prep

Public Workshop

Research

7/10 8/7 9/4 7/17 7/24 7/31 8/14 8/21 8/28

Core @ HQ

Red Team Midterm

Performance

Designs / Concepts 

of Operations

Mitigation Performance

Model / Tool Development

Concept / Program 

Alternatives Synthesis

c e d

Document

Interactions with Steering Group

Full @ HQ

Red Team

7/10 8/7 9/4 7/17 7/24 7/31 8/14 8/21 8/28

Core @ HQ

Red Team Midterm

Performance

Designs / Concepts 

of Operations

Mitigation Performance

Model / Tool Development

Concept / Program 

Alternatives Synthesis

c e d

Document

Interactions with Steering Group

Full @ HQ

Red Team

Assess Performance 

Assess Performance 

Figure 68. Study Calendar 
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Appendix C. Derived Requirements 

Number Requirement Trace and Notes 

1.0 Detect, track, catalogue and 

characterize (Survey) near-Earth 

asteroids and comets, 

Public Law “to provide warning and 

mitigation of the hazard. 

 1.1 Survey will Survey potentially 

hazardous objects (PHOs), a subset of 

near-Earth asteroids and comets. 

PHOs represent the true hazard and a 

more reasonable number for 2020 

completion. 

 1.2 Track determination shall consist of at 

least 2 nights of observation on the 

same object within one week. 

Sufficient orbit characterization to 

classify as a potential PHO. 

 1.3 Establish a third observation within 40 

days of track determination. 

Sufficient orbit characterization to 

classify and catalog as a PHO. 

 1.4 Additional observations shall be 

performed to quantify, improve, and 

update the actual PHO risk 

Mitigation strategies will require more 

accuracy than initial PHO determination. 

 1.5 The system shall propagate orbits; make 

impact predictions, and predictions of 

impact energy. 

Public Law “to provide warning and 

mitigation of the hazard. 

 1.6 The system shall provide warnings and 

alerts as appropriate to the hazard or 

threat condition. 

Public Law “to provide warning and 

mitigation of the hazard. 

2.0 Plan, develop, and implement a NEO 

Survey program to Survey the 

physical characteristics of near-Earth 

objects equal to or greater than 140 

m. 

Public Law to assess the threat of such 

objects to the Earth. 

 2.1 PHO mass variability shall be 

determined by characterizing a 

statistically valid number of PHOs. 

All mitigation strategies will require 

knowledge of the mass of the PHO. 

 2.2 A range of solutions for characterizing 

other parameters of interest (size, shape, 

rotation rate, composition, etc.) shall be 

cross-checked with mitigation 

alternatives. 

Some mitigation strategies will require 

detailed physical understanding of the 

PHO. 

3.0 Achieve 90% completion of the NEO 

catalogue (based on statistical 

predictions) within 15 years. 

Public Law – 15 years from December 

28, 2005 
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4.0 Provide alternatives to carry out the survey 

program with technical descriptions. 

Public Law 

 4.1 Systems must be capable of searching the entire 

sky down to 25th visible magnitude at least three 

times per lunation. 

System sensitivity consistent with PHO threat 

definition. 

 4.2 Systems must have an astronometric accuracy of 

< 0.5 arc second. 

Observation accuracy needed for sufficient orbit 

determination for PHO cataloguing. 

 4.3 The system shall minimize mean time to 

determine miss distance. 

Provides highest accuracy information for orbit 

and potential keyhole events. 

 4.4 A fault tolerant data collection and preliminary 

orbit determination center must be maintained. 

Architecture requires maintaining a reference and 

warning archive. 

 4.4.1 A geographically distinct "mirror" data center 

shall be maintained. 

Provides geographic diversity to limit impact of 

natural disaster. 

 4.4.2 Offsite data backups shall be maintained. Disaster recovery capability. 

 4.4.3 Capacity shall be for at least 2,000,000 

observations per day for at least 500,000 NEOs, 

and discovery of at least 15 NEOs (3 PHOs) per 

day. 

Matched to threat. 

 4.4.4 Provide web-based access to data for 

“precovery” (analysis of archived data) and other 

data mining activities. 

Leverage archived data and collaboration with 

internationals and amateurs. 

5.0 Provide an analysis of possible alternatives to 

divert an object on a likely collision course 

with Earth. 

Public Law 

 5.1 Reduce PHO impact risk to the nominal PHO 

risk (~1 in one million) 

Reduce PHO impact risk to the nominal 

background risk of PHOs. 

 5.2 Provide a means for verification of object 

diversion. 

Determination whether mitigation action was 

successful. 
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Appendix D. Public Workshop 

On May 15, 2006, the NEO study team solicited public input to help fulfill the 

requirements of this study. The workshop was held June 26-29 in Vail, Colorado and was 

named the Near-Earth Object (NEO) Detection, Characterization and Threat Mitigation 

Workshop. Participation in the workshop was by invitation only. 

The study team was divided into three areas to focus discussion with the public: 

• Detection, tracking and cataloging NEOs 

• Characterization of NEOs, and 

• Deflection or other forms of NEO threat mitigation 

Interested parties were invited to submit a one-page abstract outlining concept proposals 

in the study focus areas by May 26, 2006. The NEO study team divided into working 

groups and these groups during the first week of June to review the abstracts. Invitees 

were chosen based upon their expertise and contribution to understanding the focus areas. 

Authors of accepted abstracts received a workshop invitation by June 7, 2006, and were 

invited to submit a five- to 10-page white paper addressing the topic(s) of inquiry and/or 

detailing the proposed concept, including estimated required schedule and cost 

information. Most authors, except foreigners or those with other specific issues also were 

invited to present their papers at the workshop. The Call for Papers made it clear that 

concept proposals were to be evaluated only for the purposes of assessing alternatives; no 

proposals would be awarded funding as part of the PA&E study process. 

The workshop was held over 4 days, and significantly increased the study team’s 

knowledge of current capabilities and plans, and provided much-needed context for both 

highly specialized and less-experienced team members. Spirited and extremely useful 

dialogue was generated on each of the 4 days, and the public input has significantly 

influenced and contributed to the study’s results. 

Lists of external workshop attendees and their presentation or white paper topics are 

listed on the following pages. 
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Table 43. List of Vail NEO Workshop Attendees and Contributors 

Full Name Organization Title 

Abell, Paul NASA Johnson Space Center 

The Compositions Of Binary Near-Earth 

Objects: Implications For Their Internal 

Structure And Hazard Potential 

Adams, Rob NASA - MSFC 

Summation Of NASA-TP-2004-213089 

"Survey Of Technologies Relevant To Defense 

From Near-Earth Objects (NEO's)" (Paper) 

Arentz, Robert Ball Aerospace and 

Technologies Corp. Invitation Only 

Barbee, Brent Emergent Space 

Technologies, Inc. 

Optimal Deflection Of Hazardous Near-Earth 

Objects By Standoff Nuclear Detonation And 

NEO Mitigation Mission Design 

Bartlett, Paul W Honeybee Robotics 

Implementation Study And Technology 

Development For A Near-Earth Object 

Deflection Mission Using Mass Drivers 

Bekey, Ivan Bekey Designs, Inc. 

Extremely Large Yet Very Low Weight And 

Low Cost Space Based Telescopes For 

Detection Of 140 Meter Diameter Asteroids At 

5.7 AU, And Obtaining 6 Year Warning Times 

For 1 km Diameter Comets 

Bowell, Edward Lowell Observatory 

Searching For Neos Using Lowell 

Observatory's Discovery Channel Telescope 

(DCT) 

Buie, Marc W. Lowell Observatory Physical Characterization Of Phas With 

Ground-Based Telescopes 

Campbell, Donald B Cornell University 

Near-Earth Asteroid Astrometry And 

Characterization By Radar: Current Radar 

Systems And Future Possibilities 

Chapman, Clark R. Southwest Research Institute Mitigation: Interfaces Between NASA, Risk 

Managers, And The Public 

Conway, Bruce A University of Illinois Optimization Of The Deflection Of A 

Hazardous NEO 

Cruikshank, Dale P. NASA Ames Research 

Center 

Determining The Physical Properties Of NearEarth Objects (White Paper) 

Davis, Jeremy J Texas A&M University Impact Keyholes And Collision Probability 

Analysis For Resonant Encounter Asteroids 

Do, Khanh Q The Boeing Company Boeing Space-Based Characterization Sensor 

Concept 

Doyle, Monica SAIC Assessment Of Current Ground-Based 

Telescope Cost Modeling Methodologies 

Dunham, Ed Lowell Observatory Invitation Only 

Fork, Richard L. University of Alabama in 

Huntsville 

Solar Powered Modelocked Lasers For 

Deflection Of Earth Threatening Asteroids 

Gaffey, Michael J University of North Dakota 

The Compositional And Physical 

Characterizations Of Neos From VNIR 

Spectroscopy 

Gertsch, Leslie S. University of Missouri-Rolla Distributed-Energy Blasting For NEO 

Mitigation 

Gorevan, Stephen Honeybee Robotics Enabling Technology For NEO 

Characterization Of Return Samples 

Holsapple, Keith A. University of Washington Existing Methods Of Asteroid Deflections Will 

Work: Impacts And Nuclear Bombs 
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Full Name Organization Title 

Howard, Regan E Orbital Sciences Corp. 

Adapting GEO Spacecraft To Reduce The Cost 

Of NEO Detection, Tracking And Cataloging 

Missions 

Huebner, Walter F. Southwest Research Institute White Paper: Seismology Measurements Of 

NEO Properties For Earth Impact Mitigation 

Izenberg, Noam R. JHU/APL Shield: A Comprehensive Earth Protection 

System 

Junkins, John Texas A&M Invitation Only 

Lambert, John Boeing Invitation Only 

Larson, Stephen M University of Arizona Observational Follow Up Needs For The Next 

Generation Of NEO Surveys 

Lissauer, Jack J. NASA Ames 

Advantages Of Low-Cost Missions To Study 

NEO's: Why Extensive Study Will Tell Us 

More Than Intensive Study 

McMillan, Robert Scott University of Arizona Spacewatch Preparations For The Era Of Deep 

All-Sky Surveys - White Paper 

Mills, Robert Lowell Observatory Invitation Only 

Morrison, David Ames Research Center Near-Earth Asteroid Trailblazer 

Ostro, Steve Jet Propulsion Laboratory 

Ground-based Radar In Near-Earth Object 

Tracking, Characterization And Threat 

Mitigation 

Pitchford, Brian E. NASA/KSC 

Near-Earth Object Deflection System 

(NEODESYS) - Nullifying The Threat Posed 

By Asteroid 99942 Apophis 

Raeth, Peter Ball Aerospace Invitation Only 

Reitsema, Harold J Ball Aerospace Neo-Vis, The Visible-Light, Near-Earth Object, 

Survey Mission. 

Remo, John L Sandia National Laboratories Near-Earth Object Short Term Threat 

Mitigation White Paper 

Ryan, Eileen V. New Mexico Institute of 

Mining and Technology 

Physical Characterization And Follow-Up 

Studies Of Faint Near-Earth Objects Using The 

Magdalena Ridge Observatory's 2.4-Meter 

Telescope 

Scheeres, Daniel J University of Michigan Spacecraft At Small Near-Earth Objects 

Schweickart, Russell B612 Foundation Threat Characterization: Trajectory Dynamics 

Solem, Johndale Los Alamos National 

Laboratory Invited Speaker 

Stuart, Joseph Scott MIT Lincoln Laboratory 

Serendipitous Discovery Of Near-Earth 

Asteroids With The DARPA Space 

Surveillance Telescope 

Tedesco, Edward University of New 

Hampshire 

Albedos And Sizes Of NEOs Using 

Polarimetry 

Tokunaga, Alan T. University of Hawaii A Dedicated Program For The Reconnaissance 

Of NEO Surface Reflectance Characteristics 

Tyson, Anthony University of California LSST: Comprehensive NEO Detection, 

Characterization, And Orbits 

VanCleve Ball Aerospace & 

Technologies Corp. 

NEO Retro-Reflectors And Beacons: 

Concepts, System Trades And Performance 

Issues 

Whiteley, Robert J University of Arizona Rapid Response Characterization Of Small 

Neas 

Wie, Bong Arizona State University Solar Sail Kinetic Energy Impactor 
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Appendix E. Descriptions of Survey Alternatives 

E.1. Detailed Descriptions of Survey Alternatives 

E.1.1. Spaceguard [D-1]

With 10 observatories in seven locations across the world, the Spaceguard Foundation is 

a dedicated international consortium in NEO observations. Telescope apertures range 

from 0.5 meters to 1.5 meters, with a typical field of view of 3° x 3°, which helps achieve 

a typical limiting magnitude (Vlim) of approximately about 19.5. Participation includes 

scientists from Italy, Germany, Australia, United Kingdom, U.S., Japan, New Zealand, 

Russia, Uruguay, Finland, Sweden, Canada, Spain, Czech Republic, China, and Croatia. 

These telescopes, all currently operational, function in the visible-wavelength range. 

Spaceguard includes the following observatories or associations participating in PHO 

detection: LINEAR, Catalina Sky Survey, Pan-STARRS-1, Spacewatch, NEAT, Campo 

Imperatore NEO Survey, U.K. Spaceguard, Japan Spaceguard Association, Spaceguard 

Croatia, Germany Spaceguard Foundation. Two members (LINEAR and Catalina Sky 

Survey), alone, have accounted for approximately 70% of NEO discoveries over the last 

10 years. Although the Spaceguard observatories do not coordinate their search patterns 

for observations, all their detections are reported through the Minor Planet Center. 

Spaceguard telescopes cannot reach limiting magnitudes to detect 140-meter PHOs and 

each organization maintains a separate budget. Funding varies from year-to-year. 

D-1: Spaceguard

Technical Description

• Consortium of ground-based NEO detection efforts

• Visible, 3° x 3° FOV (typical), Vlim= 19.5 (typical)

• Participation includes observatories and scientists from: Italy, Germany, 

Australia, UK, US, Japan, New Zealand, Russia, Uruguay, Finland,

Sweden, Canada, Spain, Czech Republic, China, Croatia

Rationale for consideration

• Currently operational & active in NEO detection & surveying

• Baseline of current NEO detection capabilities

Capability

• Worldwide observation capability that reports findings to the Minor Planet 

Center

Pros

• Often dramatic improvements are possible when pre-discovery observations 

are found (e.g. Apophis 2004 March data extended observed interval 3 months 

and allowed the removal of the 2029 Earth impact possibility).

• Dedicated international collaboration towards this global concern

Cons

• Current telescopes cannot reach limiting magnitudes to detect 140m NEOs

• Daytime, weather, moonlight & location limit sky access

• Each organization maintains a separate budget, funding may vary year-toyear

CONOPS

• Operational: Currently operational

• Worldwide detection observatories: Catalina Sky Survey, Pan-Starrs1, Spacewatch, NEAT, LINEAR, Campo Imperatore NEO Survey, UK 

Spaceguard, Japan Spaceguard Association, Spaceguard Croatia, 

Germany Spaceguard Foundation

Aperture (m) 0.5 to 1.5

Telescopes in System 10+

Locations 7+

Aperture (m) 0.5 to 1.5

Telescopes in System 10+

Locations 7+

System Sizing
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E.1.2. Shared LSST [D-2]

Description

The Shared Large Synoptic Survey Telescope (LSST) is a planned (2014) 8.4-meter 

aperture telescope. It will spend about 75% of its observation time in a survey mode 

compatible with PHO detection. LSST is the largest aperture ground-based optical 

telescope currently in development. Its development is underway and has received 

funding from the Department of Energy (DOE), National Science Foundation (NFS), and 

private donations. 

LSST is a ground-based survey telescope with an alt-azimuth mount. It performs 

observations in the visible and near-infrared range of the spectrum (300-1100 nm), with a 

3.1° x 3.1° field of view and a limiting magnitude (Vlim) of approximately 24.8. The 

optical design consists of a monolithic three-mirror design, with an 8.4-meter diameter 

primary mirror (effective aperture of 6.8 meters) and uses adaptive optics to correct for 

atmospheric distortions. The Shared LSST is to be located at Cerro Pachón, Chile, and is 

expected to be operational by 2014. 

LSST will spend approximately 75% of its time detecting PHOs. It will accomplish this 

by carrying out an automated all-sky survey, with a 35-second cadence and two 15-

second exposures of the same spot in the sky. LSST will take the exposures about 30 

minutes apart. LSST will have the capability to determine an asteroid’s orbital parameters 

and perform visible and near-infrared photometry. New instrumentation can be added to 

expand capabilities. LSST technology is mature and well understood. It is based on 

monolithic, ground-based observatories dating back centuries. 

Qualitative Assessment

Shared LSST offers a low risk of catastrophic failure. As opposed to a space-based asset, 

it can be easily maintained and upgraded. The operation, purpose, and scientific benefits 

of the Shared LSST extend well beyond the proposed survey. It will have fewer 

opportunities to detect PHOs in Earth-like orbits because daylight, weather, moonlight, 

and location limit its access to the sky. LSST may be required to perform follow-up 

observations to determine the orbital parameters of the faintest objects within a limited 

time. 

Requiring the Shared LSST to concentrate on follow-up observations for characterization 

may reduce the time Shared LSST could perform PHO-detection observations. LSST can, 

however, automatically perform follow-on observations while surveying the sky. That is 

because it can cover the entire sky in about 3 nights and therefore would typically 

observe a new PHO about three times within a month. Lastly, it will have a slightly lessthan-optimal survey strategy and hardware since it is not purely a survey telescope. 
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D-2: Shared Large Synoptic Survey Telescope (LSST)

Technical Description

• Ground based survey telescope, alt-azimuth mount

• Visible (300-1100 nm) 5-band, 3.1° x 3.1° FOV, Vlim= 24.8

• Monolithic, 8.4m aperture, 3-mirror design

Rationale for consideration

• Development efforts well underway, NSF & DOE backing

• Largest aperture proposed ground-based optical survey 

•Capability

• NEO orbital parameters, visible & NIR photometry

Pros

• Low risk of catastrophic failure, easy to maintain & upgrade

• Will have lifetime & science value beyond NEO survey

Cons

• If only 1 built, may be needed to provide own follow-up observations

• Infrequent detection opportunities for NEOs in Earth-like orbits 

• Daytime, weather, moonlight & location limit sky access

Aperture (m) 8.4

Telescopes in System 1

Locations 1

Aperture (m) 8.4

Telescopes in System 1

Locations 1

CONOPS

• Location: Cerro Pachon, Chile 

• Operational: 2014

• Availability: 75% in NEO detection mode

• Survey Strategy: All sky survey

• Cadence: 35sec, two 15-sec exposures of same area with 30min separation

System Sizing

E.1.3. Shared Pan-STARRS 4 (PS4) [D-3]

Description

The Shared Panoramic Survey Telescope and Rapid Response System (Pan-STARRS) 4 

(PS4) is an observatory under development consisting of a four 1.8-meter telescopes 

searching the same spot in the sky at the same time, thus combining to provide an 

effective aperture of 3.6 meters. It is expected to spend 30% of its observation time in 

survey mode. PS4 may show that it is more cost effective to build several small 

telescopes than to build a larger, single system. Its development is well underway and has 

received funding from the U.S. Air Force (USAF). The first 1.8-meter telescope (PS1) 

started test operations in June 2006. 

PS4 is a ground-based survey telescope with an alt-azimuth mount. It performs 

observations in the visible and near-infrared range of the spectrum (400-1100 nm), with a 

1.8° x 4.0° field of view and a limiting magnitude (Vlim) of approximately 24.0. Designed 

as a monolith mount, the telescope includes a 1.8-meter diameter primary mirror for each 

of the four telescopes (3.6-meter effective aperture). It uses adaptive optics to correct for 

atmospheric distortions. The Shared PS4 is to be located at Mauna Kea, Hawaii, and is 

expected to be operational by 2010. 

Shared PS4 will spend about 30% of its time detecting PHOs by performing an 

automated all-sky survey, with a 35-second cadence and two 15-second exposures of the 

same spot in the sky roughly 15 to 30 minutes apart. The survey strategy is a balance 

between PHO detection and general astronomy. Sixty-percent of surveying time will be 

utilized in a cadence suitable for linking solar system objects. Additionally, 
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approximately 5% of its surveying time will be devoted to surveying PHO-rich “sweetspots,” the sky within about 10o of the ecliptic and within about 60 ̊ to 90 ̊ from the Sun’s 

line of sight. The system is expected to have at least a 10-year operational lifetime. 

PS4 will have the capability to determine an asteroid’s orbital parameters and perform 

visible and near-infrared photometry. New instrumentation can be added to expand 

capabilities. PS4 technology is mature and well understood and based on monolithic, 

ground-based observatories dating back centuries. 

Qualitative Assessment

As opposed to a space-based asset, Shared PS4 runs a low risk of catastrophic failure and 

can be easily maintained and upgraded. PS4’s lifetime and, therefore, science value will 

extend beyond the completion of the PHO survey. Some of its drawbacks include 

infrequent opportunities to detect PHOs in Earth-like orbits and limited sky access due to 

daylight, weather, moonlight, and a location with limited access to the sky. Lastly, it will 

have a slightly non-optimal survey strategy and hardware since the asset is not purely a 

survey telescope. 

D-3: Shared Pan-STARRS 4 (PS4)

Technical Description

• Four ground based survey telescopes, alt-azimuth mount

• Visible (400-1100 nm), 1.8° x 4.0° FOV, Vlim= 24.0

• Monolithic, 1.8m aperture x 4 telescopes

Rationale for consideration

• Development efforts well underway, backing of USAF

• May be more economically feasible to build coordinated set of small 

telescopes than a single larger one 

•Capability

• NEO orbital parameters, visible & NIR photometry

Pros

• Low risk of catastrophic failure, easy to maintain & upgrade

• Will have lifetime & science value beyond NEO survey

Cons

• Shared, thus non-optimal survey strategy

• Infrequent detection opportunities for NEOs in Earth-like orbits

• Daytime, weather, moonlight & location limit sky access

Aperture (m) 1.8 x 4

Effective Aper. (m) 3.6

Telescopes in System 4

Locations 1

Aperture (m) 1.8 x 4

Effective Aper. (m) 3.6

Telescopes in System 4

Locations 1

CONOPS

• Location: Mauna Kea, Hawaii 

• Operational: 2010

• Availability: 30% in NEO detection mode

• Survey Strategy: All sky survey (when in NEO mode)

• Cadence: 35sec, two 15-sec exposures of same area with 15-30min interval

• All 4 telescopes simultaneously image same spot in sky

• 10 year operational lifetime

System Sizing

E.1.4. Dedicated LSST [D-4]

Description

The Dedicated Large Synoptic Survey Telescope (LSST) is a possible rebuild of the 

Shared LSST, and would be dedicated and optimized for PHO search. A copy of a 

proposed ground-based telescope, dedicated to the PHO survey, would reduce the cost of 

developing a new telescope, while guaranteeing that it spent 100% of its observation time 

in survey mode. The Dedicated LSST will have the same basic design as the Shared 
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LSST, but optimization for PHO detection and an increased exposure time to 22.5 

seconds will increase its limiting magnitude (Vlim) to approximately 25.4. Dedicated 

operation will allow time for the follow up of faint PHOs when necessary. 

The Dedicated LSST was assumed to be located at San Pedro Mártir in Mexico’s Baja 

Peninsula, the Shared LSST’s runner-up location. It would be operational by 2016. The 

Dedicated LSST would spend 100% of its observation time detecting PHOs. The 

Dedicated LSST would have the capability to determine an asteroid’s orbital parameters. 

New instrumentation can be added to expand capabilities. 

Qualitative Assessment

The Dedicated LSST provides redundancy to the Shared LSST. The Shared LSST could 

be leveraged for follow-on observations of PHOs, thus the Dedicated LSST would not 

need to slow its survey to perform follow-up observations. If both the Shared and 

Dedicated LSST image the same PHO, it would provide a parallax view of the PHO, 

allowing for greater accuracy in determining an object’s orbit. Because it is a groundbased asset, it has a lower risk of catastrophic failure as opposed to a space-based asset. It 

can be easily maintained and upgraded. A Dedicated LSST’s lifetime and, therefore, 

science value will extend beyond the completion of the PHO survey. Some of its 

drawbacks include infrequent opportunities to detect PHOs in Earth-like orbits and 

limited sky access due to daylight, weather, moonlight, and location. 

D-4: Dedicated LSST

Technical Description

• Ground based survey telescope, alt-azimuth mount

• Visible (300-1100 nm) 5-band, 3.1° x 3.1° FOV, Vlim= 25.4

• Monolithic, 8.4m aperture, 3-mirror design

• Optimized for NEO detection versus shared asset

• Largest aperture proposed ground-based optical survey 

Rationale for consideration

Aperture (m) 8.4 Capability

Telescopes in System 1

Locations 1

Aperture (m) 8.4

Telescopes in System 1

Locations 1

CONOPS

• Location: San Pedro Mártir, Mexico

• Operational: 2016

• Availability: 100% in NEO detection mode

• Survey Strategy: All sky survey

• Cadence: 50sec, two 22.5-sec exposures of same area with 30min 

separation

• Shared LSST would still be available for follow-on observations

Pros

• Can leverage shared LSST for follow-on observations & co-survey 

• Low risk of catastrophic failure, easy to maintain & upgrade

Cons

• Infrequent detection opportunities for NEOs in Earth-like orbits

• Daytime, weather, moonlight & location limit sky access

System Sizing

• NEO orbital parameters, visible & NIR photometry
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E.1.5. Dedicated Pan-STARRS 4 (PS4) [D-5]

Description

The Dedicated PS4 is a possible rebuild of the Shared PS4 dedicated to PHO search. 

Copying a proposed ground-based telescope dedicated to the PHO survey would save 

money and guarantee that the telescope would spend 100% of its observation time in 

survey mode. The Dedicated PS4 would have the same design as the Shared PS4. The 

Dedicated PS4 is to be located either in Hawaii (Northern Hemisphere) or in Chile 

(Southern Hemisphere). It would be operational by 2012 and will spend 100% of its time 

detecting PHOs. The Dedicated PS4 will be able to determine an asteroid’s orbital 

parameters. New instrumentation can be added to expand capabilities. Dedicated 

operation will allow for the critical follow-up of faint PHOs when necessary. 

Qualitative Assessment

The Dedicated PS4 will provide redundancy to the Shared PS4. If both the Shared and 

Dedicated PS4 in different hemispheres image the same PHO, they would provide a 

parallax view of the PHO, which would enhance observers’ ability to accurately 

determine an object’s orbit. As a ground-based asset, it also has a low risk of catastrophic 

failure, especially compared with a space-based asset; it can be easily maintained and 

upgraded. A Dedicated PS4’s lifetime and, therefore, science value will extend beyond 

the completion of the PHO survey. As with other ground-based observatories, it has 

drawbacks. As with the others, it will have limited opportunities to detect PHOs in Earthlike orbits because of limited access to the sky due to daylight, weather, moonlight, and 

location. 

D-5: Dedicated PS4 (North or South)

Technical Description

• Four ground based survey telescopes, alt-azimuth mount

• Visible (400-1100 nm), 1.8° x 4.0° FOV, Vlim= 24.0

• Monolithic, 1.8m aperture x 4 telescopes

• Optimized for NEO detection versus shared asset

• May be more economically feasible to build coordinated set of small 

telescopes than a single larger one 

Rationale for consideration

Aperture (m) 1.8 x 4 Capability

Effective Aper. (m) 3.6

Telescopes in System 4

Locations 1

Aperture (m) 1.8 x 4

Effective Aper. (m) 3.6

Telescopes in System 4

Locations 1

CONOPS

• Location: Hawaii (North) or Chile (South)

• Operational: 2012

• Availability: 100% in NEO detection mode

• Survey Strategy: All sky survey 

• Cadence: 35sec, two 15-sec exposures of same area with 15-30min interval

• All 4 telescopes simultaneously image same spot in sky

Pros

• Low risk of catastrophic failure, easy to maintain & upgrade

• Will have lifetime & science value beyond NEO survey

Cons

• Infrequent detection opportunities for NEOs in Earth-like orbits 

• Daytime, weather, moonlight & location limit sky access

System Sizing

OR

• NEO orbital parameters, visible & NIR photometry
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E.1.6. Dedicated Pan-STARRS 8 (PS8) [D-6]

Description

The Panoramic Survey Telescope and Rapid Response System (Pan-STARRS) 8 (PS8) is 

a proposed system consisting of two dedicated PS4 observatories, for a total of eight 1.8-

meter telescopes searching the same spot in the sky at the same time, with an effective 

aperture of 5.1 meters. PS8 would be dedicated to PHO search. PS8 may show it is more 

cost effective to build several small telescopes than it is to build a larger, single system. 

PS8 would be composed of two identical observatories. Although its design is based on 

that of the PS4, doubling the number of 1.8-meter telescopes to eight increases its 

limiting magnitude (Vlim) to 24.4. PS8 would be located either in Hawaii (Northern 

Hemisphere) or Chile (Southern Hemisphere) and is expected to be operational by 2012. 

PS8 will spend 100% of its time searching for PHOs. The PS8 would have the capability 

to determine an asteroid’s orbital parameters. New instrumentation could be added to 

expand capabilities. 

Qualitative Assessment

PS8 offers a low risk of catastrophic failure, as opposed to a space-based asset, and can 

be easily maintained and upgraded. PS8’s lifetime and, therefore, science value would 

extend beyond the completion of the PHO survey. Its drawbacks are similar to those of 

other ground-based assets discussed in previous sections. 

Design parameters are similar to those for PS4 except that eight telescopes are used. 

E.1.7. Dedicated Pan-STARRS 16 (PS16) [D-7]

Description

The Dedicated Panoramic Survey Telescope and Rapid Response System (Pan-STARRS) 

16 (PS16) is a proposed system consisting of two PS8 observatories, one located in the 

Northern Hemisphere and the other in the Southern Hemisphere. Each PS8 observatory, 

equipped with a total of eight 1.8-meter telescopes, would search the same spot in the sky 

at the same time, with an effective aperture of 5.1 meters. PS16 would be dedicated to 

PHO search and, by observing a PHO from both hemispheres, would provide a parallax 

view of the asteroid. PS8 may show it is more cost effective to build several small 

telescopes than a larger, single system. 

PS16 would be composed of two identical observatories based on the PS8 design. Their 

geographic separation (Northern and Southern Hemispheres) would allow parallax 

viewing. One of the two observatories would be located in Hawaii (PS16 North) and the 

other in Chile (PS16 South). The full PS16 system could be operational by 2012. PS16 

will spend 100% of its time detecting PHOs. Each PS16 observatory would have the 

capability to independently determine an asteroid’s orbital parameters. Planners expect 

the two observatories to provide greater accuracy in determining orbits. New 

instrumentation can be added to expand capabilities. 

Qualitative Assessment

PS16 offers a lower risk of catastrophic failure, especially compared with a space-based 

asset. PS16 can be easily maintained and upgraded. If both the PS16 North and South 
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DRAFT Pre-Decisional Material 165 

observatories image the same PHO, they would provide a parallax view of the PHO and 

allow for greater accuracy in determining orbits. PS16’s lifetime and, therefore, science 

value will extend beyond the completion of the PHO survey. Its drawbacks include 

infrequent opportunities to detect PHOs in Earth-like orbits and limited sky access due to 

daylight, inclement weather, moonlight, and location. Lastly, the operation of two 

identical facilities in different hemispheres may present operational challenges due to 

shared personnel between the two distant locations, although observatory automation 

may mitigate these issues. 

Design parameters are similar to those for PS4 except that 16 telescopes are used. 

E.1.8. 2-meter Visible (Vis) LEO/L1/L2 [D-8]

Description

The 2-meter Visible LEO/L1/L2 concept is a space-based, survey telescope that would 

scan the sky to search of NEOs. It can be placed into low-Earth Orbit (LEO) or at the 

Sun-Earth L1 or L2 point. This concept has the largest aperture of any space-based, 

optical system considered in this study to increase detection rate. This detection asset will 

search for PHOs and can be used to calculate their orbital parameters. This design traces 

its heritage to other visible, space-based telescopes, and would be similar to the proposed 

design of NEO-VIS. 

The telescope would have a 2.0-meter aperture that detects light in the visible spectrum. 

It would have a field of view of 2.81° x 2.81°, and be able to achieve a limiting 

magnitude (Vlim) of approximately 23.6. A Delta II 2920 would place the spacecraft in a 

low-Earth orbit or an Atlas V 401 would deliver it to a L1/L2 orbit. It would be 

operational around 2014. It is dedicated to PHO detection and the full bandpass is used to 

maximize detection. This spacecraft would use both onboard- and ground-image 

processing. 

Qualitative Assessment

The benefit of a space-based visible sensor is that the asset is not limited to night-only 

observations. Because this asset will be deployed closer to the Earth, as compared with 

Venus-like orbit concepts, it will be less complex and easier to operate. There is 

significant heritage for space-based visible observatories and these telescopes can achieve 

relatively wide fields of view due to mature technology. Thermal requirements for optical 

LEO/L1/L2 concepts also are less strict than they are for IR systems. 

At the same time, space-based systems have a higher risk of catastrophic failure when 

compared with ground assets and are extremely difficult to maintain or upgrade in case of 

failures or unforeseen circumstances. The observatory’s location at LEO/L1/L2 is not 

optimal for PHO detection because it will have fewer opportunities to discover asteroids 

with Earth-like orbits, especially when compared with an observatory in a ~0.7AU, 

Venus-like orbit. Higher data rates are required with optical systems when compared with 

IR systems, especially with a larger aperture. The spacecraft would not be designed to 

store all raw data for precovery. In addition, it only provides limited photometric 

information about PHOs, which makes it difficult to determine an object’s size. 
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D-8: 2m Vis LEO/L1/L2

Technical Description

• Space-based survey telescope (Hubble-like)

• Visible, 2.81° x 2.81° FOV, Vlim= 23.6 

• 2 m aperture (a little smaller than Hubble)

• Earth-based orbit, all of which have virtually identical detection performance

LEO ~ 1000 km orbit

L1 / L2 ~ 1.5x106 km

Rationale for Consideration

• Space-based asset with greater access to sky

• Provide comparison with space-based IR platforms 

Capability

• NEO orbital parameters

Pros

• Can conduct survey at all times (vs. night only on ground)

• Build on heritage spacecraft (Kepler, Hubble)

• Less strict thermal requirements on optics than IR

Cons

• Higher data rates required than IR

• Non-optimal NEO survey location, especially for Earth-like orbits

• Higher risk of catastrophic failure vs. ground based asset

• Very difficult or impossible to maintain or upgrade

CONOPS

• Location: Earth-Sun L1 (nearly identical detection performance at LEO or L2)

• Operational: 2014

• Availability: 100% in NEO detection mode

• Launch Vehicle: Atlas V 401 for L1/L2, Delta II 2920 for LEO

• Full bandpass used to maximize detection 

• On-board image processing coupled with ground image processing

• 10 year mission life

System Sizing

96

102

580

0

225

74

266

1344

1255

Dry Mass (kg)

Thermal 4%

TT&C 4%

Structure & Mechanisms 22%

Propulsion 0%

Power 9%

C&DH 3%

ACDS 10%

Spacecraft Bus 52%

Payload 48%

% of Dry Mass

96

102

580

0

225

74

266

1344

1255

Dry Mass (kg)

Thermal 4%

TT&C 4%

Structure & Mechanisms 22%

Propulsion 0%

Power 9%

C&DH 3%

ACDS 10%

Spacecraft Bus 52%

Payload 48%

% of Dry Mass

Wet Mass (kg) 2599

367

1255

P/L

Dry Mass (kg) 2599

S/C

Power (W) 1311

Wet Mass (kg) 2599

367

1255

P/L

Dry Mass (kg) 2599

S/C

Power (W) 1311

E.1.9. 1-meter Visible (Vis) Venus-like [D-9]

Description

The 1-meter Visible (Vis) Venus-like concept is a space-based, survey telescope that 

would scan the sky in search of NEOs. It is based on the proposed NEO-VIS design and 

would orbit the Sun in a ~0.7 AU heliocentric, Venus-like orbit. This option is included 

to provide a comparison with space-based IR platforms. The selected orbit also offers 

greater visibility of IEOs. This detection asset will search for PHOs and can be used to 

calculate their orbital parameters. This design is based on other visible, space-based 

telescopes, and would be similar to the proposed NEO-VIS design. 

The telescope would have a 1.0-meter aperture that detects light in the visible spectrum. 

It would have a field of view of 2.81° x 2.81°, capable of achieving a limiting magnitude 

(Vlim) of approximately 22. An Atlas V 401 would launch the spacecraft. It would 

become operational around 2014. A passive Venus flyby is employed to position the 

spacecraft in approximately a 0.6 by 0.8 AU orbit. Hydrazine propulsion is used only for 

orbital corrections during the mission’s cruise phase. It is a dedicated PHO detection 

asset that uses the full bandpass to maximize detection. The spacecraft has an operational 

life of 10 years. 

Qualitative Assessment

The benefit of a space-based visible sensor is that more time can be spent on detection; in 

other words, it is not limited to night-only observations. The Venus-like orbit yields a 

greater detection rate and enables the observation of IEOs. A significant heritage exists 

for space-based visible observatories. In addition, larger fields of view can be achieved 
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2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 167 

due to mature technology. Compared with IR systems, thermal requirements also are less 

restrictive. At the same time, space-based systems have a higher risk of catastrophic 

failure and are extremely difficult or impossible to maintain or upgrade in the event that 

they fail. Venus-like orbit-based systems will be larger than Earth-orbiting systems and 

higher data rates are required. 

D-9: 1m Vis Venus-like

Technical Description

• Space-based survey telescope (NEO-VIS like)

• Visible, 2.81° x 2.81° FOV, Vlim= ~22

• 1 m aperture diameter (similar to Kepler Observatory)

• Heliocentric, Venus-like orbit (can be Venus-trailing or other elliptical orbit 

inside of 1 AU)

CONOPS

• Location: Heliocentric, Venus-like orbit

• Operational: 2014

• Availability: 100% in NEO detection mode

• Launch Vehicle: Atlas V 401

• Full bandpass used to maximize detection 

• Mostly autonomous spacecraft due to unique communication profile with 

Earth

• 10 year mission life

Rationale for Consideration

• Provide comparison with space-based IR platforms

• Venus orbit has greater Inner Earth Orbit object visibility

Capability

• NEO orbital parameters

Pros

• Better location to observe NEOs with Earth-like orbits

• Can conduct survey at all times (vs. night only on ground)

• Build on heritage spacecraft (Kepler, Hubble)

• Less strict thermal requirements on optics than IR

Cons

• Higher data rates required than IR

• Higher risk of catastrophic failure vs. ground based asset

* Note: Passive Venus flyby is employed to position S/C in approximately a 0.6 by 0.8 AU orbit. • Very difficult or impossible to maintain or upgrade

Hydrazine used for orbit corrections during cruise-phase of the mission.

Wet Mass (kg) 2070

251

539

P/L

Dry Mass (kg) 1970

S/C

Power (W) 1169

Wet Mass (kg) 2070

251

539

P/L

Dry Mass (kg) 1970

S/C

Power (W) 1169

248

119

580

25

145

86

229

1431

539

Dry Mass (kg)

Thermal 13%

TT&C 6%

Structure & Mechanisms 29%

Propulsion 1%

Power 7%

C&DH 4%

ACDS 12%

Spacecraft Bus 73%

Payload 27%

% of Dry Mass

248

119

580

25

145

86

229

1431

539

Dry Mass (kg)

Thermal 13%

TT&C 6%

Structure & Mechanisms 29%

Propulsion 1%

Power 7%

C&DH 4%

ACDS 12%

Spacecraft Bus 73%

Payload 27%

% of Dry Mass

System Sizing

E.1.10. 2-meter Visible (Vis) Venus-like [D-10]

Description

The 2-meter Visible (Vis) Venus-like concept is a space-based, survey telescope that 

would scan the sky to search for NEOs. It is based on the proposed NEO-VIS design and 

would orbit the Sun in a ~0.7 AU heliocentric, Venus-like orbit. This option is included 

to allow comparison with space-based IR platforms. The selected orbit also offers greater 

visibility of IEOs and its larger aperture increases the detection rate. This detection asset 

will search for PHOs and can be used to calculate their orbital parameters. This design is 

based on other visible, space-based telescopes, and would be similar to the proposed 

NEO-VIS design. 

The telescope would have a 2.0-meter aperture that detects light in the visible spectrum. 

It would have a field of view of 2.81° x 2.81°, capable of achieving a limiting magnitude 

(Vlim) of approximately 23.6. The spacecraft would be launched by a Delta IV Medium+ 

(5,4) and would be operational around 2014. A passive Venus flyby is employed to 

position the spacecraft in a 0.6 by 0.8 AU orbit. Hydrazine propulsion is used only for 

orbital correction during the mission’s cruise phase. It is a dedicated PHO detection asset, 

using the full bandpass to maximize detection. The spacecraft would be primarily 
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168 DRAFT Pre-Decisional Material

autonomous due to a unique communications profile and is designed for 10 years of 

operations. 

Qualitative Assessment

Compared with a ground sensor, a space-based visible sensor can spend more time 

searching for objects because it is not limited to night-only observations. The Venus-like 

orbit yields a greater detection rate and enables the observation of objects primarily 

inside Earth’s orbit. Space-based visible observatories enjoy a long history and can 

achieve larger fields of view due to mature technology. In addition, their thermal 

requirements are less restrictive. 

At the same time, space-based systems have a higher risk of catastrophic failure than do 

ground assets and are extremely difficult or impossible to maintain or upgrade. Venuslike orbit-based systems will be larger than Earth-orbiting systems. Compared with IR 

systems, they also will require higher data rates. The spacecraft would not be able to store 

all raw data for precovery. In addition, it only provides limited astrometric information 

about PHOs, which makes it more difficult to determine an object’s size. 

D-10: 2m Vis Venus-like

Technical Description

• Space-based survey telescope 

• Visible, 2.81° x 2.81° FOV, Vlim= 23.6

• 2m aperture diameter (a little smaller than Hubble Observatory)

• Heliocentric, Venus-like orbit (can be Venus-trailing or other elliptical orbit 

inside of 1 AU)

CONOPS

• Location: Heliocentric, Venus-like orbit

• Operational: 2014

• Availability: 100% in NEO detection mode

• Launch Vehicle: Delta IV Medium+ (5,4)

• Full bandpass used to maximize detection 

• Mostly autonomous spacecraft due to unique communication profile with 

Earth

• 10 year mission life

Rationale for Consideration

• Provide comparison with space-based IR platforms

• Venus orbit has greater Inner Earth Orbit object visibility

Capability

• NEO orbital parameters

Pros

• Better location to observe NEOs with Earth-like orbits

• Can conduct survey at all times (vs. night only on ground)

• Build on heritage spacecraft (Kepler, Hubble)

• Less strict thermal requirements on optics than IR

Cons

• Higher data rates required than IR

• Higher risk of catastrophic failure vs. ground based asset

• Very difficult or impossible to maintain or upgrade

Wet Mass (kg) 3056

367

1255

P/L

Dry Mass (kg) 2903

S/C

Power (W) 1311

Wet Mass (kg) 3056

367

1255

P/L

Dry Mass (kg) 2903

S/C

Power (W) 1311

* Note: Passive Venus flyby is employed to position S/C in approximately a 0.6 by 0.8 AU orbit. 

Hydrazine used for orbit corrections during cruise-phase of the mission.

289

138

648

38

169

100

266

1648

1255

Dry Mass (kg)

Thermal 10%

TT&C 5%

Structure & Mechanisms 22%

Propulsion 1%

Power 6%

C&DH 3%

ACDS 9%

Spacecraft Bus 57%

Payload 43%

% of Dry Mass

289

138

648

38

169

100

266

1648

1255

Dry Mass (kg)

Thermal 10%

TT&C 5%

Structure & Mechanisms 22%

Propulsion 1%

Power 6%

C&DH 3%

ACDS 9%

Spacecraft Bus 57%

Payload 43%

% of Dry Mass

System Sizing

E.1.11. 0.5-meter Infrared (IR) L1/L2 [D-11]

Description

The 0.5-meter Infrared (IR) L1/L2 concept is a space-based, infrared survey telescope 

that would scan the sky in search of NEOs. It is based on the proposed NEOCam design, 

with an orbit at the Sun-Earth L1 point. This option is included due to PHO signatures 

that are stronger in the IR-spectral bandpass; consequently, they would be easier to detect 
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DRAFT Pre-Decisional Material 169 

with this asset. This detection asset will search for PHOs and can be used to calculate 

their orbital parameters. This design is based on other infrared, space-based telescopes, 

and would be similar to the proposed NEOCam design. 

The telescope would have a 0.5-meter aperture that detects light in the infrared region 

between 6-10 μm. It would have a field of view of 1.7° x 6.8°, capable of achieving a 

limiting magnitude (Vlim) of approximately 23.6-25.4. Because the spacecraft would be 

placed in an L1 orbit, the IR detectors would be passively cooled. A Delta II 7920 would 

launch the spacecraft, which is expected to become operational around 2012. It is a 

dedicated PHO detection asset, using the full bandpass to maximize detection. The 

spacecraft has a 10-year operational life. 

Qualitative Assessment

The benefit of a space-based infrared sensor, especially compared with a ground sensor, 

is that atmospheric absorbance would not block incoming IR light. Besides PHO 

signatures are stronger in the IR wavelengths, which results in less source confusion and 

a lower spacecraft data rate. The solar avoidance zone also is smaller for IR detectors. At 

the same time, space-based systems have a higher risk of catastrophic failure and are 

extremely difficult or impossible to maintain or upgrade. The observatory’s location at 

LEO/L1/L2 is not optimal for PHO detection. It will have fewer opportunities to discover 

asteroids with Earth-like orbits, especially when compared with an observatory in a 

~0.7AU, Venus-like heliocentric orbit. Compared with IR systems, the system requires 

higher data rates. In addition, infrared space systems enjoy less historical flight heritage, 

although it will be able to use detector technologies developed for JWST and the Widefield Infrared Survey Explorer (WISE). 

D-11: 0.5m IR L1/L2

Technical Description

• Space-based survey telescope (NEOCam-like)

• Infrared (6-10 μm), 1.7° x 6.8° FOV, Vlim= 23.6-25.4 

• 0.5m aperture

• Passively cooled 

• Earth-based orbit, cannot be LEO due to cooling requirements

L1 / L2 ~ 1.5x106 km

Rationale for Consideration

• NEO signature is stronger in IR domain, easier to detect

• Provide comparison with space-based Vis platforms 

Capability

• NEO orbital parameters

Pros

• No atmospheric absorbance in IR vs. ground-based asset

• Can conduct survey at all times (vs. night only on ground)

• Less source confusion and lower data rate than Vis

• Can look as close as ~55 degrees from Sun

Cons

• Leverages not yet flown systems (JWST, WISE)

• Higher risk of catastrophic failure vs. ground based asset

• Very difficult or impossible to maintain or upgrade

CONOPS

• Location: Earth-Sun L1 (nearly identical detection performance at L2)

• Operational: 2012

• Availability: 100% in NEO detection mode

• Launch Vehicle: Delta II 2920

• Full bandpass used to maximize detection 

• 10 year mission life

System Sizing

23

15

113

0

62

18

109

341

259

Dry Mass (kg)

Thermal 4%

TT&C 2%

Structure & Mechanisms 19%

Propulsion 0%

Power 10%

C&DH 3%

ACDS 18%

Spacecraft Bus 57%

Payload 43%

% of Dry Mass

23

15

113

0

62

18

109

341

259

Dry Mass (kg)

Thermal 4%

TT&C 2%

Structure & Mechanisms 19%

Propulsion 0%

Power 10%

C&DH 3%

ACDS 18%

Spacecraft Bus 57%

Payload 43%

% of Dry Mass

Wet Mass (kg) 600

84

259

P/L

Dry Mass (kg) 600

S/C

Power (W) 302

Wet Mass (kg) 600

84

259

P/L

Dry Mass (kg) 600

S/C

Power (W) 302
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E.1.12. 1-meter Infrared (IR) L1/L2 [D-12]

Description

The 1-meter Infrared (IR) L1/L2 concept is a space-based, infrared survey telescope that 

would scan the sky in search of NEOs. It is based on the proposed NEOCam design, with 

an orbit at the Sun-Earth L1 point. The study team included this option because PHO 

signatures are stronger in the IR-spectral bandpass; so these objects would be easier to 

detect with this asset. It also has a larger aperture that could enhance the PHO detection 

rate. This detection asset will search for PHOs and can be used to calculate their orbital 

parameters. This design has heritage to other infrared, space-based telescopes, and would 

be similar to the proposed NEOCam design. 

The telescope would have a 1.0-meter aperture that detects light in the infrared region 

between 6-10 μm. It would have a field of view of 1.5° x 6.0°, capable of achieving a 

limiting magnitude (Vlim) of approximately 24.2-26.0. The IR detectors are passively 

cooled because of its L1 orbit. A Delta II 2925 would launch the spacecraft, which is 

expected to become operational around 2014. It is a dedicated PHO detection asset, using 

the full bandpass to maximize detection. The spacecraft has a 10-year operational life. 

D-12: 1m IR L1/L2

Technical Description

• Space-based survey telescope 

• Infrared (6-10 μm), 1.5° x 6.0° FOV, Vlim= 24.2-26.0 

• 1m aperture

• Passively cooled

• Earth-based orbit, cannot be LEO due to cooling requirements

L1 / L2 ~ 1.5x106 km

Rationale for Consideration

• Larger aperture IR may enhance detection rate & characterization

• Provide comparison with space-based Vis platforms

Capability

• NEO orbital parameters

Pros

• No atmospheric absorbance in IR vs. ground-based asset

• Can conduct survey at all times (vs. night only on ground)

• Less source confusion and lower data rate than Vis

• Can look as close as ~55 degrees from Sun

Cons

• Leverages not yet flown systems (JWST, WISE)

• Higher risk of catastrophic failure vs. ground based asset

• Very difficult or impossible to maintain or upgrade

• Reduced FOV due to bigger aperture for same telescope length

CONOPS

• Location: Earth-Sun L1 (nearly identical detection performance at L2)

• Operational: 2014

• Availability: 100% in NEO detection mode

• Launch Vehicle: Delta II 2925

• Full bandpass used to maximize detection 

• 10 year mission life

System Sizing

41

27

203

0

111

32

195

609

427

Dry Mass (kg)

Thermal 4%

TT&C 3%

Structure & Mechanisms 20%

Propulsion 0%

Power 11%

C&DH 3%

ACDS 19%

Spacecraft Bus 59%

Payload 41%

% of Dry Mass

41

27

203

0

111

32

195

609

427

Dry Mass (kg)

Thermal 4%

TT&C 3%

Structure & Mechanisms 20%

Propulsion 0%

Power 11%

C&DH 3%

ACDS 19%

Spacecraft Bus 59%

Payload 41%

% of Dry Mass

Wet Mass (kg) 1036

143

427

P/L

Dry Mass (kg) 1036

S/C

Power (W) 449

Wet Mass (kg) 1036

143

427

P/L

Dry Mass (kg) 1036

S/C

Power (W) 449
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DRAFT Pre-Decisional Material 171 

Qualitative Assessment

The benefit of a space-based infrared sensor is that atmospheric absorbance does not 

block the incoming IR light. Besides PHO signatures are stronger in the IR wavelengths, 

resulting in less source confusion and a lower spacecraft data rate. The solar avoidance 

zone is also smaller for IR detectors. On the flip side, however, space-based systems have 

a higher risk of catastrophic failure compared with ground assets and are extremely 

difficult or impossible to maintain or upgrade. In addition, infrared space systems have 

less historical flight heritage, even though they could use detector technologies developed 

for the James Webb Space Telescope and the Wide-field Infrared Survey Explorer 

(WISE). 

E.1.13. 0.5-meter IR Venus-like [D-13]

Description

The 0.5-meter IR Venus-like concept is a space-based, infrared survey telescope that 

would scan the sky searching for NEOs. Its design is based on the proposed NEOCam, 

but it would orbit the Sun in a ~0.7 AU heliocentric, Venus-like orbit. The team included 

this option because PHO signatures are stronger in the IR-spectral bandpass; therefore, 

these objects would be easier to detect with this asset. This concept’s orbit also offers 

greater visibility of IEOs. This detection asset will search for PHOs and can be used to 

calculate their orbital parameters. This design traces its heritage to other infrared, spacebased telescopes, and would be similar to the proposed NEOCam design. 

The telescope would detect light in the infrared region between 6-10 μm. It would have a 

field of view of 1.7° x 6.8° and capable of achieving a limiting magnitude (Vlim) of about 

23.6-25.4. The IR detectors are passively cooled because the spacecraft would be in a 

heliocentric orbit. Operational around 2014, the spacecraft would travel to space by way 

of a Delta II 2926. A passive Venus flyby is employed to position the spacecraft in 

approximately a 0.6 by 0.8 AU orbit. Hydrazine propulsion is used only for orbital 

correction during the mission’s cruise phase. A dedicated PHO-detection asset, it would 

use a full bandpass to maximize detection. The spacecraft has a 10-year operational life. 

Qualitative Assessment

The benefits of a space-based infrared sensor are comparable to those of other IR-based 

systems; namely, no atmospheric absorbance can block incoming IR light. The Venuslike orbit yields a greater detection rate and enables the observation of objects in Earthlike orbits. Besides the fact that PHO signatures are stronger in the IR wavelengths, these 

systems offer less source confusion and a lower spacecraft data rate. The solar avoidance 

zone also is smaller for IR detectors. As with other space-based systems, this asset runs a 

higher risk of catastrophic failure and it is impossible to maintain or upgrade. In addition, 

dust in the galactic plane can confuse the IR structure. And last, infrared systems have 

less flight heritage, although this concept would be able to use detector technologies 

developed for JWST and the Wide-field Infrared Survey Explorer (WISE). 
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D-13: 0.5m IR Venus-like

Technical Description

• Space-based survey telescope 

• Infrared (6-10 μm), 1.7° x 6.8° FOV, Vlim= 23.6-25.4 

• 0.5m aperture

• Heliocentric, Venus-like orbit (can be Venus-trailing or other elliptical orbit 

inside of 1 AU)

• Passively cooled

CONOPS

• Location: Heliocentric, Venus-like orbit

• Operational: 2014

• Availability: 100% in NEO detection mode

• Launch Vehicle: Delta II 2926

• Full bandpass used to maximize detection 

• 10 year mission life

Rationale for Consideration

• Provide comparison with space-based Vis platforms

• Venus orbit has greater Inner Earth Orbit (IEO) object visibility

Capability

• NEO orbital parameters

Pros

• Better location to observe NEOs with Earth-like orbits

• No atmospheric absorbance in IR vs. ground-based asset

• Can conduct survey at all times (vs. night only on ground)

• Less source confusion and lower data rate than Vis

• Can look as close as ~55 degrees from Sun

Cons

• Leverages not yet flown systems (JWST, WISE)

• Higher risk of catastrophic failure vs. ground based asset

• Very difficult or impossible to maintain or upgrade * Note: Passive Venus flyby is employed to position S/C in approximately a 0.6 by 0.8 AU orbit. 

Hydrazine used for orbit corrections during cruise-phase of the mission.

Wet Mass (kg) 698

84

259

P/L

Dry Mass (kg) 663

S/C

Power (W) 302

Wet Mass (kg) 698

84

259

P/L

Dry Mass (kg) 663

S/C

Power (W) 302

69

20

125

9

46

25

109

404

259

Dry Mass (kg)

Thermal 10%

TT&C 3%

Structure & Mechanisms 19%

Propulsion 1%

Power 7%

C&DH 4%

ACDS 16%

Spacecraft Bus 61%

Payload 39%

% of Dry Mass

69

20

125

9

46

25

109

404

259

Dry Mass (kg)

Thermal 10%

TT&C 3%

Structure & Mechanisms 19%

Propulsion 1%

Power 7%

C&DH 4%

ACDS 16%

Spacecraft Bus 61%

Payload 39%

% of Dry Mass

System Sizing

E.1.14. 1-meter Infrared (IR) Venus-like [D-14]

Description

The 1-meter Infrared (IR) Venus-like concept is a space-based, infrared survey telescope 

that would scan the sky in search of NEOs. It is based on the proposed NEOCam, but 

would orbit the Sun in a ~0.7 AU heliocentric, Venus-like orbit. The study team included 

the option because PHO signatures are stronger in the IR-spectral bandpass. This 

concept’s orbit also offers greater IEO visibility and its telescope has a larger aperture 

that could enhance the PHO detection rate. This detection asset will search for PHOs and 

can be used to calculate their orbital parameters. The design is based on other infrared, 

space-based telescopes. 

The telescope would detect light in the infrared region between 6-10 μm. It would have a 

field of view of 1.5° x 6.0° and a limiting magnitude (Vlim) of approximately 24.2-26.0. 

The IR detectors are passively cooled because the spacecraft would be in a heliocentric 

orbit. Expected to become operational by 2014, the spacecraft would fly on a Delta IV 

Medium. A passive Venus flyby is employed to position the spacecraft in approximately 

a 0.6 by 0.8 AU orbit. Hydrazine propulsion is used only for orbital correction during the 

mission’s cruise phase. A dedicated PHO detection asset, it uses the full bandpass to 

maximize detection. The spacecraft has an operational life of 10 years. 
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D-14: 1m IR Venus-like

Technical Description

• Space-based survey telescope

• Infrared (6-10 μm), 1.5° x 6.0° FOV, Vlim= 24.2-26.0 

• 1m aperture

• Heliocentric, Venus-like orbit (can be Venus-trailing or other elliptical orbit 

inside of 1 AU)

• Passively cooled

CONOPS

• Location: Heliocentric, Venus-like orbit

• Operational: 2014

• Availability: 100% in NEO detection mode

• Launch Vehicle: Delta IV Medium

• Full bandpass used to maximize detection 

• 10 year mission life

Rationale for Consideration

• Larger aperture IR may enhance detection rate & characterization

• Venus orbit has greater Inner Earth Orbit (IEO) object visibility

Capability

• NEO orbital parameters

Pros

• Better location to observe NEOs with Earth-like orbits

• No atmospheric absorbance in IR vs. ground-based asset

• Can conduct survey at all times (vs. night only on ground)

• Less source confusion and lower data rate than Vis

• Can look as close as ~55 degrees from Sun

Cons

• Leverages not yet flown systems (JWST, WISE)

• Higher risk of catastrophic failure vs. ground based asset

• Very difficult or impossible to maintain or upgrade

• Reduced FOV due to bigger aperture for same telescope length * Note: Passive Venus flyby is employed to position S/C in approximately a 0.6 by 0.8 AU orbit. 

Hydrazine used for orbit corrections during cruise-phase of the mission.

Wet Mass (kg) 1210

143

427

P/L

Dry Mass (kg) 1149

S/C

Power (W) 449

Wet Mass (kg) 1210

143

427

P/L

Dry Mass (kg) 1149

S/C

Power (W) 449

124

36

225

15

83

44

195

722

427

Dry Mass (kg)

Thermal 11%

TT&C 3%

Structure & Mechanisms 20%

Propulsion 1%

Power 7%

C&DH 4%

ACDS 17%

Spacecraft Bus 63%

Payload 37%

% of Dry Mass

124

36

225

15

83

44

195

722

427

Dry Mass (kg)

Thermal 11%

TT&C 3%

Structure & Mechanisms 20%

Propulsion 1%

Power 7%

C&DH 4%

ACDS 17%

Spacecraft Bus 63%

Payload 37%

% of Dry Mass

System Sizing

Qualitative Assessment

This system offers all of the same benefits of other infrared systems. In addition, its 

Venus-like orbit yields a greater detection rate and enables the observation of objects in 

Earth-like orbits. It also has the same disadvantages as other infrared space-based 

systems: it is difficult, if not impossible, to upgrade and maintain in the event of a failure 

and NASA has less experience building and flying them. There also is less historical 

flight heritage for infrared systems in space, although it will be able to use detector 

technologies developed for JWST and the Wide-field Infrared Survey Explorer (WISE). 

E.2. Precision Orbit Determination Alternatives 

E.2.1. Radar

Two planetary radars are capable of observing NEOs. The Goldstone radar, equipped 

with a 70-meter antenna and a 450 kW transmitter, is located in southern California’s 

Mojave Desert. Because the antenna can be steered, Goldstone can observe most of the 

sky and follow the often-rapid motions exhibited by many NEOs. The second radar, 

equipped with a 305-meter aperture and a 900 kW transmitter, is located at Arecibo, 

Puerto Rico. Its reach is farther than Goldstone’s, but because it has a fixed antenna, it 

can only look about 20 degrees off its zenith. As of early September 2006, planetary 

radars had observed 195 near-Earth asteroids and 11 comets. 

Both Arecibo and Goldstone radars are heavily subscribed; only a percentage of their 

time is available for asteroid radar. Financial support for each system also has declined 

during the past several years and is inadequate for sustaining the radar capabilities. As a 
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result, the systems have experienced difficulties mainly because of transmitter problems 

and difficulties maintaining personnel. 

Both sites need adequate, reliable financial support to continue operations. As the current 

survey operates and the rate of discoveries increases, it will become increasingly clear 

that their limitations preclude most NEO radar reconnaissance. If radar observations of 

“short warning” PHOs are required, it is likely that a dedicated NEO radar observatory 

will be needed. 

The availability of radar to observe an object over various time periods is shown in 

Figure 24, and the benefits of radar observations for characterization are discussed further 

in Reference [53]. 

E.3. Data Management System Alternatives 

E.3.1. Scale Existing Data Management Systems [D-15]

Description

When the next-generation surveys begin, the Minor Planet Center (MPC) will need to 

increase its staff and modernize some of the computing resources to handle the expected 

increase in data, which may be up to 100 times greater than what is generated by current 

missions. 

The MPC serves as the international clearinghouse for all worldwide asteroid, comet, and 

satellite astrometric and positional measurements. It is responsible for distributing 

astrometric observations and orbits via the Minor Planet Electronic Circulars (issued as 

necessary, generally once per day) and related catalogs. It specifically identifies NEOs, 

determines their short-arc orbits, and disseminates information about them. The MPC 

catalog is expected to grow to about 20,000 to 100,000 NEO records and perhaps 10 

million total solar system objects after 10 years of Pan-STARRS (PS 4) operation. 

The Smithsonian Astrophysical Observatory, in coordination with the International 

Astronomical Union (IAU), operates MPC. The MPC is located in Cambridge, MA, and 

has a current staff of two to three full-time equivalents (FTEs). As future detection 

capabilities grow 10 to 100 times the current observation rates, the staff size would likely 

need to increase to five or six FTEs. A one-to-two order of magnitude increase in 

observations will result in 200 to 10 million observations per day for minor planets and 

comets, and 20 to 1,000 observations per day for NEOs. The increased data flow also will 

require an increase in the number of objects needing follow-up.

In addition to distributing orbit and astrometric catalogs for said objects, the MPC 

facilitates follow-up observations of potential new NEOs. It places candidate sky-plane 

ephemeredes and uncertainty maps on the Web via the NEO Confirmation Page. In most 

cases, NEO observations are distributed to the public, free of charge, within 24 hours of 

receipt. The MPC also provides a variety of tools to support the NEO initiative, including 

sky-coverage maps, lists of known NEOs, lists of NEO discoverers, and a page that 

allows users to select a list of known NEOs in need of astrometric follow-up. The MPC 

also maintains a suite of programs to calculate the probability that any object is a new 

NEO based simply upon two sky-plane positions and a magnitude. 
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Qualitative Assessment

The MPC has a well-functioning infrastructure that is operated by the Smithsonian 

Astrophysical Observatory and coordinated with the IAU. As the next generation of 

large-scale surveys begin observing more NEOs, additional computer resources and a 

larger staff are necessary to support the increase in data rate. However, scaling the MPC 

is dependent on these detection systems producing a one-to-two order of magnitude 

increase in NEO observation data rate. Without this data increase it may be unnecessary 

to scale the MPC or similar data management systems. 

E.3.2. Adopt Similar Systems [D-16]

Description

A new data-management system, or a combination of data-management systems, may 

have better functionality than existing NEO data systems. NEO detection programs may 

leverage existing systems to save money or reduce schedule and other risks. Multimission ground data system infrastructures currently exist to support astrophysical and 

geophysical data acquisition, processing, product development, and dissemination. 

For example, the Infrared Processing Analysis Center (IPAC), the Multi-Mission Archive 

at Space Telescope, and the National Geophysical Data Center (NGDC) provide good 

examples. They can provide level-0 data acquisition and archival, data product 

development and dissemination, data fusion from multiple sources, and asset 

prioritization and tasking. It is important to fund and support similar services for NEO 

detection. This could involve developing and installing mission-unique hardware and 

software as required, using existing processes and data-archival services, and developing 

and distributing products. 

Qualitative Assessment

Current NEO data-management infrastructures can be upgraded with potentially new 

features to improve the data-management capabilities. Better data-storage techniques, 

improved query functionality, and better dissemination channels are some features that 

can be adopted. Prioritization of services also could be a concern because of potential 

conflicts among users. 

E.3.3. New Central Repository [D-17]

Description

Breakthroughs in telescope, detector, and computer technology allow astronomical 

surveys to produce terabytes of images and catalogs. [46] The U.S. National Virtual 

Observatory (NVO) can provide proper data management to catalog this increase in data 

flow, and is therefore the proposed framework for a NEO data management virtual 

observatory. NVO is developing new protocols and standards for data exchange and 

access to catalogued NEO observations, which is making it easier to use the data. The 

NVO has built prototypes to demonstrate the effectiveness of these new protocols and 

standards. 

The NVO is a member of the International Virtual Observatory Alliance (IVOA), whose 

mission includes facilitating “...the international coordination and collaboration 
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necessary for the development and deployment of the tools, systems, and organizational 

structures necessary to enable the international utilization of astronomical archives as an 

integrated and interoperating virtual observatory.” [47] As of January 2005, 15 different 

countries funded NVO projects. In the U.S., the NVO effort began in August 2001. More 

than 30 collaborators, principally U.S. professors with international liaisons in Europe, 

Japan, and Australia, support the NSF 5-year grant. 

Qualitative Assessment

The NVO framework describes multi-terabyte online databases with interlinked catalogs. 

Query engines will become more sophisticated and the research results from the online 

data can potentially be as rich as those from real telescopes. New standards are being 

developed with the international astronomical community. The framework only is being 

researched and proposed; a virtual observatory has not been developed yet. 

E.3.4. Backup Facilities [D-18]

Description

As the next generation of large-scale surveys begins observing NEOs, the Minor Planet 

Center (MPC) will need to increase staff and modernize some of the computing resources 

to handle a 10x to 100x increase in the observation data rate. During the interim, the 

NVO can be developed into a virtual observatory as the MPC is scaled to provide longterm data management. After it is fully scaled, the MPC will be the primary system and 

the NVO will become the backup facility. 

As the MPC is being scaled to accommodate an increase in NEO observation data, the 

NVO framework can be implemented to manage the incoming data. The NVO is 

developing new protocols and standards for data exchange and access of cataloging NEO 

observations. The NVO will make astronomical data easier to use through the creation 

and adoption of standards. The NVO will build a few new protocols and standards for 

data exchange and access. The NVO has the main collaborators in the US with liaisons in 

Europe, Japan and Australia. The MPC is located in Cambridge, MA. 

The MPC serves as the international clearinghouse for all asteroid, comet and satellite 

astrometric and positional measurements obtained worldwide. It is responsible for the 

dissemination of astrometric observations and orbits via the Minor Planet Electronic 

Circulars (issued as necessary, generally once per day) and related catalogs. It focuses 

specifically on identification, short-arc orbit determination, and dissemination of 

information pertaining to NEOs. The MPC catalog will grow to 20k-100k NEO records 

after 10 years of Pan-STARRS (PS 4) operation, and perhaps 10 million total objects. 

After the system is scaled, it will be used as the primary data-management facility. For 

purposes of redundancy, the NVO will become the secondary system. 

Qualitative Assessment

A one-to-two order of magnitude increase in observations will result in 200 to 10 million 

observations/day for minor planets and comets and 20 to 1,000 observations per day for 

NEOs. This increase in data flow also will require an increase in the number of objects 

needing follow-up. The MPC will be scaled over a several years to manage this data 

increase. During this time, the NVO will be developed to process this data. The NVO 
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framework is a multi-terabyte online database with interlinked catalogs. Query engines 

will become more sophisticated and the research results from the online data can 

potentially be just as rich as that from real telescopes.

This will address the near-term and long-term NEO observation data-management needs. 

The NVO framework is a proposed protocol and standard for data exchange and 

cataloging; so it is uncertain when it will be funded and become a functioning system. 
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Appendix F. Descriptions of Characterization Alternatives 

F.1. Object Characterization Alternatives 

F.1.1. Remote Characterization – Optical and Near Infrared (to 1um)

In the 1970s, astronomers began classifying asteroids into taxonomically distinct groups 

based on observations through three or four spectral filters. The broad groups included 

“C” for dark carbonaceous objects and “S” for stony (silicaceous) objects. Since then, the 

approach has been expanded to include several widely used systems using a progressively 

larger number of narrower bands or medium resolution spectro-photometry over the 0.4-

um to 1-um region of the spectrum. These results will drive the follow-on observations, 

but taxonomy only can crudely bracket crucial parameters, such as mass, and only 

approximate composition. Asteroid taxonomy is covered in greater detail in Appendix 

Section K.4. 

Spitzer Space Telescope (formerly SIRTF)

The Spitzer Space Telescope was designed to track solar system objects moving at 1”/sec 

(3600”/hr) and is capable of observing a large fraction of the NEO population. Spitzer 

provides both imaging and spectroscopic data for size and compositional analyses. When 

compared with ground-based surveys, Spitzer is very strong in the areas of limiting 

sensitivity and wavelength. Although observers will find many as-yet-undiscovered 

NEOs closer to the Sun than the Earth, Spitzer will not find them because of its strict 

Sun-avoidance constraints (Spitzer observations are only possible within 80-120o solar 

elongation). 

Spitzer also does not automatically examine data within hours and it cannot routinely 

command rapid follow-up observations. Although Spitzer is not an optimal facility for 

NEO search, it will make major contributions in the characterization of the NEO 

population. It is the best available facility to provide spectroscopic and photometric 

studies of recently discovered objects that are too faint for detailed study with groundbased telescopes. The spectroscopic information will provide clues to an asteroid’s 

composition, and the photometric information will derive an infrared albedo. As a result, 

observers will obtain a highly accurate measure of the object’s size, especially compared 

with what an optical system could provide alone. Spitzer’s imaging instruments are 

designed to routinely take simultaneous or near-simultaneous observations of the same 

region of sky at multiple wavelengths, which may lead to more robust detection of 

serendipitous NEO trails. 

Spitzer sensitivity to these objects depends not on the integration time of the frame, but 

the amount of time that the object spends on each pixel. This is governed by the object’s 

rate of motion. For a rapidly moving object, this may be only a few seconds. However, 

depending on the geometry with which astronomers view these objects, NEOs may NOT 

move rapidly. They may instead move as fast as the main belt asteroids. As a 

consequence, they appear near stationary for days at a time. In the latter case, reasonable 

integration times of 120 seconds for IRAC at 8um would allow astronomers to detect and 
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characterize objects as small as 32 meters in diameter at a distance of 0.5 AU from 

Spitzer. [48] 

F.1.2. Remote Characterization – Infrared

The infrared spectral region is most important for discerning an object’s composition 

(carbonaceous, stony, or metallic), which is needed to constrain the NEO mass and to 

determine the most effective deflection or mitigation strategy. Figure 85 of Appendix 

Section K.4 demonstrates usefulness of the 1- to 2.5-micron spectral region. Therefore, 

coverage of the near-infrared spectral region has the highest priority. Remote infrared 

observations have a much greater range than radar observations, and thus afford more 

planning time. 

Observations using modestly sized telescopes to observe NEOs at close approach and 

observations of newly identified PHOs shortly after their discovery are preferred. This 

requires access to large telescopes and a cadre of observers. In addition, for ground-based 

applications, polarimetry may be more effective than thermal IR observations. Table 44 

details information derived from selected techniques. 

Table 44. Mapping of Characterization Technique to Derived Information 

Technique 

Observed 

Parameter Derived Information 

Photometry Light curve Spin state, axial ratio, presence of 

satellites, constraint on density 

(cohesion?), absolute magnitude, 

phase curve, obliquity and shape 

modeling 

Multi-wavelength 

photometry 

Broad classification First-order information: 

carbonaceous, stony, iron 

Spectroscopy Surface mineralogy First-order information: density 

and porosity by analogy to 

meteorites and spacecraft data 

Polarimetry (preferably 

at visible wavelengths 

but can also be done a 

near-infrared 

wavelengths) 

Polarization as 

function of 

heliocentric phase 

angle 

Polarimetric albedo: size can be 

estimated when used with the 

absolute magnitude; surface 

particle size and roughness 

information 

Thermal photometry Thermal emission Albedo (when combined with 

visible photometry); thermal 

inertia gives estimate of the 

beaming parameter 

Adaptive optics Satellite detection 

and orbits 

Mass of primary 

F.1.3. Remote Characterization – Radar

As noted in section 5.10.2, radar data are capable of quickly refining orbits of some 

PHOs, especially for those objects with only a short optical data interval. Radial radar 
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range measurements (i.e., round-trip light time) and radar radial velocity measurements 

(Doppler) can be accurate to the 8-meter and 1-mm-per-second levels, respectively, and 

hence have a fractional precision that is orders of magnitude better than optical position 

measurements. 

Radar observations are capable of remotely characterizing, or constraining, a PHO’s size, 

shape, surface roughness, and bulk density. In addition, of the 30 NEOs for which 

satellites were identified, 20 were found using radar techniques. The tracking of an 

asteroid’s satellite can be used, through Kepler’s third law, to determine the mass of the 

primary body. This mass, together with the radar-determined size (i.e., volume), can be 

used to determine the asteroid’s bulk density. The information available from the radar 

characterization of a PHO is comparable to that of a spacecraft mission. 

Images built from Doppler radar, which measures the frequency shift of a signal 

reflection to measure the speed of the target, can reconstruct the asteroid’s actual shape. 

[49] [50] These analyses are made easier by knowing the object’s rotational period and 

orientation, which ground-based photometric observations may help to derive. 

In most modern radar observations, the transmission is circularly polarized and the ratio 

of the strength of two parallel receiving channel signals are used to gauge the target 

body’s near-surface roughness at scale lengths comparable to the radar wavelengths 

(3.5 cm for Goldstone and 13 cm for Arecibo). The radar albedo, or reflectivity, also 

provides a useful constraint upon the object’s characteristics, such as porosity and metal 

abundance. Radar imagery also may be used to determine the presence of an orbiting 

satellite, which can reveal the mass of the primary body. 

Figure 69 and Figure 70 illustrate the ability of radar observations to characterize the 

shapes of asteroids and to provide constraints upon the object’s surface properties. The 

radar observations in Figure 69 show that the main belt asteroid (216), Kleopatra, is 

shaped like a dog bone. Its surface is porous and loosely consolidated. Its interior may be 

composed of an arrangement of solid-metal fragments and loose metallic rubble. Figure 

courtesy of Steve Ostro, JPL. [51] 

Figure 69. Radar Observations of Asteroid (216) Kleopatra 

Figure 70 shows radar images of PHO (4179), Toutatis, which were made during the 

object’s close approach to Earth on December 8, 1992. The images reveal two irregularly 

shaped, cratered objects about 4 and 2.5 km in average diameter. They are probably in 

contact with each other. The four frames shown here (from left to right) were obtained on 

Dec. 8, 9, 10 and 13, when Toutatis was about 4 million km from Earth. On each day, the 

asteroid was in a different orientation with respect to Earth. In these images, the radar 
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DRAFT Pre-Decisional Material 181 

illumination comes from the top of the page; so parts of each component facing toward 

the bottom are not seen. The large crater shown in the Dec. 9 image (upper right) is about 

700 meters in diameter. Courtesy, Steve Ostro, JPL [52]. 

The availability of radar to observe an object over various time periods is shown in 

Figure 24. The benefits of radar observations for characterization are discussed further in 

Reference [53]. 

Figure 70. Radar Images of Asteroid (4179) Toutatis 

F.1.4. In-Situ Characterization – Flyby

In-situ characterization missions are divided into two classes. Flyby missions, which also 

may include probes or impactors, are the least complex in-situ missions and may visit 

multiple NEOs. Orbital missions may include optional landers and surface samplers and 

are significantly more complex. The lower cost and complexity of flyby missions, along 

with their ability to visit multiple NEOs with a single launch, make them ideal for 

characterizing diverse objects in a population. 

The most important objectives of a flyby mission would be to determine the key physical 

parameters of the target body. Flyby missions can provide accurate measures of size, 

shape, volume, albedo, general structure, rotation rate, pole location, and whether the 

body has any satellites. A flyby mission cannot measure the mass directly because the 

spacecraft is moving too quickly to measure the gravitational deflection of its trajectory 

by a small NEO. However, a flyby mission does provide the key parameters that are 

important for determining the body’s mass. Once scientists know the body’s volume, they 

only need to infer the porosity and average density to calculate the mass. 

The objectives of a flyby mission can be classified into Tier-1 objectives, which are key 

to many deflection techniques and are readily measured by flybys. Tier-2 objectives are 

items that would be very helpful to understand the NEO population and may be important 

to know for some deflection alternatives. It may be very difficult for a flyby mission to 
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obtain accurate measures of these Tier-2 objectives. However, the addition of an impactor 

to a flyby mission, similar to the Deep Impact experience, could help to achieve some of 

the Tier-2 objectives. These objectives are shown in Table 45. 

Table 45. Possible Flyby Characterization Objectives 

Tier-1 Objectives Tier-2 Objectives (Difficult) 

• Size 

• Shape 

• Volume 

• General Structure (rubble pile, 

or cohesive object) 

• Rotation rate and approximate 

axis 

• Albedo 

• Satellites? 

• Mass 

• Density 

• Porosity 

• Detailed orbital elements (by 

delivering a coherent transponder to 

the surface) 

• Interior structure (by employing 

explosives and seismometers, or 

radar tomography, or other 

techniques for examining the 

interior) 

• Cohesiveness and strength of the 

NEO near surface 

• Composition of the NEO (both 

mineralogical and atomic) 

There are two basic options for examining a number of NEOs with a single flyby mission. 

One option uses the flexibility of the trajectory to visit multiple NEOs, while the other 

uses a great deal of onboard propulsion. There are cost and mission objective tradeoffs 

involved in selecting which option to employ. 

The first option is a multi-body flyby tour mission. Its orbit is similar to those of the 

CONTOUR mission and the Galileo and Cassini tours of Jupiter and Saturn’s moon 

systems. The concept uses a high-energy Earth-return orbit. This mission would fly by 

one or two NEOs per year and then return to Earth where planners would program it for 

the next NEO flyby. Each of these yearly excursions forms a petal of the overall 

trajectory. This approach requires relatively little propulsion, as the spacecraft swings by 

Earth each year before it begins its next flyby. 

Figure 71 shows some examples of Earth-return orbits that could employ chemical 

propulsion systems. The principal advantage of this type of mission is that the number of 

objects visited only depends on the mission’s duration. This type of mission requires very 

little onboard propulsion and could easily fly by a large number of objects. The flyby 

speeds for these missions are several km/s, and therefore probably could only accomplish 

Tier-1 objectives without impactors. With impactors, they could get information on the 

targets’ interior structures, thereby fulfilling a number of the Tier-2 objectives. 
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Figure 71. Examples of 1-year and multi-year Earth-return orbits 

The second option for a multiple NEO flyby mission is an electric propulsion mission, 

either solar-electric or radioisotope-electric. While the exact amount of onboard ΔV 

required depends on the NEO targets chosen, most targets require at least 500 m/s to 

1,500 m/s of onboard ΔV. Because of their high specific impulse (specific impulse from 

600 seconds to greater than 3,000 seconds, depending on the type of electric thrusters), 

electric propulsion flyby missions can carry several km/s of propulsion. This should be 

sufficient to visit several NEOs, where the exact number depends on their orbital 

parameters. 

The payload required for a flyby mission to accomplish its Tier-1 objectives includes 

imagers and spectrographs, plus some optional instruments that also could help it achieve 

some of the Tier-2 objectives. At least some imagers must have sufficiently high 

resolution to find the target body many days before their scheduled flyby. The 

information can help guide the spacecraft to the target, especially when the target’s 

ephemeris is not known precisely. Other imaging assets must be able to resolve the size, 

shape, and detailed surface features during the high-speed flyby. This will require a 

tracking-mirror system for the medium-resolution camera and spectrograph. The 

spectrograph must have broad spectral coverage in the visible and near infrared to be able 

to observe the expected features of the various minerals on the NEO’s surface. Additional 

payload elements may include a laser or radar altimeter magnetometer, neutron 

spectrometer, and X-ray spectrometer. 

F.1.5. In-Situ Characterization – Rendezvous

While flyby missions can determine many of the NEO’s physical characteristics, only a 

rendezvous mission with an orbital phase and possibly a lander can completely 

characterize a target. Flyby missions can tell a great deal about the diversity of the NEO 

population, but to learn more about a particular body’s mass, structure, and composition, 

a rendezvous mission is required. 

Rendezvous missions are designed to characterize a single body in great detail. The 

mission would get a range of characteristics about the target body, which are particularly 

important for selecting an appropriate deflection technique. If the mission deploys 

separate landers, the information they gather can reveal much about the body’s interior, 

more so that what an orbiter alone can tell. 

Rendezvous missions fulfill a wider range of Tier-1 objectives than flyby missions. 

Tier-1 objectives include those of a flyby mission, plus an accurate measure of the body’s 

mass, its precise orbital motion, and a number of surface characteristics. However, a 
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2006 Near-Earth Object Survey and Deflection Study 

184 DRAFT Pre-Decisional Material

rendezvous mission requires more propulsion to arrive at the target body with zero 

relative velocity. This usually means that it must carry several times the onboard 

propulsion of a flyby mission to the same body. It also may require a trajectory that 

includes up to several planetary swingbys to reach the target body. See Table 46 for 

Tier-1 and Tier-2 rendezvous characterization objectives. 

Table 46. Possible Rendezvous Characterization Objectives 

Tier-1 Objectives Tier-2 Objectives 

• Mass 

• Detailed orbital elements 

• Size and Shape 

• General Structure (rubble pile, 

or cohesive object) 

• Rotation rate and axis 

• Albedo 

• Scan for satellites or other 

bodies 

• Average density 

• Porosity 

• Surface structure and its 

variations 

• Surface composition 

• Determine the interior structure 

(seismic, radio tomography, ground 

penetrating RADAR, etc.) 

• Demonstrate landing 

• Demonstrate anchoring to the 

surface with a given tensile strength 

• Demonstrate drilling the surface 

• Determine cohesiveness and 

strength near the surface 

Missions with a landed component may enhance the return of information to support 

multiple deflection techniques. The payload of a landed component can provide more 

information about the body’s surface composition and its interior structure. It also can 

experiment with landing, anchoring, and possibly drilling techniques, which could be 

used in a mitigation strategy. The payload for a rendezvous mission has all of the 

instruments of a flyby mission; however, it does not require a tracking-mirror system 

because there is about zero velocity between the orbiter and the NEO. In addition to the 

imagers, spectrographs, altimeter, and magnetometer, a landed component could have a: 

• Low-power microscope to examine surface structure 

• Laser-ablation mass spectrometer to study the body’s molecular, atomic and 

isotopic composition 

• Neutron spectrometer to look for hydrated minerals 

• Alpha-proton-X-ray spectrometer or solar X-ray fluorescence spectrometer to 

examine the atomic composition of the surface 

• Gamma-ray spectrometer to examine the atomic composition of the subsurface 

• Plasma spectrometer provide information about its interaction with the solar wind 
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F.1.6. Ground-Based Characterization Concept Sheets

• Location: Mauna Kea, HI

• Operational: Currently operational

• Availability: On request, prioritized with others

• Provide follow-on observations of NEO detected by surveying telescopes 

upon request

C-1: Shared InfraRed Telescope Facility (IRTF)

Technical Description

• Currently operational, NASA owned 

• Follow-on observations of NEOs without disturbing detection effort & 

improved accuracy

Rationale for consideration

Telescopes in System 1

Locations 1

Aperture (m) 3.0

Telescopes in System 1

Locations 1

Aperture (m) 3.0

CONOPS

Pros

• Much improved estimate of size, shape & albedo vs. only visible

• NEOs most detectable in IR

• Low risk of catastrophic failure, easy to maintain & upgrade vs. space asset

Cons

• Atmosphere absorbs majority of IR light, thus limited viewing of NEOs

• Limited access, observing time allocated 2x per year 

System Sizing

• Ground based telescope, alt-azimuth mount

• Infrared (0.8 - 5.5 μm), 80’ x 80’ FOV, Vlim= 22.1 – 23.9 

• Monolithic, 3.0m aperture

• Current camera is not optimized for NEO characterization, a new, optimized 

camera could improve NEO characterization by a factor of 2

Capability

• NEO orbital parameters, composition, size

• Location: Mauna Kea, HI (North) and Cerro Pachón, Chile (South)

• Operational: Currently operational

• Availability: On request

• Provide follow-on observations of NEO detected by surveying telescopes 

upon request

C-2: Gemini

Technical Description

• Currently operational, multi-nation backing (including NSF) 

• Follow-on observations of NEOs without disturbing detection effort & 

improved accuracy

Rationale for consideration

Telescopes in System 2

Locations 2

Aperture (m) 8.1

Telescopes in System 2

Locations 2

Aperture (m) 8.1

CONOPS

Pros

• Much improved estimate of size, shape & albedo vs. only visible

• NEOs most detectable in IR

•Two 8m class telescopes at each hemisphere maximize sky access & allow 

parallax for orbital determination

Cons

• Atmosphere absorbs majority of IR light, thus limited viewing of NEOs

System Sizing

• Twin ground based telescopes in Northern and Southern hemispheres, altazimuth mount

• Visible (360-1100 nm) and Infrared (1-5.5 μm), 0.75° x 0.75° FOV, Vlim= 25

• Monolithic, 8.1m aperture

AND

Capability

• NEO orbital parameters, composition, size
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• Location: Happy Jack, 40 miles SE of Flagstaff, AZ

• Operational: 2009

• Availability: On request, less limited

• Provide follow-on observations of NEO detected by surveying telescopes 

upon request

C-3: Discovery Channel Telescope (DCT)

Technical Description

• Development underway, partnership with Discovery Communications, Inc. 

• Follow-on observations of NEOs without disturbing detection effort

Rationale for consideration

• NEO orbital parameters, visible spectroscopy

Capability

Telescopes in System 1

Locations 1

Aperture (m) 4.2

Telescopes in System 1

Locations 1

Aperture (m) 4.2

CONOPS

Pros

• Inexpensive follow-on of NEOs with Vlim < 23

Cons

• Vlim= 23 will not be able to provide follow up for smaller NEOs detected 

by an LSST or PS4-like telescope

System Sizing

• Ground based telescope, alt-azimuth mount

• Two modes: Prime Focus (2° FOV) and Ritchey-Chrétien Focus (30' FOV)

• Visible (330-1000 nm) Vlim= 23 

• Monolithic, 4.2m aperture

• Location: Mauna Kea, HI

• Operational: Currently operational

• Availability: 100% NEO characterization mode

• Provide follow-on observations of NEO detected by surveying telescopes 

upon request

C-7: Dedicated IRTF

Technical Description

• Optimized NEO characterization versus shared asset 

• Follow-on observations of NEOs without disturbing detection effort & 

improved accuracy

• Dedicated, thus can quickly respond to characterizing new NEO detections

Rationale for consideration

Telescopes in System 1

Locations 1

Aperture (m) 3.0

Telescopes in System 1

Locations 1

Aperture (m) 3.0

CONOPS

Pros

• Unlimited & rapid access, dedicated to NEO characterization

• Much improved estimate of size, shape & albedo vs. only visible

• NEOs most detectable in IR

• Low risk of catastrophic failure, easy to maintain & upgrade vs. space asset

Cons

• Atmosphere absorbs majority of IR light, thus limited viewing of NEOs

System Sizing

• Ground based telescope, alt-azimuth mount

• Infrared (0.8 - 5.5 μm), 80’ x 80’ FOV, Vlim= 22.1 – 23.9 

• Monolithic, 3.0m aperture

• Current camera is not optimized for NEO characterization, a new, optimized 

camera could improve NEO characterization by a factor of 2

Capability

• NEO orbital parameters, composition, size
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C-10: Arecibo

Technical Description

• Ground based telescope, fixed in ground

• Radar (3 – 600 cm), 305m dish

• Bowl-shaped reflector fixed in the ground, with a movable Gregorian 

reflector system suspended 137 meters hanging above it 

• Approximately 40 kilometers of cabling supports the reflector. This 

prevents the Gregorian reflector system from changing shape as 

temperatures fluctuate and winds blow.

Power (S-band) 1 MW

Zenith angle range 1.06 - 19.69 deg

Elevation Slew Rate 0.04 deg/sec

Azimuth Slew Rate 0.4 deg/sec

Frequency Range 327 – 8500 MHz

Surface accuracy 2 mm rms

Reflector diameter 304.8-m

Wavelength (S-band) 13-cm

Power (S-band) 1 MW

Zenith angle range 1.06 - 19.69 deg

Elevation Slew Rate 0.04 deg/sec

Azimuth Slew Rate 0.4 deg/sec

Frequency Range 327 – 8500 MHz

Surface accuracy 2 mm rms

Reflector diameter 304.8-m

Wavelength (S-band) 13-cm

Pros

• Inexpensive follow on characterization of NEOs

• Large focusing antenna requires a few minutes of observation for collecting 

enough energy for analysis

• NSF supports operational, maintenance and upgrading costs

Cons

• Is limited to observations ~0.1AU or closer

• Declining financial support despite requiring receiver upgrade, small optics 

changes, upgrading of data acquisition system

• Small percentage of time available for asteroid radar

Rationale for consideration

• Currently operational, backed by NSF

• Can provide high-precision characterization for NEOs and PHAs 

CONOPS

• Location: Arecibo, Puerto Rico

• Operational: Currently operational

• Availability: On request 

• 13-cm wavelength (S-band) transmitter

System Sizing

Capability

• NEO orbital parameters, composition, mass, size, binarity, rotation , shape

C-11: Goldstone

Technical Description

• Ground based telescope, steerable

• Radar (3.5 cm), 70m dish

Wavelength (X-band) 3.5-cm

Power (X-band) 500 kW

Reflector diameter 70-m

Wavelength (X-band) 3.5-cm

Power (X-band) 500 kW

Reflector diameter 70-m

Pros

• Large focusing antenna requires a few minutes of observation for collecting 

enough energy for analysis

• Steerable vs. Arecibo

Cons

• Small percentage of time available for asteroid radar

• Declining financial support 

• Is limited to observations ~0.1AU or closer

CONOPS

• Location: Goldstone, CA

• Operational: Currently operational

• Availability: On request 

•Communicates in X-band (8.6GHz) and S-band (2.3 GHz)

• Primarily uses 3.5cm wavelength (X-band) transmitter

Rationale for consideration

• Currently operational

• Can provide high-precision characterization for NEOs and PHAs 

System Sizing

Capability

• NEO orbital parameters, composition, mass, size, binarity, rotation , shape
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C-12: Bistatic 100m

Technical Description

• Two ground based telescopes, steerable

• Radar (0.9 – 3.5 cm), 100m dish

• Ka and X band

• Antennas, receivers, and transmitters will be one order of magnitude more 

sensitive than Arecibo

• Megawatt transmitter, each antenna’s gain ~88dB

• Megawatt Transmitter: TRL 8 (Arecibo - 1 megawatt transmitter)

• Steerable 100m antenna: TRL 8 (NRAO Greenbank Telescope, West Virginia)

CONOPS

• Location: In conceptual design, no location identified

• Operational: TBD

• Availability: Dedicated (assuming funding incurred)

Pros

• No switching of antenna pointing between transmit and receive directions

• No interruption of data acquisition, coherent integration or orientational

coverage

• Doubled data integration time

• Accessibility of arbitrarily close targets, including those within a few EarthMoon distances (RTT < 10s)

Cons

• Very early in design process (concept stage), many unknowns

• Is limited to observations ~0.1AU or closer

Number of Scopes 2

Wavelengths (Ka and 0.9 to 3.5-cm

X-band)

Gain 88 dB

Power 1 MW

Reflector diameter 100-m

Number of Scopes 2

Wavelengths (Ka and 0.9 to 3.5-cm

X-band)

Gain 88 dB

Power 1 MW

Reflector diameter 100-m

System Sizing

Rationale for consideration

• Would be leap forward from current radar telescopes

• Dedicated to NEOs, could compile more observations in one year than 

Arecibo and Goldstone can in a decade

Capability

• NEO orbital parameters, composition, mass, size, binarity, rotation , shape

F.1.7. Space-based Characterization Concept Sheets

• Location: LEO, 600km altitude

• Operational: Currently operational

• Availability: On request, very limited

• Provide follow-on observations of NEO detected by surveying telescopes 

upon request

C-13: Hubble

Technical Description

• Currently operational 

• Follow-on observations of NEOs in emergencies

Rationale for consideration

CONOPS

Pros

• Already launched & operational

• Above atmosphere which absorbs majority of IR light

Cons

• Limited lifetime, will not be available long term (deorbit ~2010)

• Shared asset, thus compete for telescope time

• Space based telescope, pointable

• Ultraviolet, Visible & IR (115 – 2500 nm), 28’ x 28’ FOV, Vlim=??

• Monolithic, 2.4m aperture

5802

Payload

Dry Mass (kg) 11972

Observatory

Power (W) 2400

5802

Payload

Dry Mass (kg) 11972

Observatory

Power (W) 2400

System Sizing

Capability

• NEO orbital parameters, spectroscopy
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C-8: 2m Vis LEO/L1/L2 with Filter Wheel

Technical Description

• Space-based dual mode (detect & characterize) telescope, pointable

• Visible, 2.81° x 2.81° FOV, Vlim= 23.6 in detect mode 

• FOV & Vlim much less in characterize mode

• 2m aperture (a little smaller than Hubble)

• Modification to solely detection system

• Filter wheel added to provide narrow band pass for characterization

CONOPS

• Location: Earth-Sun L1 (nearly identical detection performance at LEO or L2)

• Operational: 2014 (TBR, may be affected by dual mode)

• Availability: Split between NEO detection and characterization modes

• Launch Vehicle: Atlas V 401 to L1/L2, Delta II 2920 for LEO

• 10 year mission life

Rationale for Consideration

• Add characterization capability with relatively minor modification to an 

asset that may be desired for detection purposes

Capability

• NEO orbital parameters, spectroscopy (if filter wheel used)

Pros

• If already in use for detection purposes, then may be cost effective design 

for dual (detect & characterize) purpose

Cons

• While asset in characterization mode, it would be unable to perform survey 

mission (will lose time which It could have been detecting NEOs)

• Will only be able to characterize NEOs with a Vlim much less than the 

telescope’s Vlim in detect mode

• Increased moving parts & complexity vs. no filter wheel

System Sizing

96

102

580

0

225

74

266

1344

1285

Dry Mass (kg)

Thermal 4%

TT&C 4%

Structure & Mechanisms 22%

Propulsion 0%

Power 9%

C&DH 3%

ACDS 10%

Spacecraft Bus 51%

Payload 49%

% of Dry Mass

96

102

580

0

225

74

266

1344

1285

Dry Mass (kg)

Thermal 4%

TT&C 4%

Structure & Mechanisms 22%

Propulsion 0%

Power 9%

C&DH 3%

ACDS 10%

Spacecraft Bus 51%

Payload 49%

% of Dry Mass

Wet Mass (kg) 2629

367

1285

P/L

Dry Mass (kg) 2629

S/C

Power (W) 1311

Wet Mass (kg) 2629

367

1285

P/L

Dry Mass (kg) 2629

S/C

Power (W) 1311

C-9: 2m Vis Venus-like with Filter Wheel

Technical Description

• Space-based dual mode (detect & characterize) telescope, pointable

• Visible, 2.81° x 2.81° FOV, Vlim= 23.6 in detect mode

•FOV & Vlim much less in characterize mode

• 2m aperture diameter (a little smaller than Hubble Observatory)

• Modification to solely detection system

• Filter wheel added to provide narrow band pass for characterization

• Venus-like orbit (can be Venus-trailing or other elliptical orbit inside of 1 AU)

CONOPS

• Location: Heliocentric, Venus-like orbit

• Operational: 2014 (TBR, may be affected by dual mode)

• Availability: Split between NEO detection and characterization modes

• Launch Vehicle: Delta IV Medium+ (5,4)

• Mostly autonomous spacecraft due to unique communication profile with 

Earth

Rationale for Consideration

• Add characterization capability with relatively minor modification to an 

asset that may be desired for detection purposes

• Venus orbit has greater Inner Earth Orbit (IEO) object visibility

Capability

• NEO orbital parameters, spectroscopy (if filter wheel used)

Pros

• If already in use for detection purposes, then may be cost effective design 

for dual (detect & characterize) purpose

• No atmospheric absorbance in IR vs. ground-based asset

Cons

• While asset in characterization mode, it would be unable to perform survey 

mission (will lose time which It could have been detecting NEOs)

• Will only be able to characterize NEOs with a Vlim much less than the 

telescope’s Vlim in detect mode

• Increased moving parts & complexity vs. no filter wheel

Wet Mass (kg) 3095

367

1285

P/L

Dry Mass (kg) 2942

S/C

Power (W) 1311

Wet Mass (kg) 3095

367

1285

P/L

Dry Mass (kg) 2942

S/C

Power (W) 1311

* Note: Passive Venus flyby is employed to position S/C in approximately a 0.6 by 0.8 AU orbit. 

Hydrazine used for orbit corrections during cruise-phase of the mission.

289

138

657

38

169

100

266

1657

1285

Dry Mass (kg)

Thermal 10%

TT&C 5%

Structure & Mechanisms 22%

Propulsion 1%

Power 6%

C&DH 3%

ACDS 9%

Spacecraft Bus 56%

Payload 44%

% of Dry Mass

289

138

657

38

169

100

266

1657

1285

Dry Mass (kg)

Thermal 10%

TT&C 5%

Structure & Mechanisms 22%

Propulsion 1%

Power 6%

C&DH 3%

ACDS 9%

Spacecraft Bus 56%

Payload 44%

% of Dry Mass

System Sizing
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C-16: Multiple-channel 0.5m IR L1/L2

Technical Description

• Space-based dual mode (detect & characterize) telescope, pointable

• Infrared (6-10 μm), 1.7° x 6.8° FOV, Vlim= 23.6-25.4 in detect mode

• FOV & Vlim much less in characterize mode

• Modification to solely detection system

• Filter wheel or spectrograph added to provide narrow band pass for 

characterization

• 0.5m aperture

• Passively cooled 

CONOPS

• Location: Earth-Sun L1 (nearly identical detection performance at L2)

• Operational: 2012 (TBR, may be affected by dual mode)

• Availability: Split between NEO detection and characterization modes

• Launch Vehicle: Delta II 2920

• 10 year mission life

Rationale for Consideration

• Add characterization capability with relatively minor modification to an 

asset that may be desired for detection purposes

Capability

• NEO orbital parameters, composition, mass, size

Pros

• If already in use for detection purposes, then may be cost effective design 

for dual (detect & characterize) purpose

• No atmospheric absorbance in IR vs. ground-based asset

Cons

• While asset in characterization mode, it would be unable to perform survey 

mission (will lose time which It could have been detecting NEOs)

• Will only be able to characterize NEOs with a Vlim much less than the 

telescope’s Vlim in detect mode

• Increased moving parts & complexity vs. single channel

System Sizing

23

15

113

0

62

18

109

341

267

Dry Mass (kg)

Thermal 4%

TT&C 2%

Structure & Mechanisms 19%

Propulsion 0%

Power 10%

C&DH 3%

ACDS 18%

Spacecraft Bus 56%

Payload 44%

% of Dry Mass

23

15

113

0

62

18

109

341

267

Dry Mass (kg)

Thermal 4%

TT&C 2%

Structure & Mechanisms 19%

Propulsion 0%

Power 10%

C&DH 3%

ACDS 18%

Spacecraft Bus 56%

Payload 44%

% of Dry Mass

Wet Mass (kg) 608

84

267

P/L

Dry Mass (kg) 608

S/C

Power (W) 302

Wet Mass (kg) 608

84

267

P/L

Dry Mass (kg) 608

S/C

Power (W) 302

C-17: Multiple-channel 0.5m IR Venus-like

Technical Description

• Space-based dual mode (detect & characterize) telescope, pointable

• Infrared (6-10 μm), 1.7° x 6.8° FOV, Vlim= 23.6-25.4 in detect mode

• FOV & Vlim much less in characterize mode

• Modification to detection system

• Filter wheel or spectrograph added to provide narrow band pass for 

characterization

• 0.5m aperture

• Passively cooled

CONOPS

• Location: Heliocentric, Venus-like orbit

• Operational: 2014 (TBR, may be affected by dual mode)

• Availability: Split between NEO detection and characterization modes

• Launch Vehicle: Delta II 2926

• 10 year mission life

Rationale for Consideration

• Add characterization capability with relatively minor modification to an 

asset that may be desired for detection purposes

• Venus orbit has greater Inner Earth Orbit (IEO) object visibility

Capability

• NEO orbital parameters, composition, mass, size

Pros

• If already in use for detection purposes, then may be cost effective design 

for dual (detect & characterize) purpose

• No atmospheric absorbance in IR vs. ground-based asset

Cons

• While asset in characterization mode, it would be unable to perform survey 

mission (will lose time which It could have been detecting NEOs)

• Will only be able to characterize NEOs with a Vlim much less than the 

telescope’s Vlim in detect mode

• Increased moving parts & complexity vs. single channel

Wet Mass (kg) 708

84

267

P/L

Dry Mass (kg) 672

S/C

Power (W) 302

Wet Mass (kg) 708

84

267

P/L

Dry Mass (kg) 672

S/C

Power (W) 302

* Note: Passive Venus flyby is employed to position S/C in approximately a 0.6 by 0.8 AU orbit. 

Hydrazine used for orbit corrections during cruise-phase of the mission.

69

20

127

9

46

25

109

406

267

Dry Mass (kg)

Thermal 10%

TT&C 3%

Structure & Mechanisms 19%

Propulsion 1%

Power 7%

C&DH 4%

ACDS 16%

Spacecraft Bus 60%

Payload 40%

% of Dry Mass

69

20

127

9

46

25

109

406

267

Dry Mass (kg)

Thermal 10%

TT&C 3%

Structure & Mechanisms 19%

Propulsion 1%

Power 7%

C&DH 4%

ACDS 16%

Spacecraft Bus 60%

Payload 40%

% of Dry Mass

System Sizing
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C-18: Multiple-channel 1m IR L1/L2

Technical Description

• Space-based dual mode (detect & characterize) telescope, pointable

• Infrared (6-10 μm), 1.5° x 6.0° FOV, Vlim= 24.2-26.0 in detect mode

• FOV & Vlim much less in characterize mode

• Modification to solely detection system

• Filter wheel or spectrograph added to provide narrow band pass for 

characterization

• 1m aperture

• Passively cooled 

CONOPS

• Location: Earth-Sun L1 (nearly identical detection performance at L2)

• Operational: 2014 (TBR, may be affected by dual mode)

• Availability: Split between NEO detection and characterization modes

• Launch Vehicle: Delta II 2925

• 10 year mission life

Rationale for Consideration

• Add characterization capability with relatively minor modification to an 

asset that may be desired for detection purposes

Capability

• NEO orbital parameters, composition, mass, size

Pros

• If already in use for detection purposes, then may be cost effective design 

for dual (detect & characterize) purpose

• No atmospheric absorbance in IR vs. ground-based asset

Cons

• Reduced FOV due to bigger aperture for same telescope length

• While asset in characterization mode, it would be unable to perform survey 

mission (will lose time which It could have been detecting NEOs)

• Will only be able to characterize NEOs with a Vlim much less than the 

telescope’s Vlim in detect mode

• Increased moving parts & complexity vs. single channel

System Sizing

41

27

203

0

111

32

195

609

442

Dry Mass (kg)

Thermal 4%

TT&C 3%

Structure & Mechanisms 19%

Propulsion 0%

Power 11%

C&DH 3%

ACDS 19%

Spacecraft Bus 58%

Payload 42%

% of Dry Mass

41

27

203

0

111

32

195

609

442

Dry Mass (kg)

Thermal 4%

TT&C 3%

Structure & Mechanisms 19%

Propulsion 0%

Power 11%

C&DH 3%

ACDS 19%

Spacecraft Bus 58%

Payload 42%

% of Dry Mass

Wet Mass (kg) 1051

143

442

P/L

Dry Mass (kg) 1051

S/C

Power (W) 449

Wet Mass (kg) 1051

143

442

P/L

Dry Mass (kg) 1051

S/C

Power (W) 449

C-19: Multiple-channel 1m IR Venus-like

Technical Description

• Space-based dual mode (detect & characterize) telescope, pointable

• Infrared (6-10 μm), 1.5° x 6.0° FOV, Vlim= 24.2-26.0 in detect mode

• FOV & Vlim much less in characterize mode

• Modification to detection system

• Filter wheel or spectrograph added to provide narrow band pass for 

characterization

• Passively cooled

• 1m aperture

CONOPS

• Location: Heliocentric, Venus-like orbit

• Operational: 2014 (TBR, may be affected by dual mode)

• Availability: Split between NEO detection and characterization modes

• Launch Vehicle: Delta IV Medium

• 10 year mission life

Rationale for Consideration

• Add characterization capability with relatively minor modification to an 

asset that may be desired for detection purposes

• Venus orbit has greater Inner Earth Orbit (IEO) object visibility

Capability

• NEO orbital parameters, composition, mass, size

Pros

• If already in use for detection purposes, then may be cost effective design 

for dual (detect & characterize) purpose

• No atmospheric absorbance in IR vs. ground-based asset

Cons

• While asset in characterization mode, it would be unable to perform survey 

mission (will lose time which It could have been detecting NEOs)

• Will only be able to characterize NEOs with a Vlim much less than the 

telescope’s Vlim in detect mode

• Increased moving parts & complexity vs. single channel

Wet Mass (kg) 1229

143

442

P/L

Dry Mass (kg) 1168

S/C

Power (W) 449

Wet Mass (kg) 1229

143

442

P/L

Dry Mass (kg) 1168

S/C

Power (W) 449

* Note: Passive Venus flyby is employed to position S/C in approximately a 0.6 by 0.8 AU orbit. 

Hydrazine used for orbit corrections during cruise-phase of the mission.

124

36

228

15

83

44

195

726

442

Dry Mass (kg)

Thermal 11%

TT&C 3%

Structure & Mechanisms 20%

Propulsion 1%

Power 7%

C&DH 4%

ACDS 17%

Spacecraft Bus 62%

Payload 38%

% of Dry Mass

124

36

228

15

83

44

195

726

442

Dry Mass (kg)

Thermal 11%

TT&C 3%

Structure & Mechanisms 20%

Propulsion 1%

Power 7%

C&DH 4%

ACDS 17%

Spacecraft Bus 62%

Payload 38%

% of Dry Mass

System Sizing
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2006 Near-Earth Object Survey and Deflection Study 

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F.1.8. In-Situ Characterization Concept Sheets

Technical Description

• Spacecraft with tracking mirror system, narrow and wide field-of-view 

imager, visible/infrared imaging spectrograph, magnetometer, laser / 

RADAR altimeter, gamma ray and neutron spectrometer, near infrared 

spectrometer, and multispectral imager 

• 5 year design life

• 125m/s delta-V with 225s Isp hydrazine

• 4600m/s delta-V with 3000s Isp Xenon

• Optional STAR 37F SRM would provide up to 2467m/s delta-V

CONOPS

• Multi-body flyby tour, visiting as many NEOs as possible for 

characterization

• Launch vehicle: Delta II 2925 or EELV Medium

Rational for Consideration

• Can visit and characterize several NEOs with single mission

• Electric propulsion may allow for an increased number of NEO flybys

Capability

• Characterize general structure, size, shape, improved mass estimate, 

binarity, and rotation of NEOs

Pros

• Multiple NEOs can be targeted

• Build on heritage spacecraft (Deep Space 1)

Cons

• Complicated mission design and long flight times with electric propulsion

• Limited encounter duration vs. an orbiter

C-21: Flyby Tour (Electric)

Optional STAR 37F SRM (kg) 1288

Payload Power (W) 70

Spacecraft Dry Mass (kg) 440

Propellant – Xenon (kg) 80

Propellant – Hydrazine (kg) 30

Spacecraft Wet Mass (kg) 550

BOL Power (W) 2500

380

60

Spacecraft Bus Mass (kg)

Payload (kg)

Optional STAR 37F SRM (kg) 1288

Payload Power (W) 70

Spacecraft Dry Mass (kg) 440

Propellant – Xenon (kg) 80

Propellant – Hydrazine (kg) 30

Spacecraft Wet Mass (kg) 550

BOL Power (W) 2500

380

60

Spacecraft Bus Mass (kg)

Payload (kg)

System Sizing

Orbit Options Flyby Spacecraft

Technical Description

• Spacecraft with tracking mirror system, narrow and wide field-of-view 

imager, visible/infrared imaging spectrograph, magnetometer, laser / 

RADAR altimeter, gamma ray and neutron spectrometer, near infrared 

spectrometer, and multispectral imager

• 5 year design life

• 436m/s delta-V with 225s Isp

• Optional STAR 37F SRM would provide up to 2638m/s delta-V

CONOPS

• Multi-body flyby tour, visiting as many NEOs as possible for 

characterization

• Design and approach similar to the CONTOUR mission

• Launch vehicle: Delta II 2925 or EELV Medium

Rational for Consideration

• Can visit and characterize several NEOs with single mission

• Simpler mission design than electric propulsion

Capability

• Characterize general structure, size, shape, improved mass estimate, 

binarity, and rotation of NEOs

Pros

• Multiple NEOs can be targeted

• Build on heritage spacecraft (CONTOUR)

• Simpler mission design and shorter flight times than electric propulsion

Cons

• Limited encounter duration vs. an orbiter

C-22: Flyby Tour (Chemical)

Optional STAR 37F SRM (kg) 1288

Payload Power (W) 70

Spacecraft Bus Mass (kg) 330

BOL Power (W) 697

475

390

60

Spacecraft Wet Mass (kg)

Spacecraft Dry Mass (kg)

Payload Mass (kg)

Optional STAR 37F SRM (kg) 1288

Payload Power (W) 70

Spacecraft Bus Mass (kg) 330

BOL Power (W) 697

475

390

60

Spacecraft Wet Mass (kg)

Spacecraft Dry Mass (kg)

Payload Mass (kg)

System Sizing
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2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 193 

Technical Description

• Flyby –

• Payload same as Flyby In Situ Characterization Alternative

• 2 year design life

• 203m/s delta-V with 225s Isp

• Optional STAR 37F SRM would provide up to 1721m/s delta-V

• Impactor (1 used) –

• High precision star camera (Impactor Target Sensor, ITS), and 

algorithms developed for DS-1, small hydrazine propulsion for 

attitude correction (25m/s, 1750Ns impulse)

• Delivers 19 Gigajoules of KE (370kg at 10.2 km/s)

CONOPS

• Hit NEO with impactor and characterize debris with flyby spacecraft

• Launch vehicle: Delta II 2925 or EELV Medium

Rational for Consideration

• Can provide greater characterization than a only flyby mission

• Simpler mission design than an orbiter or flyby tour

Capability

• Characterize NEO orbital parameters, composition, improved mass

estimate, size, binarity, rotation, shape, surface topography, & internal 

properties

Pros

• Build on heritage spacecraft (Deep Impact)

• Additional data from impactor ejecta

Cons

• Accurate targeting from flyby spacecraft needed

• Only a single NEO can be targeted

• Impactor can only target one location on the NEO

• Limited encounter duration vs. an orbiter

C-23: Flyby + Impactor

Flyby Spacecraft Impactor

Optional STAR 37F SRM (kg) 1288

Payload Power (W) 80

Impactor Flyby

2.8 kW-hr for 

24 hours

377

370

185

185

Spacecraft Bus Mass (kg) 530

BOL Power (W) 727

1060

967

377 (impactor), 

60 (Instruments)

Wet Mass (kg)

Spacecraft Dry Mass (kg)

Payload (kg)

Optional STAR 37F SRM (kg) 1288

Payload Power (W) 80

Impactor Flyby

2.8 kW-hr for 

24 hours

377

370

185

185

Spacecraft Bus Mass (kg) 530

BOL Power (W) 727

1060

967

377 (impactor), 

60 (Instruments)

Wet Mass (kg)

Spacecraft Dry Mass (kg)

Payload (kg)

System Sizing

Technical Description

• Spacecraft with tracking mirror system, narrow and wide field-of-view 

imager, visible/infrared imaging spectrograph, magnetometer, laser / 

RADAR altimeter, gamma ray and neutron spectrometer, near infrared 

spectrometer, and multispectral imager 

• 5 year design life

• 1597m/s delta-V with 315s Isp

• Optional STAR 37F SRM would provide up to 1948m/s delta-V

CONOPS

• Characterize NEO by orbiting it

• Launch vehicle: Delta II 2925 or EELV Medium

Rational for Consideration

• Can provide detailed characterization of a NEO for a long period

Capability

• Characterize NEO orbital parameters, composition, mass, size, 

binarity, rotation, shape, and surface topography

Pros

• Build on heritage spacecraft (NEAR)

• Can directly measure NEO mass

Cons

• Only a single NEO can be targeted

• Complex mission design due to orbiting requirement

• Can only perform remote characterization vs. a lander or penetrator

C-24: Orbiter

Optional STAR 37F SRM (kg) 1288

Payload Power (W) 70

Spacecraft Bus Mass (kg) 454

BOL Power (W) 697

862

514

60

Spacecraft Wet Mass (kg)

Spacecraft Dry Mass (kg)

Payload Mass (kg)

Optional STAR 37F SRM (kg) 1288

Payload Power (W) 70

Spacecraft Bus Mass (kg) 454

BOL Power (W) 697

862

514

60

Spacecraft Wet Mass (kg)

Spacecraft Dry Mass (kg)

Payload Mass (kg)

System Sizing
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194 DRAFT Pre-Decisional Material

C-25: Orbiter + Lander

Optional STAR 37F SRM (kg) 1288

Payload Power (W) 25 80

133

120

100

90

10

Lander Orbiter

Spacecraft Bus Mass (kg) 653

BOL Spacecraft Power (W) 727

1281

1073

420 (60kg + 

3 landers)

Spacecraft Wet Mass (kg)

Spacecraft Dry Mass (kg)

Payload Mass (kg)

Optional STAR 37F SRM (kg) 1288

Payload Power (W) 25 80

133

120

100

90

10

Lander Orbiter

Spacecraft Bus Mass (kg) 653

BOL Spacecraft Power (W) 727

1281

1073

420 (60kg + 

3 landers)

Spacecraft Wet Mass (kg)

Spacecraft Dry Mass (kg)

Payload Mass (kg)

System Sizing

Technical Description

• Orbiter –

• Payload same as Orbiter In Situ Characterization Alternative

• 5 year design life

• 391m/s delta-V with 225s Isp

• Optional STAR 37F SRM would provide up to 1525m/s delta-V

• Lander (3 used) –

• Small lander deployed as black box payload from orbiter

• Payload includes optical camera, Raman spectrometer, 

microscope, alpha proton x-ray detector, lidar, and penetrometer

• 2 year design life

CONOPS

• Characterize NEO by orbiting and landing on it

• Launch vehicle: Delta II 2925 or EELV Medium

• Lander can be launched with a SRM from a small launch vehicle

Rational for Consideration

• Can provide detailed characterization of a NEO for a long period

• Lander can provide in-situ characteristics of a NEO from its surface

Capability

• Characterize NEO orbital parameters, composition, mass, size, shape, 

binarity, rotation, surface topography, and sub-surface internal properties

Pros

• Can directly measure NEO mass

• Multiple landers can target several locations on the NEO

• Large amount of additional data from lander

Cons

• Only a single NEO can be targeted

• Complex mission design due to orbiting requirement

• May be very complicated/impossible to land on fast rotating asteroid

• Lander may require active anchoring to counteract any moving parts

C-26: Orbiter + Penetrator

Spacecraft with 

Multiple Penetrators Single Penetrator

Technical Description

• Orbiter –

• Payload same as Orbiter In Situ Characterization Alternative 

• 5 year design life

• 391m/s delta-V with 225s Isp

• Optional STAR 37F SRM would provide up to 1488m/s delta-V

• Penetrator (3 used) –

• Miniature dart-like spacecraft with communications, command 

and data handling, navigation, power, thermal, and propulsion 

subsystems with a neutron and mass spectrometer and camera

• 12-hour mission life

CONOPS

• Characterize NEO by orbiting and firing multiple penetrators into target

• Launch vehicle: EELV Medium

Rational for Consideration

• Can provide detailed characterization of a NEO for a long period

• Penetrator can provide in-situ characteristics of a NEO from its surface

• Penetrator design is simpler & more robust than a lander

Capability

• Characterize NEO orbital parameters, composition, mass, size, shape, 

binarity, rotation, surface topography, and sub-surface internal properties

Pros

• Can directly measure NEO mass

• Multiple penetrators can target several locations on the NEO

• Penetrators can be used in fast rotating asteroid

Cons

• Only a single NEO can be targeted

• Short lifetime limits amount of additional data from penetrator

• Complex mission design due to orbiting requirement Optional STAR 37F SRM (kg) 1288

Payload Power (W) 80

200 W-Hr

133

58

51

7

Penetrator Orbiter

Spacecraft Bus Mass (kg) 653

BOL Spacecraft Power (W) 727

1329

1113

460 (60kg inst. 

P/L + 3 

Penetrators)

Spacecraft Wet Mass (kg)

Spacecraft Dry Mass (kg)

Payload Mass (kg)

Optional STAR 37F SRM (kg) 1288

Payload Power (W) 80

200 W-Hr

133

58

51

7

Penetrator Orbiter

Spacecraft Bus Mass (kg) 653

BOL Spacecraft Power (W) 727

1329

1113

460 (60kg inst. 

P/L + 3 

Penetrators)

Spacecraft Wet Mass (kg)

Spacecraft Dry Mass (kg)

Payload Mass (kg)

System Sizing
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2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 195 

Appendix G. Description of Deflection Alternatives 

G.1. Deflection Alternatives Using Impulsive Techniques 

Impulsive techniques act instantaneously on the object and include kinetic impactors, 

which collide into the object at a high speed, and explosive techniques, which explode 

either on a high-velocity impact (similar to Deep Impact) or detonate during a highvelocity flyby or after they have landed on the object. 

G.1.1. Conventional Explosive – Contact

A surface-level detonation of a conventional explosive potentially would add impulse to a 

kinetic impactor as a consequence of the ejection of near-surface material. A spacecraft 

on an impact trajectory with a PHO could deliver such a device and an impact fuse would 

detonate the explosive on contact. The mission profile and targeting would be similar to 

those used by a kinetic impactor. 

As with a subsurface explosive, the structure and density of the asteroid is very important 

when trying to calculate the amount of conventional explosives needed to produce the 

desired change in velocity; however, a blast on the surface transfers roughly half the 

energy of a subsurface explosion. [34] This means the mission would have to double the 

amount of explosives to get the same effect. This method uses the same assumptions as 

the subsurface methods. Without knowing the asteroid’s inner structure, predicting the 

explosive’s cratering effects would be difficult. 

For impact velocities of 1 km/s, an explosive may double the momentum transferred to 

the PHO, but research has found that for impact velocities between 2 and 3 km/s, the 

explosive device no longer is beneficial and might actually decrease the overall 

effectiveness. The gaseous products produced by the chemical detonation could prevent 

ejecta from leaving the surface at high velocities. [53] As a result, a simple kinetic 

impactor is preferred for higher-impact velocities. 

Expected Performance

The team calculated the amount of chemical explosives that the Delta IV Heavy and the 

proposed Ares V could carry. For example, for C3 of 25 km2

/s2, the nominal effective 

momentum changes for contact explosives were calculated to be about 106

 kg-m/s and 

107

 kg-m/s, respectively. 

These calculations suggest that the ΔV potential for contact chemical explosives ranges 

from 10 to 100 cm/s for very small PHOs (smaller than 40 meters), to 1 to 10 cm/s for 

small PHOs (smaller than 100 meters), to 0.1 to 1 cm/s for moderately size PHOs 

(smaller than 200 meters). The calculations also suggest that, with few exceptions, 

contact chemical explosives could not be expected to provide sufficient deflections of 

larger PHOs. 
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196 DRAFT Pre-Decisional Material

Characterization Required

Overall effectiveness is defined by the surface and subsurface material displaced in 

response to the explosion (as noted, the efficiency will be degraded as the impact velocity 

increases). Consequently, good information about the body’s surface, internal structure, 

and density are very helpful for predicting the outcome. Since the device would be 

required to impact at a reasonably high velocity (several km a second), the spin of the 

PHO is not an issue, but knowing the shape and orientation of the surface at impact could 

be helpful because the blast wave might remove material from a sloping surface. While 

mission planners could design a deflection mission and spacecraft without detailed 

information about the target PHO, these factors should be examined further to eliminate 

uncertainties. It may be more effective to include several spacecraft with impact 

explosives in a single deflection mission. 

Technology Readiness

Conventional explosives have been thoroughly tested on Earth, providing a wealth of 

knowledge on how different materials react to an explosion. The mechanism of action 

requires that the explosion remove material from near the surface at a high velocity. The 

effectiveness of such a technique is somewhat hypothetical and untested. 

Explosive Devices — The characteristics of explosive devices are well known on Earth. 

Though no one has flown large explosive devices in space for some time, pyrotechnic 

devices have been used extensively on satellites and deep space probes. This technology 

area is rated as having a relatively high level of technology readiness. 

Delivery Systems — The Deep Impact mission demonstrated a delivery system that could 

be used to deliver an explosive payload to a PHO. The risk level is rated as having a 

relatively high level of technology readiness. 

Fusing — Fusing of conventional explosive devices is well understood, but its 

characteristics for high-impact velocities (on the order of 10 km/s) need to be verified. 

This technology area is rated as having a relatively medium level of technology 

readiness. 

Overall Technology Readiness and Effectiveness

Conventional explosives on Earth are low-risk technology. Fusing, detonating, and 

developing shape charges are all well understood. However, these technologies have not 

been demonstrated or used in space, which adds an element of risk. As the impact 

velocity increases, the additional momentum gained by the explosive (as compared with 

the kinetic impact alternative) becomes a small percentage of the total momentum 

delivered. However, given the relatively well-developed status of conventional explosive 

technology, the technique is rated as having a high overall level of technology readiness. 

Compared with other options, contact chemical explosives can provide only a moderate 

amount of momentum change to a PHO. However, most PHOs are relatively small. For 

this reason, the study judges the overall effectiveness of this approach as medium. 
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2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 197 

G.1.2. Conventional Explosives - Subsurface

There are two possible approaches for carrying out a subsurface explosion of 

conventional explosives. The first would require rendezvousing with and landing on the 

PHO, implanting the device, and timing the detonation to assure that the ejecta from the 

blast would be in the correct direction and at the optimal point in the object’s orbit. [55] 

This option requires extensive operations near or when attached to the PHO. It also 

requires technologies and detailed characterization methods that currently are unavailable. 

The second option involves containing the explosive inside an impacting spacecraft. The 

explosive would detonate some time after the spacecraft came into contact with the 

object’s surface, allowing it to penetrate some distance into the surface. This would be 

similar to a “bunker-buster” aerial bomb. The approach for delivering the explosive is 

similar to the one used by a kinetic impactor and would not require a rendezvous. The 

additional weight and characterization information needed to design a reliable penetrator 

is comparable to what is needed to develop a similarly sized contact explosive. 

This assessment does not account for any momentum transfer from ejecta caused by the 

detonation. As is discussed in the kinetic impact alternative, cratering can impart a 

significant momentum change, but is beyond the scope of this analysis. However, since 

the energy density of a conventional explosive is relatively small, the ability to create 

ejecta is expected to be small. Therefore, the overall assessment of its effectiveness is 

expected to be valid. The many possible surface structures, densities, and material 

properties may have a profound effect on mission results. These are among the 

uncertainties that mission planners would need to verify through testing and take into 

account when designing a spacecraft and mission. 

Expected Performance

The team calculated the amount of chemical explosives needed to carry out “bunkerbuster” approach, basing the estimate on the payload capacities of the Delta IV Heavy 

and the proposed Ares V. For example, for C3 of 25 km2

/s2, the nominal effective 

momentum changes for contact explosives were calculated to be somewhat greater than 

that of the contact chemical explosives case: 3x 106

 kg-m/s and 3x 107 kg-m/s, 

respectively. 

These calculations suggest that the ΔV potential for subsurface chemical explosives 

ranges from 10 to 100 cm/s for very small PHOs (smaller than 50 meters), to 1 to 10 cm/s 

for small PHOs (smaller than 150 meters), to 0.1 to 1 cm/s moderate PHOs (smaller than 

300 meters). The calculations also suggest that, with few exceptions, subsurface chemical 

explosives could not be expected to sufficiently deflect larger PHOs. 

Characterization Required

The mass of the PHO and the ΔV required are the most important parameters for defining 

the overall deflection mission. However, density and subsurface characteristics also 

might help better refine a deflection campaign using this technique because it relies on 

the proper functioning of a subsurface device. Specific surface properties would not be 

required as long as the explosive device can imbed itself to the desired depth over a wide 

range of impact conditions and materials. 
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2006 Near-Earth Object Survey and Deflection Study 

198 DRAFT Pre-Decisional Material

Technology Readiness

Explosive Devices — Characteristics of explosive devices are well known on Earth. 

Though no large explosive devices have flown in space for some time, pyrotechnic 

devices have been used extensively on satellites and deep space probes. This technology 

area is rated as having a relatively high level of technology readiness. 

Delivery Systems/Implantation — The Deep Impact mission demonstrated a delivery 

system that mission planners could use to deliver an explosive payload to a PHO, but the 

ability to implant an explosive either by kinetic energy from the impact (high-relative 

velocity with no rendezvous) or by mechanical means after a soft landing (requires 

rendezvous) have not been verified. The complexity of this approach may be problematic. 

This technology area is rated as having a relatively low level of technology readiness. 

Fusing — Fusing a conventional explosive device is well understood. However, fusing 

for these high relative velocities and the allowances that must be made for unknown PHO 

material properties must be verified. This technology area is rated as having a relatively 

medium level of technology readiness. 

Overall Technology Readiness and Effectiveness

This alternative is a higher risk than its surface counterpart mainly because planners 

would need to understand the object’s surface structure and internal composition to 

effectively design a mission. It is rated as having a medium level of overall technology 

readiness. 

Compared with other options, subsurface chemical explosives can provide only a 

moderate amount of momentum change to a PHO. However, most PHOs are relatively 

small. For this reason, the overall effectiveness of this approach is judged to be medium. 

G.1.3. Kinetic Impactor

A kinetic impactor mission would target a mass so that it impacts in a particular direction, 

similar to the approach taken with the Deep Impact mission. The objective would be to 

impact the impactor at a sufficiently high-impact velocity that the correct momentum is 

transferred to the PHO. The relative velocity at impact would be determined by the 

specific intercept trajectory. High-efficiency propulsion systems such as solar sails could 

be used to increase delivered impact mass and increase the impact velocity [56], but with 

a higher technology risk and much longer launch-to-deflection timeline. 

Kinetic impactors greatly depend on the asteroid’s structure. Generally, momentum 

transfer depends on the mechanics of the impact. Whether the object is solid metal or a 

loose pile of particles will determine the efficiency of the impact. For greatest momentum 

transfer, the impact should cause a large amount of cratering and expulsion of material. 

For the kinetic impactor, the transfer of momentum is calculated by 

( )

M asteroid

V

mimpactuimpact Δ = β (2)

where β is the momentum multiplication constant (momentum exchange efficiency) and 

ΔV represents the velocity change imparted to the object. If β = 1, the collision is 
DRAFT Pre-Decisional Material

2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 199 

perfectly plastic, no ejecta is produced, and momentum is imparted directly. If the PHO is 

sufficiently weak and the impactor penetrates, then β<1 and the impact is less effective. If 

β>1, ejecta is released by the impact, and the impact is more effective. Some estimate 

that β could have a magnitude of 10 or more, depending on the material properties of the 

PHO, but a correlation of measurable asteroid characteristics and β is difficult. 

The required mass is obtained as 

⋅β

Δ ⋅ =

relative

deflection asteroid

impact v

V M

m (3)

and the momentum transferred by the impact is 

Δp = mimpact ⋅vrelative ⋅ β (4)

Thus, by knowing the mass of the PHO, the desired velocity increment, and the relative 

velocity of the impactor at impact, one can estimate the momentum imparted to the PHO 

and the impactor’s required mass to effectively carry out the mission. β is expected to 

vary from 0.1 to 100 with values between 1 and 10 used for analyses in this study. 

Expected Performance

Calculations were carried out for hypothetical kinetic impactors based on various impact 

velocities, various values for β, and based on the payload capacities of the Delta IV 

Heavy and the proposed Ares V. For example, for C3 of 25 km2

/s2, the nominal effective 

momentum changes for kinetic impactors were calculated to range from about 5 x 107

 to 

2 x 109

 kg-m/s for systems launched by the Delta IV Heavy and 4 x 108 to 2 x 1010 kgm/s for the Ares V. 

These calculations suggest that kinetic impactors could be deployed to deflect moderateto-large PHOs, perhaps even as large as a few km, if the ΔV requirement is on the order 

of 1 mm/s. Such small ΔV requirements are probably unusual cases, corresponding to 

circumstances such as keyhole encounters or cases where the deflection attempt is 

decades in advance of the potential collision with Earth. 

More typically, the calculations for kinetic impactors suggest that they could reasonably 

be deployed for cases where ΔV requirements might range from 1 to 10 cm/s for 

moderately large PHOs (smaller than several hundred meters in size). 

Characterization Required

For the kinetic impact mission, information on the PHO spin rate and shape are not 

required, though the location of the center of mass would be helpful. The mission and 

impactor can be designed using assumed size and mass estimates and assumed material 

properties, although definitive estimates would be helpful. The primary uncertainty 

related to kinetic impactors is the momentum multiplication factor, but the above 

approach is conservative in that it does not account for any ejecta or enhanced outgassing 

(as seen after the Deep Impact intercept of comet Tempel 1), which could actually 

provide much more effective ΔV than anticipated. As Deep Impact illustrates, a kinetic 

impact mission can be designed without detailed information on the target PHO, but 
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2006 Near-Earth Object Survey and Deflection Study 

200 DRAFT Pre-Decisional Material

again, a deflection mission must take into account uncertainties in the multiplication 

factor and other parameters. 

M-4: Kinetic Impactor

Spacecraft Wet Mass (kg) 2640

BOL Power (W) 1000

2134

520

1614 (14 targeting, 

1600 copper ballast)

Spacecraft Dry Mass (kg)

Spacecraft Bus Mass (kg)

Payload Mass (kg)

Spacecraft Wet Mass (kg) 2640

BOL Power (W) 1000

2134

520

1614 (14 targeting, 

1600 copper ballast)

Spacecraft Dry Mass (kg)

Spacecraft Bus Mass (kg)

Payload Mass (kg)

System Sizing

Technical Description

• A single spacecraft that is Near Earth Object Deflection System

(NEODESYS) like

• Designed to impart a 4.6mm/s delta-V on the asteroid Apophis

• Propulsion: 362m/s, hydrazine (3kg attitude control, 503kg trajectory 

corrections)

CONOPS

• Imparts a momentum change on the NEO by impact

• Provides another spacecraft to provide assessment of impact

• Launch vehicle: EELV Medium

Rational for Consideration

• Non-nuclear option – more politically feasible

• Simple, flexible design

Capability

• Imparts a momentum change on the NEO to prevent collision with Earth

Pros

• Fewer political concerns compared to other mitigation alternatives

• Can be performed in a single mission or salvos

Cons

• Large mass or multiple salvos needed for large asteroids 

• Requires more complex targeting

• Less/not effective on rubble piles and binary NEOs

Kinetic impactors do have the potential to fragment an asteroid, particularly if the 

asteroid has a diameter of 1-3 km. Kinetic-energy impact deflection techniques are 

probably viable for stronger, smaller bodies (diameter of less than 1 km) without concern 

for fragmentation, but planners would need to undertake a disruption analysis for 

asteroids a few km in diameter. For example, a kinetic impactor of sufficient size to 

impart a 1 cm/s ΔV to a 1-km asteroid would likely fragment the body. For targets with 

multiple large masses (e.g. binary objects), missions targeted at each primary mass may 

be required. Fragmentation is discussed further in Section 6.14. 

Technology Readiness

Key technologies and systems related to kinetic impactors have been tested in space (e.g., 

Deep Impact), but their effectiveness for PHO deflection has not been characterized. 

Risks associated with critical technologies include: 

Delivery Systems — Delivery systems required to deliver the impactor to the vicinity of a 

NEO have been demonstrated by the Deep Impact mission, which allowed no maneuvers 

during the last 6 minutes before impact, almost the last 4,000 km. It was still successful at 

hitting its target. This technology area is rated as having a relatively high level of 

technology readiness. 

Targeting — The Deep Impact impactor was successfully targeted for a 10 km/s impact. 

[58] Some additional work may be necessary to verify this approach for smaller objects 
DRAFT Pre-Decisional Material

2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 201 

with shorter acquisition ranges and higher relative velocities. This technology area is 

rated as having a relatively high level of technology readiness. 

Overall Technology Readiness and Effectiveness

Because other Earth-based and space-based missions have verified many of the required 

technologies, the overall technology readiness associated with this technique is 

judged as high. Since it is expected to be applicable to a variety of PHOs and is highly 

scaleable, its effectiveness is rated high. 

G.1.4. Nuclear Explosive – Surface Contact

The mission concept for detonating a nuclear explosive on contact with a PHO is very 

similar to the concept proposed for a conventional explosive. The concept is to place a 

nuclear explosive on an interceptor with an impact fuse. Blasting material from the PHO 

using the energy from the explosive would produce the momentum shift. 

Differences exist between how nuclear and conventional explosives interact with a 

material, making nuclear explosives more efficient than their relative advantage in 

explosive energy. Nuclear explosives do not merely form a crater and eject material, but 

through radiative heating (principally from X-rays and neutrons) broadly ablate the 

surface of the object, transferring additional momentum. 

M-2: Nuclear Surface

Technical Description

• Carrier –

• Communication relay for confirmation of detonation

• 2 year design life

• Interceptor –

• Visible / Near IR and LIDAR navigation systems for terminal 

guidance

• Neutron optimized device (1.0 MT yield)

• 2 year design life 

• 1200m/s delta-V

CONOPS

• Carrier releases interceptor which explodes at surface of NEO 

imparting a momentum change

• Launch vehicle: Delta IV Heavy

Rational for Consideration

• Valid for wide range of asteroid types

• Surface location may provide greater energy transfer than a standoff

Capability

• High energy transfer into NEO causes ablation and ejection of material 

altering the course of the NEO

Pros

• Mature mitigation technology

• Requires a relatively low level of characterization to perform and is 

effective despite NEO binarity, rotation or structure

• Can be performed in a single mission or salvos

Cons

• Difficult political considerations, nuclear detonation ban

• Highly precise timekeeping and ranging for surface detonation

5000

2000

1100

900 (70 targeting, 

830 weapon)

Interceptor Carrier

Spacecraft Bus Mass (kg) 442

BOL Spacecraft Power (W) 401

5900

5442

5000

Spacecraft Wet Mass (kg)

Spacecraft Dry Mass (kg)

Payload Mass (kg)

5000

2000

1100

900 (70 targeting, 

830 weapon)

Interceptor Carrier

Spacecraft Bus Mass (kg) 442

BOL Spacecraft Power (W) 401

5900

5442

5000

Spacecraft Wet Mass (kg)

Spacecraft Dry Mass (kg)

Payload Mass (kg)

System Sizing

For a surface burst, a nuclear explosive delivers approximately 8% of its energy to the 

object. [59] Theoretical calculations predict that a surface burst of a nuclear explosive is 

roughly 3-5% as effective as that of a conventional explosive, but this shortcoming is 

overcome by the much higher yield per kilogram for nuclear explosives over that for 

conventional explosives [60]. Thus, the yield of a contact nuclear explosive required to 
DRAFT Pre-Decisional Material

2006 Near-Earth Object Survey and Deflection Study 

202 DRAFT Pre-Decisional Material

deliver a specified impulse can be conservatively calculated by dividing the yield of the 

conventional explosives required by 0.03. 

Expected Performance

Calculations were carried out for hypothetical contact nuclear explosives based on the 

payload capacities for the Delta IV Heavy and the proposed Ares V. For example, for C3 

of 25 km2

/s2, the nominal effective momentum changes for contact explosives were 

calculated to be about 2 x 1011 kg-m/s and 2 x 1012 kg-m/s, respectively. 

These calculations suggest that the ΔV potential for contact nuclear explosives ranges 

from 10 to 100 cm/s for km-scale PHOs, to 0.1 to 1 cm/s for PHOs that are 10 km or 

larger. Surface nuclear explosives, therefore, could be expected to deflect most PHOs, 

given sufficient lead times (in some cases, decades to a century in advance). It is likely 

that, at the end of the Survey program, typical lead times for very large PHOs will 

probably be one or more centuries. 

Characterization Required

A mission utilizing detonation of nuclear explosive on contact with a target PHO can be 

designed without detailed knowledge of the PHO. However, nuclear surface detonations 

have the potential to fracture and fragment the PHO, and some information on the 

composition and integrity of the PHO would be required to address this concern. For 

purposes of determining which characteristics are required to design the mission, it is 

assumed that fragmentation is not a problem, even if it does occur. Rotation rate is not 

required for this technique, but the orientation at impact could affect the angle of the 

explosion products, which might lower the efficiency in the required direction. Design of 

the mission should examine this possibility and include sufficient margin in the size of 

the explosive device, if required. 

Technology Readiness

Explosive Devices — Nuclear explosives are well characterized; there is a wealth of 

information about the effects of nuclear weapons on materials that may be found on 

NEOs (e.g., sand, granite, basalt, ice, and tundra or “dirty ice”). [62] Nuclear explosives 

have been tested at an altitude of 400 km in space [63] and have been designed for 

delivery on missiles. A wide variety of designs have been tested and, while high-yield 

devices have been removed from the world’s nuclear-weapon stockpile, discussions with 

the staff at Los Alamos and Lawrence Livermore National Laboratories indicate that if 

specialized or high-yield explosives were needed, they could be designed and built to 

operate with no need for testing. This technology area is rated as having a relatively high 

level of technology readiness. 

Delivery Systems — Delivery systems for these devices would be very similar to the one 

used in the Deep Impact mission. This technology area is rated as having a relatively 

high level of technology readiness. 

Targeting and Fusing — Nuclear explosive detonation systems are well known and well 

characterized. A ranging system capable of providing accurate information during the 

final approach is required, and such systems have been demonstrated on Deep Impact and 

other missions. Unlike conventional explosives, a nuclear explosive does not need to 
DRAFT Pre-Decisional Material

2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 203 

come into direct contact with the surface to achieve a reasonable energy coupling. 

Therefore, this mission could use radar-controlled fusing, which has been tested on 

conventional and nuclear bombs for 60 years. The fusing and targeting technology area is 

rated as having a relatively high level of technology readiness. 

Overall Technology Readiness and Effectiveness

This approach is assessed as having a high overall level of technology readiness. 

Because it could be used against any PHO, it is expected to provide a very high level of 

effectiveness. 

G.1.5. Nuclear Explosive — Standoff

Unlike a surface explosion, a nuclear standoff detonation does not use its energy to add 

impulse to the asteroid, but rather to vaporize some of the asteroid’s surface to produce 

the desired change in the NEO’s velocity. [30] [64] By using a standoff detonation, the 

impulse absorbed will theoretically be less than the energy required to break up the 

asteroid, but sufficient to vaporize enough material to impart the necessary ΔV. 

For this mission, the spacecraft would be designed to detonate at a specific height above 

the object’s surface. The radiation produced by the explosion- X-rays, gamma rays, and 

neutrons - would bombard the surface, effectively vaporizing the surface layer. When 

material is vaporized and blown off an asteroid, an impulse is given to the asteroid due to 

conservation of momentum. Using this law and assuming the mass of the ejecta is very 

small compared with the mass of the asteroid, the amount of mass and average velocity of 

the ejecta needed to change the asteroid’s velocity by a certain ΔV can be calculated as 

asteroid

ejecta ejecta

deflection M

m v

ΔV = (5)

where mejecta is the mass of the ejecta, vejecta is the average velocity of the ejecta, Masteroid is 

the mass of the asteroid, and ΔVdeflection is the change in the asteroid’s velocity. Using this 

relationship, one can see that by maximizing either the mass of the ejecta or its velocity, 

one can impart a greater momentum transfer to the asteroid. 
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2006 Near-Earth Object Survey and Deflection Study 

204 DRAFT Pre-Decisional Material

M-1: Nuclear Standoff

Technical Description

• Carrier –

• Communication relay for confirmation of detonation

• 2 year design life

• Interceptor –

• Visible / Near IR and LIDAR navigation systems for terminal 

guidance

• Neutron optimized device (1.9 MT yield)

• 2 year design life 

• 1000m/s delta-V

CONOPS

• Carrier releases interceptor which explodes within close proximity of 

NEO imparting a momentum change

• Launch vehicle: Delta IV Heavy

Rational for Consideration

• Valid for wide range of asteroid types

• Standoff location may provide a better geometry to deflect NEO

Capability

• High energy transfer into NEO causes ablation and ejection of material 

altering the course of the NEO

Pros

• Relatively mature mitigation technology

• Requires a relatively low level of characterization to perform and is 

effective despite NEO binarity, rotation or structure

• Can be performed in a single mission or salvos

Cons

• Difficult political considerations, nuclear detonation ban

• Precise timing required on detonation altitude required

5000

2800

1230

1570 (70 targeting, 

1500 weapon)

Interceptor Carrier

Spacecraft Bus Mass (kg) 442

BOL Spacecraft Power (W) 401

5900

5442

5000

Spacecraft Wet Mass (kg)

Spacecraft Dry Mass (kg)

Payload Mass (kg)

5000

2800

1230

1570 (70 targeting, 

1500 weapon)

Interceptor Carrier

Spacecraft Bus Mass (kg) 442

BOL Spacecraft Power (W) 401

5900

5442

5000

Spacecraft Wet Mass (kg)

Spacecraft Dry Mass (kg)

Payload Mass (kg)

System Sizing

The velocity of the gaseous ejecta is proportional to the square root of the temperature. 

[65] For this reason, it is better to have more mass vaporized to a relatively low 

temperature than little mass vaporized to a much higher pressure. To achieve this desired 

effect, an explosive custom made to emit mostly neutrons would be best for this scenario. 

Such devices have been designed and tested and are discussed in the Appendix P. 

Most fusion-based explosives produce the majority of their radiation as X-rays. Although 

X-rays can carry more energy then neutrons, they are able to penetrate the surface to a 

depth of roughly 10-50 microns depending on surface structure and material. [66] 

Neutron radiation has the ability to penetrate to a depth on the order of 10 cm, effectively 

burning off more mass at a lower temperature and creating a higher-momentum transfer. 

Tailored neutron bombs have the ability to transfer roughly 10% of the blast energy into 

neutrons, which could vaporize the asteroid’s surface. [69] 

To transfer the highest amount of momentum to the NEO, it is important to find the 

detonation height above the PHO’s surface. This allows the most mass to receive the 

most energy and to vaporize and exit with the necessary escape velocity. Reference [30] 

relates the momentum change, energy needed, and the optimum height of the explosion 

above the PHO. 

Achieving the correct detonation height above the surface requires accurate measurement 

and fusing during the encounter. The relative velocity between the spacecraft and asteroid 

may range anywhere from a km per second to 10 km per second near impact. The 

optimum height for a standoff explosion varies widely in literature, from 6-80% of the 

mean asteroid diameter. For this study, the value chosen was 20 meters, or 10% of the 

diameter of the object. A strong correlation exists between standoff distance and 
DRAFT Pre-Decisional Material

2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 205 

effectiveness, and the standoff distance chosen (or realized due to timing errors) could 

affect effectiveness by one-to-two orders of magnitude. 

An advantage of this technique is that it spreads the impulse over a large area of the NEO, 

reducing the possibility of fragmentation. Further, information on its rotation rate and 

other features are not required. 

Expected Performance

Calculations were carried out for standoff nuclear explosives. They were based on 

hypothetical conventional systems and tailored neutron nuclear explosives as well as on 

the payload capacities of the Delta IV Heavy and the proposed Ares V. For example, for 

C3 of 25 km2

/s2, the nominal effective momentum changes for tailored neutron nuclear 

explosives in a standoff mode were calculated to be about 1010 kg-m/s and 5 x 1010 kgm/s, respectively. 

These calculations suggest that the ΔV potential for standoff nuclear explosives ranges 

from 10 to 100 cm/s for PHOs up to about 1 km in size, to 0.1 to 1 cm/s for PHOs that are 

up to several km in size. Standoff nuclear explosives, therefore, could be expected to 

deflect many large PHOs, given sufficient lead times (in some cases, decades to a century 

in advance). 

Characterization Required

Without knowledge of an asteroid’s elemental composition, it is difficult to predict the 

effectiveness of a standoff detonation due to uncertainties in the material’s vaporization. 

These uncertainties must be included in the mission design. Increasing the size of the 

explosive device probably could compensate for these uncertainties. Specific information 

on the size and shape of the target PHO is not required because the device is activated 

above the surface and the device could be sized to counter any size- or shape-related 

uncertainties. Similarly, the device would be sized to compensate for uncertainties in 

surface properties; therefore, specific information is not required. 

An advantage of this technique is that it spreads the impulse over a large area of the PHO, 

reducing the possibility of fragmentation. Further, information on the rotation rate and 

other features are not required, but a single spacecraft can include an explosive device 

sized to deliver sufficient energy to account for these uncertainties. 

Technology Readiness

This approach has fundamentally lower technology requirements than the nuclear surface 

contact alternative. Only the differences are discussed here. 

Explosive Devices — This technology area is rated as having a relatively high level of 

technology readiness. 

Targeting and Fusing — A ranging system capable of providing accurate information 

during final approach is required, and such systems have been demonstrated on Deep 

Impact and other missions. However, as noted earlier, a high-closing velocity is very 

challenging. This mission could use radar-controlled fusing, which has been tested on 

munitions for 60 years. This technology area is rated as having a relatively high level of 

technology readiness. 
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2006 Near-Earth Object Survey and Deflection Study 

206 DRAFT Pre-Decisional Material

Overall Technology Readiness and Effectiveness

The primary issue for standoff nuclear systems relates to timing the explosion so that it 

occurs at the proper altitude above the PHO surface and the effectiveness of the explosion 

in imparting a momentum change. Nuclear explosives deliver the highest energy per unit 

mass to the PHO and the explosive energy can be sized to account for uncertainties. Also, 

the standoff devices will not create large amounts of dust or other debris that will make 

assessing its effectiveness difficult [70]. 

This approach has a high overall level of technology readiness and, because it could be 

used against any PHO, it is expected to provide very high effectiveness. 

G.1.6. Nuclear Explosive - Subsurface

A subsurface nuclear explosive would be delivered to a PHO in a manner similar to a 

kinetic impactor, but the impact velocity would have to be set to the range required to 

achieve the desired penetration. This could require some maneuvering of the interceptor 

before impact. The explosive device would be detonated at the prescribed depth. 

Another approach would be to rendezvous with and land on the NEO. The payload could 

drill a hole to the proper depth and insert the explosive device. The detonation would be 

timed so that the ejecta move in the direction for maximum effect. This would require a 

sensor system capable of determining the orientation of the PHO in space and detonating 

the device when the PHO is oriented correctly. Since this approach requires technologies 

that have not yet been designed or tested, it is not considered further here. 

The effectiveness of subsurface nuclear explosives is very dependent on the internal 

structure of the asteroid. 

Expected Performance

Subsurface nuclear explosives are expected to provide the greatest momentum change per 

system mass of any deflection technique considered in this analysis. Calculations were 

carried out for subsurface nuclear explosives. They were based on hypothetical 

conventional systems, tailored neutron nuclear explosives, and the payload capacities of 

the Delta IV Heavy and the proposed Ares V. For example, for C3 of 25 km2

/s2, the 

nominal effective momentum changes for subsurface nuclear explosives were calculated 

to be about 5 x 1011 kg-m/s and 4 x 1012 kg-m/s. 

These calculations suggest that the ΔV potential for subsurface nuclear explosives ranges 

from 10 to 100 cm/s for PHOs up to several km in size, to 0.1 to 1 cm/s for PHOs that are 

up to tens of km in size. Subsurface nuclear explosives, therefore, could be expected to 

deflect many large PHOs, given sufficient lead times (in some cases, decades to a century 

in advance). 

Characterization Required

To estimate the size of the explosive required, planners would need to know the PHO’s 

mass. They would not necessarily need to know its size, except to estimate the likelihood 

of fragmentation. The density is helpful for designing the penetrator. Uncertainties about 

the PHO’s internal and external structure (e.g., if the impactor struck a steep slope and 
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2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 207 

the ejecta did not move in the optimal direction) could affect the device’s efficiency. 

Increasing the size of the explosive device could compensate for this lack of knowledge. 

Technology Readiness

This approach has fundamentally higher technology requirements than the nuclear 

surface devices. They are discussed here. 

Explosive Devices — This technology area is rated as having a relatively high level of 

technology readiness. 

Delivery Systems – Although explosive penetrators are well understood in the physical 

environment of the Earth, less is known about how explosive penetrators will interact 

with PHOs in space. This technology area is rated as having a relatively low level of 

technology readiness. 

Targeting and Fusing —Nuclear explosive systems are well known and well 

characterized; however, this case requires detonation below the PHO’s surface for which 

there is no precedent. This technology area is rated as having a relatively high level of 

technology readiness. 

Overall Technology Readiness and Effectiveness

The use of nuclear explosives below the surface of a PHO raises issues similar to those 

raised for conventional explosives. In addition, fragmentation of the PHO might be a real 

possibility, and a PHO fracture analysis would be required. Due to the challenges of 

detonating the nuclear explosive beneath the PHO’s surface, this approach is assessed as 

medium in overall technology readiness. In addition, since it could be used against a 

fairly large number of PHOs, but may be more difficult to implement with the smaller 

PHOs, it is assessed as providing medium overall effectiveness. 

G.2. Deflection Alternatives Utilizing Slow Push 

To produce a momentum change on an asteroid, some sort of impulse, whether it is 

instantaneous or over time, must be imparted to the mass. In general, slow push 

techniques have the advantage that they impart very small amounts of force and, as a 

result, do not run the risk of fragmenting the PHO. However, spreading the force over a 

longer arc is less efficient than impulsive techniques, sometimes much less efficient. If 

the general orbital period of a PHO is 2.3±1.1 years [71], slow-push techniques, which 

act over more than a quarter of an orbital period, are expected to have large finite burn 

arc (gravity) losses. The momentum required for these techniques is developed from 

equation (18), which includes a loss for the application of force over a finite orbit arc. 

G.2.1. Enhanced Yarkovsky

Russian civil engineer Ivan Osipovich Yarkovsky first predicted the Yarkovsky Effect in 

1900. Yarkovsky hypothesized that the diurnal heating of a rotating body would create a 

force due to the escape of thermal radiation as the sunlit side rotated into darkness and 

cooled. As the heated surface turns away from the Sun, photons are radiated from the 

asteroid as it cools. Photons carry momentum off the asteroid and, due to Newton’s Law, 

provide a pushing force. Although the force produced by the photons is extremely small, 

over a long period of time it can change an asteroid’s orbit. 
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2006 Near-Earth Object Survey and Deflection Study 

208 DRAFT Pre-Decisional Material

Even though this effect was theorized in 1900, it was not until 2003 that the effect was 

actually proven. That is when a 500-meter wide asteroid named 6489 Golevka, which 

astronomers had tracked since 1991, had taken a 15-km detour from its projected orbit. 

[72] Once proven, the Yarkovsky Effect was suggested as a deflection technique. The 

concept is to enhance the asteroid’s radiation capabilities as a way to increase the force 

experienced by the asteroid, and then let the force slowly change the asteroid’s orbit and 

reduce the threat. 

Although simple in concept, implementation of this method may prove to be difficult. To 

enhance the effect on the asteroid, the asteroid must be covered with paint or other 

material. Even if done properly, the resultant force on the asteroid is extremely small and 

only will be able to produce enough change in momentum to successfully deflect the 

asteroid after years of operation. As discussed in Section 5.9, orbit perturbation may take 

many decades to produce a change that guarantees that an object will not strike Earth. It 

also may take years to measure any effects on the asteroid’s orbit. Finally, the Yarkovsky 

Effect increases as the surface area of the asteroid increases (varies with radius squared), 

while mass and momentum increase as the mass (and total momentum) increases (varies 

with radius cubed). Therefore, larger objects are relatively more difficult than smaller 

objects to divert using this method. 

Asteroid deflection using Yarkovsky is developed further in References [73] and [74]. 

Expected Performance

No calculations were performed to estimate the performance of deflection efforts that 

utilize the Yarkovsky Effect. 

Characterization Required

The Yarkovsky Effect requires a rotating PHO. It also requires changing the PHO’s 

surface so that the surface interacts with solar energy. Over time, the interaction will 

modify the object’s orbit. As a result, designing a deflection mission using this technique 

would require information on the targeted PHO’s spin rate, its mass and size, information 

about its surface properties (to ensure that mission planners select the correct coating 

material), the PHO’s axis of rotation, and its rotation rate. These quantities would be 

required before applying the coating. Uncertainties in the rotation rate, for example, 

might preclude a mission using this technique. 

Technology Readiness

Coating concepts — Although chalk has been suggested, it is unclear what coating types 

would work best in this application and how they would perform in the PHO environment. 

While engineers have used paint coatings to control thermal properties on spacecraft, 

their application to this type of mission faces many unknowns. Since no specific designs 

are presented, this technology area is rated as having a relatively low level of technology 

readiness. 

Coating Delivery — The application of a coating on such a large scale has never been 

attempted before in space. The technologies necessary are largely conceptual. This 

technology area is rated as having a relatively low level of technology readiness. 
DRAFT Pre-Decisional Material

2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 209 

Overall Technology Readiness and Effectiveness

The Yarkovsky Effect is a very small force created by sunlight acting on the surface of a 

NEO. Using this force for deflection would require that a method be developed to deliver 

a coating to the PHO’s surface. This technique could be used only when the impact threat 

is decades or more into the future. 

Neither the spacecraft used to apply the coating, or the coating itself, has been described 

to any level of detail. The effect of this small force will likely depend on the size and 

rotation rate of the PHO to which it is applied, and uncertainties in the object’s size and 

rotation could considerably extend the required action time. The overall technology 

readiness is, therefore, judged to be low. 

In general, the Yarkovsky Effect can work against small- to moderate-size objects, 

particularly in cases where enough time is available, where there is substantial time to 

generate a momentum change, and where small ΔVs are required. For this reason, the 

overall effectiveness of the Yarkovsky Effect is judged to be low. 

G.2.2. Focused Solar

This concept would use a large, thin-film mirror surface to collect sunlight and focus it on 

a small spot on the PHO. Heating would vaporize a small amount of material and the 

material’s departure would impart a small ΔV. The reflector spacecraft would not directly 

contact the PHO; rather it would stand off at a distance from the surface. As a result, its 

operation would not depend on the rotational motion or shape of the PHO, although 

rotation will decrease the efficiency of this approach. 

This concept requires that the spacecraft rendezvous with the PHO and assume a station 

at a fixed distance from the object. At this point, the solar reflector would be erected or 

inflated. Energy would be reflected via a mirror on the spacecraft to a spot on the PHO’s 

surface. The spacecraft must remain precisely positioned near the PHO for as long as the 

system operates. 

Issues related to this technique include the effect of the vaporized material on the energy 

delivered to the NEO; that is, as dust and other material leave the surface, they potentially 

could reflect or absorb some of the energy, preventing it from reaching the surface. 

Potentially, this could require that the spacecraft’s reflecting mirror move its focus spot, 

which would reduce its efficiency because new material then would have to be heated. 

Expected Performance

No calculations were performed to estimate the performance of this deflection technique. 

Characterization Required

The PHO’s mass and surface properties are important for this technique because they 

define the effectiveness of the energy transfer and the amount of time required to impart a 

momentum change. Information about the object’s rotation rate would help because it 

would help assure that dust and ejected material did not dissipate the incoming focused 

beam. No other information is required. Uncertainties in PHO properties would affect the 

time required to complete the deflection. This technique is unlikely to fragment the target 

PHO; so no information on density or other properties is required.
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2006 Near-Earth Object Survey and Deflection Study 

210 DRAFT Pre-Decisional Material

Technology Readiness

Large Space Structures — A large reflector system must be erected near the PHO and 

must maintain its shape and orientation for an extended period. Large structures of this 

type have not been demonstrated. This technology area is rated as having a relatively 

low level of technology readiness. 

Delivery and Proximity Operations, Long Duration Autonomous Operations — This 

approach requires that the spacecraft be delivered to a precise location relative to the 

PHO and that it maintain its position for an extended period. Further, the technique 

requires long-term, autonomous operation of a reasonably sophisticated spacecraft. This 

technology has not been demonstrated. Therefore, this technology area is rated as 

having a relatively medium level of technology readiness. 

Overall Technology Readiness and Effectiveness

Since this system requires a long period of operation, it will require years to assess its 

effectiveness in changing the target’s orbit. Given that the technique requires long-term 

autonomous operation, it is not clear how one would overcome potentially missionending problems. As described earlier, this approach is notional at this point. No space 

missions have verified critical technologies. The overall technology readiness for this 

approach is judged as low.

However, this approach conceivably could be applied to a large percentage of PHOs. For 

this reason, the overall effectiveness of the approach is judged to be medium. 

G.2.3. Gravity Tractor

The concept of a gravity tractor is to impart a momentum change using the force of 

gravity without coming into contact with the PHO. [75] [76] [17] [77] A gravity tractor 

would rendezvous with the object and hold a specific position in space adjacent to the 

PHO for an extended period. The mutual gravitational attraction between the PHO and 

the gravity tractor would slowly “pull” the asteroid to a different trajectory. Thrusters 

would be designed so that the exhaust from the station-keeping motors did not impinge 

on the PHO. 

The momentum transfer of a slow push (or pull) technique depends on two important 

factors: the time over which the thrust is applied and the thrust itself. For this case, the 

thrust applied to the asteroid is simply the gravitational attraction between two bodies. 

Using Newton’s laws of gravitation, the force applied to the bodies can be calculated as: 

2

asteroid tractor

d

G M m F = (6)

where G is the gravitational constant, mtractor is the mass of the tractor, and d is the 

distance between the two masses. This force is effectively the thrust applied to the 

asteroid. Inserting the thrust value into the momentum equation yields: 

)

2

t

) (t t

t

( d

G M m P p

s

s

p

2

asteroid tractor Δ c = − (7)
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2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 211 

This equation uses the very important assumption that the engines used by the gravity 

tractor are oriented so that the exhaust from the engines does not impede the attraction 

between the two bodies. The engines would be canted outward at an angle determined by 

the size of the bodies and the standoff distance. [77] 

The geometry of the canted engines is given in Figure 72. In this case, r is the mean 

radius of the asteroid; d is the distance to the tractor; φ is the cant angle of the thrusters; m

is the mass of the tractor operating with thrust T; M is the mass of the asteroid; and ρ is 

the mean density of the asteroid. 

Figure 72. Geometry of the Gravity Tractor Concept 

As shown, the effectiveness of the gravity tractor varies inversely with the square of the 

distance between the tractor and the PHO, and is therefore very sensitive to the distance 

chosen. While this analysis assumed a standoff distance of 1.5 times the mean asteroid 

radius, other studies [75] have assumed 0.5 times the radius. If the closer hover distance 

is feasible, the effectiveness of the gravity tractor would increase by a factor of about 

nine from that presented in this report. 

Expected Performance

Calculations were carried out for the hypothetical gravity tractor based on the payload 

capacities of the Delta IV Heavy and the proposed Ares V. For example, for C3 of 25 

km2

/s2, the the nominal effective momentum changes for the gravity tractor were 

calculated to be about 5 x 106

 kg-m/s and 6 x 107 kg-m/s, respectively. 

These calculations suggest that the ΔV potential for the gravity tractor is somewhat 

higher than for contact chemical explosives, but less than for the kinetic impactor. The 

calculations estimate the ΔV potential ranges from 10 to 100 cm/s for small PHOs 

(smaller than 90 meters), to 1 to 10 cm/s for moderately small PHOs (smaller than 200 

meters), to 0.1 to 1 cm/s for moderate-size PHOs (smaller than 400-500 meters). This is 

sufficient to consider the gravity tractor for keyhole encounters, including Apophis’s 

close encounter with Earth in 2029 and perhaps 2400 VD17’s potential collision with 

Earth in 2102, provided the deflection were attempted decades before the projected 

impact and assuming a decade of action time (or more). 
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212 DRAFT Pre-Decisional Material

The calculations also suggest that, with few exceptions, the gravity tractor could not be 

expected to provide sufficient deflections for larger PHOs. 

M-24: Gravity Tractor

Spacecraft Module Dry Mass (kg) 12810

Reactor Module Dry Mass (kg) 3177

Spacecraft Bus Mass (kg) 5991

Reactor Output (kWe) 208

Payload Power Non-thrusting (W) 10000

Payload Power Thrusting (W) 3000

Total Delta-V (m/s) 26000

Spacecraft Wet Mass (kg) 28117

16047

60

Spacecraft Dry Mass (kg)

Payload Mass (kg)

Spacecraft Module Dry Mass (kg) 12810

Reactor Module Dry Mass (kg) 3177

Spacecraft Bus Mass (kg) 5991

Reactor Output (kWe) 208

Payload Power Non-thrusting (W) 10000

Payload Power Thrusting (W) 3000

Total Delta-V (m/s) 26000

Spacecraft Wet Mass (kg) 28117

16047

60

Spacecraft Dry Mass (kg)

Payload Mass (kg)

System Sizing

Technical Description

• Nuclear Electric Propulsion (NEP) based spacecraft, JIMO like

• 12 year design life

• Autonomous rendezvous and dock in orbit

• Scientific payload for NEO characterization and ranging included

• Some scaling on mass and reactor size available

CONOPS

• Imparts a momentum change on the NEO by drawing it closer using

mild gravitational effects

• Assumes a 50 year lead time, applicable only for small asteroids

• Launch on two Delta IV Heavy vehicles and assembled in space

Rational for Consideration

• Non-nuclear option – more politically feasible

• Viable for virtually all asteroid types

Capability

• Provides thrust and propellant capacity for gravity tractor to move NEO 

from colliding with Earth

Pros

• Effective on virtually all NEO types (e.g. rubble pile, binary, etc.)

• Slow deflection allows for flexible operational concepts

Cons

• Difficult political considerations, can be used to weaponize NEO

• Larger asteroids would require incrementing mass

• Large lead time required to be effective

• High complexity, cost/risk due to in space assembly

• Must maintain close station adjacent to NEO for extended period

Characterization Required

An accurate estimate of the PHO mass is critical for the gravity tractor. NASA would 

likely use the largest tractor mass possible, and uncertainties in the mass would be 

reflected in the time the tractor must operate in close proximity to the PHO. Uncertainty 

in mass could mean that longer deflection times than available would be required, 

demanding that an alternative deflection approach be used. Rotation rate and shape also 

are needed for defining the operational environment near the PHO, particularly if the 

PHO is oblate and its motion requires the tractor to back away, thereby lowering its 

efficiency. These two parameters also may help estimate the overall efficiency by 

allowing simulations to determine the net gravitational force between the two objects. 

Information on the surface type and material properties are not required. 

Since the force exerted on the PHO is very small, this technique is unlikely to fragment 

the PHO. Further, if the PHO has a moon (some do), it is likely that the moon will simply 

continue its orbit as the tow progresses. Of course, the moon could represent a hazard to 

the towing spacecraft, and any maneuvers required to avoid the hazard would reduce the 

overall efficiency of the technique. 

Technology Readiness

Control Systems — Terrestrial GEO satellites have demonstrated extended stationkeeping operations, though not to the extent needed in this mission. The control systems 
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2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 213 

needed do not seem to exceed the state of the art, however. This technology area is rated 

as having a relatively high level of technology readiness. 

Sustained Power — Because electrical propulsion likely would be used, a stable, longterm source of power would be required. Solar photovoltaics could provide the power 

levels required for the gravity tractor. This technology area is rated as having a relatively 

high level of technology readiness. 

Autonomous Operations — While autonomous operations have been performed during 

several deep space missions, none has attempted to operate so close to another body for 

such a long period of time. This technology area is rated as having a relatively medium 

level of technology readiness. 

Overall Technology Readiness and Effectiveness

The gravity tractor requires a heavy spacecraft to rendezvous with the PHO and to fly 

autonomously in close proximity to the PHO for an extended period. The object’s shape 

and rotation could affect the tractor’s effectiveness to the extent that the tractor must 

adjust its height above the PHO as the PHO surface rotates. In addition, a small force 

imparted to the PHO by the tractor’s stationkeeping thrusters could nullify some of the 

gravitational attraction. 

The thruster and thruster control systems must operate while the tractor is interacting 

with the object and the tractor must carry a power supply capable of powering thrusters 

and onboard systems for extended periods. The small force imparted by the tractor would 

mean that a period of accurate tracking would be required before scientists could verify 

the technique’s effectiveness. 

Considering these factors, the overall technology readiness of the gravity tractor 

concept is judged to be medium. The gravity tractor is assessed to be effective in cases 

where a small total momentum change is required, such as for relatively small PHOs, 

where the time before Earth encounter is large enough to accommodate for uncertainties 

in the PHO mass. It also could be effective for instances involving a keyhole (see Section 

5.2.3). The recent example of a well-publicized asteroid impact scenario that involves a 

keyhole — asteroid Apophis or 2004 MN4 — appears to be the exception rather than the 

general rule. [17] Keyhole scenarios with long lead times and small deflection distances 

have been described as “extremely rare,” or far less than 1% of the total possible impacts. 

[16] Because the percentage of PHOs that present keyhole opportunities is unknown, the 

overall effectiveness of the gravity tractor is tentatively judged to be medium. 

G.2.4. Mass Driver

A mass driver, simply stated, is a pitching machine; in other words, it mines material 

from the asteroid and expels it at a high velocity, producing thrust. [78] [79] [80] [81] 

[82] The mission would require that a spacecraft rendezvous with the object and land a 

vehicle on the surface. The vehicle would attach a mining device to the surface and begin 

ejecting the mined material over a period of time. The attachment, mining, and expulsion 

activities would be autonomous. The device must have a power supply capable of 

sustained operation on the asteroid’s surface. 
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214 DRAFT Pre-Decisional Material

Expected Performance

No calculations were performed to estimate the performance of deflection efforts that use 

this technique. 

Characterization Required

The mass of the asteroid is critical for determining the size of the mining operation and 

the length of time needed to achieve the desired momentum transfer. The rotation rate is 

another critical item. The mass driver would only work when it can expel the mass in the 

correct direction. The types of material the mining operation would encounter and details 

of the asteroid’s internal structure (a hard, dense material may require different tools than 

would a light, porous material) also should be known. 

Technology Readiness

Mass Driver — The mass driver concept is relatively simple, but the technology is 

currently not well defined, particularly for space operation. Rails guns developed for 

Earth-based applications suffer from energy storage, material fatigue, relatively low 

muzzle velocities, and reliability issues. This technology area is rated as having a 

relatively low level of technology readiness. 

NEO Attachment and Mining — This concept requires attachment and an ongoing mining 

operation. Furthermore, the acceleration system must be able to dispose of the mined 

material. While many such concepts have been proposed, none has been built or tested. 

[83] This technology area is rated as having a relatively low level of technology 

readiness.

Autonomous Operations — The mass driver requires extensive, long-term autonomous 

operation of a complex system. Such operations have not been demonstrated. This 

technology area is rated as having a relatively low level of technology readiness. 

Overall Technology Readiness and Effectiveness

The mass driver has the advantage that, once it is affixed to an asteroid’s surface, it can 

continue to operate until the requisite ΔV has been imparted. This assumes, of course, 

that the system carries a suitable power supply (possibly a nuclear reactor). Problems 

include a reduction in efficiency because of the object’s rotation, the requirement for an 

onboard sensor system to determine the orientation of the PHO to assure the momentum 

is transferred in the correct direction, the attachment system, issues with the surface 

materials and internal structure, the possibility that large quantities of dust may be 

generated and affect the solar panels (should they be used), equipment and sensors, and 

the long period required for operation. Accordingly, the overall technology readiness of 

the mass driver technique is judged to be low. 

The mass driver may also have limited applicability. For example, it would be more 

difficult to attach and operate mass driver systems on small PHOs, where the 

gravitational attraction is smaller. Accordingly, the overall effectiveness for the mass 

driver technique is judged to be medium. 
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2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 215 

G.2.5. Pulsed Laser

A sufficiently powered ground-based laser currently does not exist. A variation of this 

concept would be to house the laser in a spacecraft stationed near the PHO. It would use 

large solar panels to capture solar energy (or draw power from a nuclear reactor) needed 

to drive a high-power laser. [84] The laser beam would be directed at a spot on the 

asteroid. Vaporization of the surface material would produce a small force on the asteroid. 

An advantage of this technique is that it imparts a small force and is unlikely to fragment 

the body. In addition, the laser beam can be directed to multiple spots, which is important 

if surface dust at a particular location degrades the transfer of energy. 

As recently as 1998, the U.S. Air Force unsuccessfully attempted to demonstrate the 

concept using space-based lasers smaller than the one required for this concept. Current 

estimates are that the spacecraft would weigh about 200 metric tons. [85] 

Expected Performance

No calculations were performed to estimate the performance of the pulsed-laser 

technique. 

Characterization Required

The PHO’s mass and type (asteroid as compared with comet) and the time before it 

strikes Earth define the size of the laser required for this mission. Since the method might 

be more effective against icy comets than asteroids (the effectiveness might be amplified 

by the creation of steam), [57] information about surface material would help estimate 

time of operation and overall effectiveness. 

Technology Readiness

Laser Technology and Surface Interactions — Development of the solar-powered, pulsed 

laser is in its infancy. This technology area is rated as having a relatively low level of 

technology readiness. 

Long-Term Autonomous Operations — As with other deflection concepts requiring 

extended periods of autonomous operation, such a capability has not been demonstrated. 

This technology area is rated as having a relatively medium level of technology 

readiness. 

Sustained Power — Solar panels or nuclear power might power a laser system. If solar 

photovoltaics were used, this technology area has a relatively high level of technology 

readiness. 

Overall Development Risk and Effectiveness

This concept requires substantial technology development and systems integration. The 

concept’s overall technology readiness is judged to be low. 

However, this approach could conceivably be applied to a large percentage of PHOs. For 

this reason, the overall effectiveness of the approach is judged to be medium. 
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216 DRAFT Pre-Decisional Material

G.2.6. Space Tug

Slow push techniques maintain better control over the asteroid as they apply a change in 

velocity. The asteroid tug is a vehicle that would dock with a PHO and push the object in 

the desired direction. 

In this case, characterization is very important. In particular, the asteroid’s center of mass 

defines a line along which the thrust vector must be applied to avoid adding rotational 

acceleration and reducing the system’s overall efficiency. And if the asteroid is rotating, 

the period when the thrust can be applied in the desired direction will be reduced, 

minimizing the technique’s efficiency, as seen in Figure 73. As illustrated, the thrust acts 

only when the orientation of the tug assures the thrust will impart momentum in the 

correct direction. 

Figure 73. Illustration of Space Tug acting on a PHO 

For example, assuming the worst possible scenario where the angle between the axis of 

rotation and the orbital plane is 90 degrees, and using the momentum equation derived 

from Izzo’s distance equation [86]: 

)

2

t

)(t

t

t

P T( p

s

s

p Δ = − (8)

Where: 

T is the thrust of the propulsion system, tp is the pushing time and ts is the time before 

impact, an equation for a rotating asteroid can be found. If the thrusters fire for half of the 

rotation along the orbital plane, the thrust efficiency can be calculated by: 
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2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 217 

π π

θ θ

β

π

π

1

 T(2 )

T cos d 2

2 = =

∫− (9)

Propulsion options for these systems include purely chemical systems (not generally 

feasible because of the large quantities of propellant and oxidizer that must be transported 

to the NEO), nuclear thermal propulsion (more efficient with lower propellant 

requirements, but not tested in space), and lower-thrust, electric-propulsion systems. Two 

of these systems were used for this analysis: the NSTAR solar-electric system using three 

and six thrusters and the NEXIS nuclear-electric system using five and 10 thrusters. 

The NSTAR thrusters are rated to produce 91 mN of thrust per thruster [87]. Placing a 

combination of thrusters on the vehicle could increase the thrust level. The NEXIS 

thrusters use electricity from a nuclear reactor to produce more thrust, roughly 473 mN 

per thruster. [88] Using many thrusters, the asteroid tug could produce more thrust and, 

as a result, more change in momentum. Using these values for the propulsion systems, 

the calculation seen in Equation (9) can be used to find the change in momentum. 

However, if the body is rotating, the thrust efficiency becomes a factor and the equation 

becomes: 

)

2

t

)(t

t

t

( T P p

s

s

p Δ rot = − π

(10)

M-5: Space Tug

Spacecraft Module Dry Mass (kg) 5572

Reactor Module Dry Mass (kg) 3177

Spacecraft Bus Mass (kg) 3130

Reactor Output (kWe) 120

Payload Power Non-thrusting (W) 10000

Payload Power Thrusting (W) 3000

Total Delta-V (m/s) 26000

Spacecraft Wet Mass (kg) 15119

9119

370

Spacecraft Dry Mass (kg)

Payload Mass (kg)

Spacecraft Module Dry Mass (kg) 5572

Reactor Module Dry Mass (kg) 3177

Spacecraft Bus Mass (kg) 3130

Reactor Output (kWe) 120

Payload Power Non-thrusting (W) 10000

Payload Power Thrusting (W) 3000

Total Delta-V (m/s) 26000

Spacecraft Wet Mass (kg) 15119

9119

370

Spacecraft Dry Mass (kg)

Payload Mass (kg)

System Sizing

Technical Description

• Nuclear Electric Propulsion (NEP) based spacecraft, JIMO like

• 12 year design life

• Payload module can accommodate lander for grapple attachment 

• Scientific payload for NEO characterization also applicable

• Some scaling on mass and reactor size available

CONOPS

• Imparts a momentum change on the NEO by moving it by pushing it or 

drawing it closer

• Launch on Delta IV Heavy vehicle

Rational for Consideration

• Given sufficient advance time, an NEP tug could provide controlled 

deflection of asteroids with anchoring

Capability

• Provides thrust and propellant capacity for anchored tug or to move 

NEO from colliding with Earth

Pros

• Slow deflection allows for flexible operational concepts

Cons

• Difficult political considerations, can be used to weaponize NEO

• High complexity, cost/risk due to NEO anchoring

• May be less/not effective on rubble piles, binaural, or rotating NEOs

• Requires very detailed NEO characterization level
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218 DRAFT Pre-Decisional Material

Expected Performance

Calculations were carried out on the space tug. The team based the calculations on 

payload capacities of the Delta IV Heavy and the proposed Ares V. The team also looked 

at various rotational rates of potentially hazardous objects. To deflect non-rotating PHOs, 

for example, C3 of 25 km2

/s2, the nominal effective momentum changes for the space tug, 

were calculated to be about 4 x 107

 kg-m/s and 6 x 108 kg-m/s, respectively. 

These calculations suggest that the ΔV potential of the space tug is somewhat greater than 

that of the gravity tractor, and in the range that might be accomplished using various 

kinetic impactors. For non-rotating PHOs, the space tug could be expected to provide 

ΔVs ranging from 10 to 100 cm/s for moderately small PHOs (smaller than 200 meters), 

to 0.1 to 1 cm/s for moderately large PHOs (smaller than about 900 meters), given 

sufficient lead times (in some cases, decades to a century in advance). The calculations 

also suggest that, with few exceptions, the space tug could not sufficiently deflect larger 

PHOs. 

Characterization Required

The space tug’s effectiveness greatly depends on the asteroid’s characteristics. It is very 

important for mission designers to know the target object’s mass and rate of rotation so 

that they can better define the action time and propulsive capabilities. They also should 

know surface characteristics since the device must attach to the surface. Orientation of 

the axis of rotation also is important because the tug can function only when oriented in 

the proper direction. The shape and location of the center of mass will affect the design 

and determine the overall effectiveness of the docking mechanisms and thrusting systems. 

Technology Readiness

Attachment Systems —- The docking mechanism is simply a concept. It has not been built, 

let alone proven in space. This technology area is rated as having a relatively low level of 

technology readiness. 

Long-Term Autonomous Operations — No one has verified the ability to operate major 

components of a space tug system for extended periods. This technology area is rated as a 

relatively medium level of technology readiness. 

Sustained Power — Power systems used in electric thrusters have flown and operated in 

space for long periods, but applying this technology to PHO deflection might require 

pulsed operation, particularly if the PHO is rotating. This would add complexity to an 

already-unproven and untested operational scenario. In addition, a nuclear thermal reactor 

propulsion system might be worth considering for this application because these systems 

could potentially provide very high thrust levels, which would sharply reduce mission 

time. This technology area is rated as having a relatively medium level of technology 

readiness. 

Overall Technology Readiness and Effectiveness

Small thrusters, while very efficient, are required to operate for extended periods. In 

many cases, the power delivered over time will be reduced by the rotation of the PHO. 

Furthermore, the line of action of the thrust must be through the center of gravity to avoid 

an unwanted rotational component to the energy delivered. Should the tug fail, a back-up 
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2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 219 

deflection system must be developed and launched. More important, it must work. Once 

the tug has been developed, it may prove to be a very effective alternative for moving 

PHOs, but given its current state of development, its overall technology readiness is 

judged as low. 

Despite this, the tug could in principle be used to deflect a high percentage of PHOs. For 

this reason, its overall effectiveness is judged to be medium. 
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220 DRAFT Pre-Decisional Material

Appendix H. Analysis Methodologies 

H.1. Analysis of Alternatives Methodology 

The analysis team started with selected concepts from the study working groups. The 

concepts were a subset of the overall concepts explored by the working groups. They 

were selected to address advanced space- and ground-based systems, systems that exist 

today, and those planned for the future. The analysis team, in coordination with the 

working groups, suggested additional concepts for evaluation to fully explore the trade 

space. The working groups endorsed all concepts that the analysis team evaluated as part 

of the Analysis of Alternatives (AoA). 

Initially, trade trees were constructed and concepts were grouped. Duplication was 

removed. In areas where the concepts could be grouped by similarity, one example was 

selected to represent the group. 

The analysis team then performed a first-order mass and power sizing on each concept, 

including spacecraft bus and instrument(s). The team also developed operational concepts 

and mission timelines. Details for existing and planned ground-based telescopes and 

radars were gathered and used to further define the concepts. System sizing tools, based 

on data from actual ground- and space-based systems, were developed to parametrically 

determine the mass and power of each space-based concept. 

The concept definition process included a review by the NASA working groups to ensure 

that concepts were correctly interpreted and that assumptions used in developing the 

concept design detail were consistent with the intent of the NASA working groups. 

In parallel, the analysis team developed the figure of merit (FOM) methodology to 

evaluate each of the concepts. This methodology included the approach; modeling and 

tools to determine life-cycle costs, nominal development time, mission risks; and 

qualitative and quantitative performance metrics for the detection, characterization, and 

deflection systems. 

Additional studies to explore issues associated with characterization and deflection 

missions were conducted. These provided additional contextual information on the 

overall feasibility of executing these missions, particularly in terms of required ΔV, 

launch capabilities, and launch and arrival timing. The analysis team also examined the 

precision tracking problem. It assessed the quality of orbit-determination capabilities for 

representative PHOs. 

Each concept was evaluated against cost, development time, development risk, mission 

risk, and the relevant performance FOMs. Concepts were then combined into 

architectures, and the overall funding profiles were developed for these architectures. The 

architectures selected consisted of sets of assets that would be used to support survey, 

characterization, and deflection programs. They represented a range in costs and 

capabilities. 
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Evaluation results of the individual concepts, program architectures, and end-to-end 

architectures were compiled. The performance of potential survey systems and tracking 

capabilities were integrated into the results. This was done to develop cost-benefit and 

overall effectiveness assessments. Analysis team results, including the detailed 

descriptions of the concepts, figure of merit methodology, and results were documented 

and integrated into the final study report. 

H.2. Survey Performance Analysis 

H.2.1. Performance Simulations

Performance simulations for a number of candidate ground-based and space-based PHO 

detection systems were conducted at JPL by Steve Chesley and Paul Chodas. The 

detection capabilities of the systems were tested using a set of 1000 test PHOs from the 

Bottke et al. population. [4] The positions of each PHO on the plane-of-sky of each 

detector system were simulated daily from the assumed operational start date of the 

system through the year 2030. The PHO was a candidate for detection on a given day by 

a given system if it fell within the search region for that system for that day, and if it was 

sufficiently bright to be detected. To study the dependence of magnitude on object size, 

each PHO was replicated into 25 different size bins, from 32 meters to 10 km in size, and 

the magnitude predicted for each assumed size. One component of the magnitude 

calculation was the geometry, the distances of the object from the Earth and Sun and the 

phase angle, but another important component was the assumption used for the object’s 

albedo (reflectivity). 

The PHO population was divided into two equal groups – one with a darker albedo of 7% 

and the other with a brighter 23% albedo. This bimodal albedo distribution is thought to 

be representative of the actual PHO population. Another correction applied to the 

magnitude calculation was due to an object’s motion during the sensor integration time, 

producing a small trail on the detector rather than a single spot of light. The predicted 

magnitude was corrected for these so-called trailing losses by using a reduction 

proportional to the square root of the trail length. Also, for ground-based systems, an air 

mass correction was made to account for the loss in sensitivity as the detector looks 

through the increasingly large air mass when looking away from the zenith. 

Each simulated PHO was considered to be detected if it fell within the region covered by 

a simulated sky scan, and if it was brighter than the system’s limiting magnitude. The 

object was not considered to be discovered until it had been detected on two separate 

nights within a one week period and then a third time within 40 days. Simulations are 

carried out under the assumption that all survey alternatives will do their own follow up 

observations as part of their operational cadences due to the large aperture needed to 

detect the fainter PHOs. 

Additionally, a magnitude correction was calculated for solar phase angle. Spacecraft in 

Venus like orbits preferentially view the PHOs near perihelion and full phase. Groundbased systems see PHOs and IEOs in particular at large phase angles which significantly 

reduced their apparent magnitude. Consequently, the same IEO would appear 1-2 

magnitudes brighter as seen from a Venus-like orbit compared to a ground-based 

telescope. 
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222 DRAFT Pre-Decisional Material

Simulation of Ground-Based Systems

The ground-based detection systems were simulated by following a monthly pattern. For 

a period of time around each new moon, the system’s search region was assumed to be 

completely scanned three to four times. The limits of the search region varied slightly 

from system to system. In addition, in all but one case, the search region was limited by 

specifying two limiting air-masses, one to be used within 10 degrees of the ecliptic, and 

the other to be used away from the ecliptic. Typically, these limits were set at 2.5 and 2.0 

air masses, respectively, corresponding to elevation limits of about 23.5 and 30 degrees, 

respectively. 

The time required to complete a full scan of the search region also varied from system to 

system, depending on such system parameters as the exposure time, field of view, and 

percentage of time allocated to PHO observations (i.e., whether the system was shared 

with non-PHO observations or was operated in a 100% dedicated mode). This cycle time 

varied from 4 to 6 days. No attempt to avoid the galactic plane was made. To account for 

losses due to poor weather and moonlight, only a portion of the available nights each 

month were assumed productive. For the shared systems, only 12 to 15 nights per month 

were assumed productive for PHO searches, while for the dedicated systems, 18 to 20 

nights were assumed productive. 

Large Aperture Synoptic Survey Telescope (LSST) - Two versions of the LSST were 

simulated, a shared system and a dedicated system. In shared mode the telescope would 

not be dedicated to PHO observations, with only 75% of the observing time assumed 

available for PHO observations, and the telescope would operate at a somewhat reduced 

sensitivity to obtain color data. In this mode, the limiting magnitude would be 24.8 and 

the search region would be more restricted (the off-ecliptic air-mass limit was set to 1.5, 

corresponding to an elevation of 42 degrees), and only 3 scan cycles were assumed to be 

available each month with each cycle requiring 5 days. 

To simulate the Dedicated LSST, the limiting magnitude was increased to 25.4, the 

search region was increased by increasing the off-ecliptic air mass limit to 2.0 air masses 

(elevation of 30 deg), and the number of scan cycles was increased to 4 per month. The 

assumed operational start dates for the shared and Dedicated LSST are 2014 and 2016 

respectively. 

PanSTARRS 4 telescope array (PS4) – For PS4, four collocated 1.8 meter aperture 

telescopes provide a 7 degree field of view, and it is assumed that a planned active seeing 

correction system will be capable of observing down to an air mass of 2.0 (2.5 near the 

ecliptic plane). Two versions of PS4 were simulated: shared and dedicated each with an 

assumed limiting magnitude of 24.0. When operated as a shared asset, only 30% of the 

time would be available for PHO observations, and it would not be possible to scan the 

entire available sky. To simulate this mode, the search region was reduced to a large box 

around the opposition point and two smaller boxes around the two quadrature points. The 

scan cycle was assumed to require 4 days, with 3 scan cycles fit into each lunation. 

To simulate a dedicated PS4 system, the search region was expanded to the entire sky 

down to an air mass limit of 2.0 (elevation of 30 deg), along with the usual near-ecliptic 

limit of 2.5 air masses. In this mode, the scan cycle would lengthen to 6 days. The 
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2006 Near-Earth Object Survey and Deflection Study 

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assumed operational start dates for the shared and dedicated PS4 systems are 2010 and 

2013 respectively. 

PS8 - The PS8 system consists of two geographically collocated PS4 systems. The two 

telescopes were assumed to operate in conjunction with one another, so that a fainter 

limiting magnitude of 24.4 could be attained with the same exposure times as used by 

PS4. 

PS16 - The PS16 system was simulated by replicating two PS8 systems, one each in the 

northern and southern hemispheres. Each operated independently so the limiting 

magnitude was the same as the PS8 system (24.4). Both the PS8 and PS16 systems are 

assumed to begin operations in 2014. 

Simulation of Space-Based Systems

Both visible and near infrared wavelength space-based systems were considered, located 

either at the first Lagrange point, L1 (323,000 km sunward from Earth on a line between 

the Sun and Earth), or in a Venus-like orbit about the Sun. Each space-based system was 

assumed to follow a daily 23-hour duty cycle; with the remaining hour being used to 

transmit the day’s data to an Earth based receiving station. No attempt to avoid the 

galactic plane was made. The operational start dates for the space-based systems are 2013 

for the 0.5 m IR telescope at L1 and 2014 for all other space-based IR and visual 

telescopes, except for the 2.0m visual system in LEO or at L1. 

Visible Space-Based System - A 2 meter aperture telescope, located at L1 or in a Venuslike orbit, was considered. 

Infrared (IR) Space-Based System - Because of the greater reflected solar flux in the IR 

band pass, smaller apertures were used, and apertures of both 0.5 and 1.0 meters were 

simulated. The zodiacal light background confusion was considered separately for the IR 

systems at both the L1 and Venus orbit locations. 

H.2.2. Survey Performance Simulation Verification 

To verify the current JPL detection performance simulations, several test cases were run 

against similar cases carried out in 2003 for the SDT report [2] using independent 

software (i.e., FROSST) provided by the MIT’s Lincoln Laboratory. For these 

verification runs, a single albedo of 14% was assumed and for the cases that are of the 

highest interest, the differences between the JPL and Lincoln Lab simulation results were 

within 2%of one another, thus verifying the current simulation software. Results may be 

considered valid to within 2-5% of expected results. 

H.3. Deflection Analysis Methodology 

H.3.1. The Asteroid Deflection Formula

The momentum change required to deflect a PHO can be estimated based on an analysis 

discussed in Reference [86]. The quantities discussed in this analysis are shown in Figure 

74. This analysis approximates PHO deflection performance by assuming that the given 

deflection produces a relatively small perturbation in the PHO orbit characteristics. This 

assumption was evaluated in Reference [89] using the JPL PHO database [71] and found 
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2006 Near-Earth Object Survey and Deflection Study 

224 DRAFT Pre-Decisional Material

to agree with accurate numerical simulations of deflections. In this formulation, the 

minimum distance by which the PHO would miss the earth is given by: 

(t t) v A dt 3 a v d n

t

0 s

e

min

p

= − ⋅ ∫ μ

γ (11)

Where: 

a is the target object orbit semimajor axis 

γ is a non-dimensional parameter that describes the encounter geometry 

ve is the Earth’s velocity at encounter 

ts is the time before impact when the deflection is started 

tp is the time to complete the deflection 

μ is the gravitational constant of the Sun 

vn is a vector representing the unperturbed velocity of the PHO 

A is the action (acceleration) applied to the PHO 

The quantity, γ, is determined from the following equations: 

γ φ () ( ) 1 2 φ 1 φ cosα 2 2 2

n e n e = v + + v − v + v (12)

α

α

φ 2 cos

cos

2 2

n e n e

n e

v v v v

v v

+ −

− = (13)

The encounter geometry between the PHO and the Earth is illustrated in Figure 74. The 

quantity γ becomes small when the relative velocity between the PHO and Earth is large 

and the angle α is small. 

ve 

vn 

Earth NEO α

dmin 

γ

s = 

Figure 74. PHO/Earth Encounter Geometry 

When a PHO deflection is carried out impulsively, either through kinetic energy impact 

or explosive energy transfer, equation (1) may be simplified to the following form for 

low eccentricity orbits. 
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2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 225 

n

s

n m

pt d C Δ

min = (14) 

p m v Δ = nΔ (15) 

H.3.2. Effective Momentum Change

The miss distance equation for slow push methods is calculated as 

dt

M

T(t) (t t) v

3 a v d

n

t

0 s

e

min

p

= − ⋅ ∫ μ

γ (16)

Where: 

T(t) is the thrust vector as a function of time 

Mn is the mass of the asteroid 

Assuming that the thrust is directed optimally along the unperturbed velocity vector, 

thrust is constant, and the eccentricity of the orbit is small; the integral can be solved to 

produce: 

n

s

s

p

m

t

t

t

C

⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛ −

= 2

Tt 1

d

p

min n

(17)

To compare impulsive and slow push deflection concepts that are started at the same time 

relative to the Earth impact date, the action which creates an equivalent minimum miss 

distance, dmin is evaluated. With all other quantities being equal, the effective momentum 

change imparted by a slow push deflection technique is: 

⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛ Δ = −

s

p

eff p t

t

p Tt

2

1 (18)

This equation shows that while the effective momentum change grows with action time, 

the momentum effect diminishes as the action time becomes significant with respect to 

the time to impact. 

H.4. Technology Readiness Assessment of Deflection Alternatives 

Table 47 summarizes the readiness of the primary technology areas that apply to 

impulsive techniques. Table 48 summarizes the readiness of the primary technology areas 

that apply to the slow push techniques. The rationale for these ratings is described in 

greater detail in Appendix G. 
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226 DRAFT Pre-Decisional Material

Table 47. Readiness for Technology Areas for Impulsive Techniques 

Subsystem Technology Readiness Overall 

Conventional 

Explosive (contact) 

Explosive 

Devices 

Delivery Systems 

Targeting Fusing High 

Conventional 

Explosive (subsurface) 

Explosive 

Devices 

Delivery Systems 

Implantation Fusing Medium 

Kinetic Impact Delivery Systems 

Targeting High 

Nuclear Explosive 

(contact) 

Explosive 

Devices 

Delivery Systems 

Targeting Fusing High 

Nuclear Explosive 

(standoff) 

Explosive 

Devices 

Delivery Systems 

Targeting Fusing High 

Nuclear Explosive 

(subsurface) 

Explosive 

Devices 

Delivery Systems 

Implantation Fusing Medium 

Nuclear Explosive 

(surface delayed) 

Explosive 

Devices 

Delivery Systems 

Attachment Fusing Medium 

Table 48. Readiness for Technology Areas for for Slow Push Techniques 

Critical Technology Areas Overall 

Enhanced 

Yarkovsky 

Coating 

concepts 

Coating 

delivery Low 

Focused 

Solar 

Large Space 

Structures 

Delivery and 

Proximity Operations Low 

Gravity 

Tractor 

Control 

Systems Sustained Power Autonomous Operations Medium 

Mass Driver Mass Driver Attachments 

& Mining 

Autonomous 

Operations Low 

Pulsed Laser Laser and Surface Sustained 

Power 

Long Term 

Operations Low 

Space Tug Sustained 

Power 

Attachment 

Systems 

Autonomous 

Operations Low 
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Appendix I. Additional Survey Performance Results 

Figure 89 through Figure 93 display the percentage of completeness for each year ending 

in 90% completeness. The following performance architectures are shown as examples: 

• Figure 75. Buildup of Baseline – Existing + Shared PS4 + Shared LSST 

• Figure 76. Buildup of Baseline + Dedicated PS8 

• Figure 77. Buildup of Baseline + Dedicated LSST 

• Figure 78. Buildup of Baseline + 0.5m IR @ L1 

• Figure 79. Buildup of Baseline + 0.5m IR @ Venus 

0%

10%

20%

30%

40%

50%

60%

70%

80%

90%

100%

2000 2005 2010 2015 2020 2025 2030

Year

Percent PHO Population Catalogued

1. Existing (Spaceguard)

 Cost: < $10M FY06

2. Add Shared PS4

 Cost: $10M FY06 

3. Add Shared LSST

 Cost: $140M FY06

90% reached 

83% by 2026

2021 All Costs are Life Cycle 

Cost through 2020

Figure 75. Buildup of Baseline – Existing + Shared PS4 + Shared LSST 
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228 DRAFT Pre-Decisional Material

0%

10%

20%

30%

40%

50%

60%

70%

80%

90%

100%

2000 2005 2010 2015 2020 2025 2030

Year

Percent PHO Population Catalogued

1. Baseline (Spaceguard, 

 Shared PS4, Shared LSST)

 Cost: $150M FY06

2. Add Dedicated PS8

 Cost: $100M FY06

90% reached 

in 2024

85% by 

2021

All Costs are Life Cycle 

Cost through 2020

Figure 76. Buildup of Baseline + Dedicated PS8 

0%

10%

20%

30%

40%

50%

60%

70%

80%

90%

100%

2000 2005 2010 2015 2020 2025 2030

Year

Percent PHO Population Catalogued

2. Add Dedicated LSST

 Cost: $500M FY06

90% in 

2020

1. Baseline (Spaceguard, 

 Shared PS4, Shared LSST)

 Cost: $150M FY06

Figure 77. Buildup of Baseline + Dedicated LSST 
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2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 229 

0%

10%

20%

30%

40%

50%

60%

70%

80%

90%

100%

2000 2005 2010 2015 2020 2025 2030

Year

Percent PHO Population Catalogued

2. Add 0.5m IR L1/L2

 Cost: 1.0 $B 2006

90% by 

2020

1. Baseline (Spaceguard, 

 Shared PS4, Shared LSST)

 Cost: 0.15 $B 2006

Figure 78. Buildup of Baseline + 0.5m IR @ L1 

0%

10%

20%

30%

40%

50%

60%

70%

80%

90%

100%

2000 2005 2010 2015 2020 2025 2030

Year

Percent PHO Population Catalogued

2. Add 0.5m IR Venus-like 

 Cost: $1.1B FY06

90% reached 

in 2017

97% by 

2021

1. Baseline (Spaceguard, 

 Shared PS4, Shared LSST)

 Cost: $150M FY06

Figure 79. Buildup of Baseline + 0.5m IR @ Venus 
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Appendix J. Current Assets and Activities 

J.1. International Activities 

J.1.1. United Nations

In its 1999 UNISPACE III conference, the standing United Nation’s Committee on the 

Peaceful Uses of Outer Space (UNCOPUOS), recommended that NEOs be a topic of 

future consideration. Since then, the UNCOPUOS Scientific and Technical 

Subcommittee (STSC) has included NEO on its agenda and has established an Action 

Team of interested member states to make recommendations to the STSC and the full 

Committee. 

The STSC currently is involved in a 3-year work plan (2005-2007) that essentially allows 

member states to report on national NEO activities and facilitates presentations at the 

STSC by interested international organizations and non-governmental organizations 

(NGOs). To date, no specific proposals have been made by UNCOPUOS member states, 

but various NGOs have made proposals during presentations to UNCOPUOS. 

In 2007, the STSC Action Team may recommend that a standing STSC Working Group 

be established to address a full slate of NEO issues, including threat deflection and 

international decision making. The U.S. is an active participant in UNCOPUOS and the 

STSC. 

J.1.2. European Space Agency (ESA)

In July 2004, ESA established and international panel, the Near-Earth Object Mission 

Advisory Panel (NEOMAP), to recommend next steps in the area of NEO investigation 

and research. NEOMAP produced a set of recommendations for observatory and 

rendezvous missions from an international context. 

Flowing from NEOMAP, ESA has established the Don Quijote mission as one of its top 

space science missions. Don Quijote is an asteroid deflection precursor mission designed 

to assess and validate technology that might be used to deflect a NEO. It will consist of 

two separate spacecraft, launched on separate interplanetary trajectories. The orbiter 

(Sancho) will arrive at the target asteroid first to assess, measure, and monitor. The 

impactor (Hidalgo) will arrive later to hit the asteroid. Sancho will monitor the result. The 

7-year mission will launch no earlier than 2011. All major ESA members, including Italy, 

France, and Germany, will have a role in Don Quijote. 

J.1.3. Italy – Italian Space Agency (ASI)

The Italian scientific community participates in the observation campaigns mounted by 

other countries. For example, the Italian Space Agency (ASI) is participating in the 

NASA Dawn Discovery mission. It is providing a visible-infrared mapping spectrometer. 

ASI also is participating in ESA’s Rosetta mission, which was launched in 2004. It will 

fly by two asteroids in 2008 and 2010 en route to an encounter with comet 

67P/Churyumov-Gerasimenko in 2014. Italy will be a prime participant in the ESA Don 
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Quijote mission. Italy also is currently engaged in discussions with the French Space 

Agency (CNES) on a potential future bilateral NEO space mission. 

Italy is very active in the detection, tracking, and characterization of NEOs, primarily 

through its NEO Dynamic Site (NEODys), which is similar to JPL’s Sentry system. 

Fundamentally, NEODyS is a database of NEO information; however, the system’s 

distinguishing feature is that each near-Earth asteroid has its own “home page.” The site 

offers sections devoted to the object’s orbit, observations, and close encounters. 

NEODyS provides interactive ephemeris services to the observer in both tabular and 

graphical forms. The graphical ephemeris depicts the uncertainty region on the celestial 

sphere, including nonlinear effects, which can be of paramount importance in the 

recovery of lost or nearly lost objects. With the database query service, one may search 

for all asteroids demonstrating some desired orbital characteristic. This makes it easy, for 

example, to find all he asteroids that are large enough to be hazardous, yet are effectively 

lost. 

The database is automatically updated daily, as new observations are released from the 

Minor Planet Center. All data files needed to reproduce the NEODyS results are freely 

available. 

J.1.4. Japan - Japan’s Aerospace Exploration Agency (JAXA)

Japan’s Aerospace Exploration Agency (JAXA) is currently involved in a major NEO 

research mission, called Hayabusa. The mission included the 2003 launch of a spacecraft 

to rendezvous with the asteroid, Itokawa, and the subsequent rendezvous and observation 

campaign in the fall of 2005. Observations will occur from as close as 3 km. The 

spacecraft’s return, including the samples it gathered, was planned for 2007, but due to an 

anomaly in the spacecraft’s propulsion system, JAXA is working on an alternative plan to 

return the spacecraft in 2010. NASA and DLR are participating with JAXA in the 

Hayabusa mission. NASA is participating in the Hayabusa mission in these areas: 

• Heat shield development at Ames 

• Tracking including data receiving at DSN stations 

• Ground observations of Itokawa for ephemeris and shape information, including 

radar observations 

J.1.5. Germany - German Aerospace Center (DLR)

DLR activities include use of ground-based and space-borne astronomical telescopes 

(primarily NASA assets) for observations, theoretical investigations, and data reduction 

and analysis. DLR is interpreting data from NASA’s Deep Impact mission and from 

JAXA’s Hayabusa mission. DLR also participates in the European Fireball Network, a 

network of all-sky cameras that record and track information on NEOs. 

DLR also has proposed establishing the German Spaceguard Center, which would serve a 

similar function as JPL’s Near-Earth Object Program Office and the United Kingdom’s 

Near-Earth Object Information Centre. 
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J.1.6. Canada

Defense Research and Development Canada (DRDC) and the Canadian Space Agency 

(CSA) are collaborating to place a microsatellite in low-Earth orbit to perform optical 

detection and tracking of both IEOs and Earth-orbiting satellites and debris (i.e., 

“Resident Space Objects,” RSOs). 

In 2009, Canada plans to launch the NEO Surveillance Satellite (NEOSSat), a spacecraft 

equipped with a 15-cm, visible-imaging sensor. The tracking mission will conduct 

repeated surveys, with the aim of finding more than 50% of all IEOs whose diameters 

exceed more than 1 km. 

J.1.7. United Kingdom (UK)

The British National Space Centre (BNSC) coordinates and monitors NEO activities and 

serves as the policy lead for all NEO-related matters. Two national centers provide 

information on NEOs. Within the NEO research community, the Spaceguard Centre 

coordinates the national Comet & Asteroid Information Network and liaises with 

Spaceguard organizations in 17 different countries. The Near-Earth Object Information 

Centre provides public outreach and information to the general population. The UK is a 

strong advocate of NEO activities within the European Space Agency and has taken a 

lead role within UNCOPUOS in this area. The UK heads UNCOPUOS Action Team #14, 

which is studying NEO-related issues. While it is taking a strong role with respect to 

NEO policy issues, the UK government’s actual direct funding of NEO research is 

relatively modest, with most UK research efforts taking place in the academic sector. 

J.2. United States Activities 

J.2.1. NASA

Prior to the NASA Authorization Act of 2005, NASA’s responsibilities included: 

1. Solicit and select all science investigations, ground-based and space-based, for the 

detection and scientific exploration of NEOs; 

2. Coordinate with other agencies and organizations, including international 

agencies and organizations, and 

3. Assess the evolving understanding of NEOs, which will result from the search 

and characterization effort to guide strategic planning and mission selection. 

After several years of studies attempting to define the threat posed by NEOs and the 

capabilities necessary to more comprehensively detect a sizable portion of the population, 

NASA in 1998 adopted the goals of the Shoemaker Spaceguard Survey and committed to 

“...find and catalog 90% of Near-Earth Objects larger than one km in the next ten 

years.” This became NASA’s Near-Earth Objects Observation (NEOO) Program, for 

which funding was identified within the Planetary Astronomy Research and Analysis 

Program to pay teams of asteroid astronomers to search near-Earth space for these 

potentially hazardous objects. Since that time, funding for the NEOO program has 

leveled out to slightly more than $4 million per year, which funds activities of the NEO 

Program Office and five NEO search teams. 
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J.2.2. Jet Propulsion Laboratory (JPL) NEO Program Office

As part of the NEO Observation Program, NASA established an NEO Program Office at 

JPL with the following responsibilities: 

1. Coordinate ground-based observations to complete the survey of NEOs and obtain 

accurate orbital elements for newly detected NEOs based on the best available 

data. 

2. Facilitate communication both within the observing community and between the 

community and the public about any PHOs that are discovered as a result of the 

expanded observational program. The office also responds to public inquiries. 

3. Establish, update, and maintain a catalog of NEOs, together with an estimate of 

the quality of the orbital elements accessible to the scientific community and the 

public. 

4. Develop and support a strategy and plan for the scientific exploration of NEOs, 

including their discovery, recovery, ephemeredes, characterization, in-situ 

investigation, and resource potential. 

5. Support NASA HQ in coordinating with other government agencies, foreign 

governments, and international organizations on all NEO issues. 

On a daily basis, the Minor Planet Center makes available NEO astrometric data to JPL’s 

NEO Program Office and to a parallel, but independent orbit computational center in Pisa, 

Italy (with a mirror site in Valladolid, Spain). At JPL, 340,000 bodies in a searchable 

Small Bodies Database are maintained for the international community. 

JPL’s Horizons on-line system is an interactive ephemeris generation site that 

automatically generates about 3,000 products per day to the international science 

community (http://horizons.jpl.nasa.gov). Within the JPL SENTRY system 

(http://neo.jpl.nasa.gov/risk/), risk analyses are automatically run on those objects that 

have a potential for Earth impact. They usually consist of newly discovered objects for 

which orbital information is unknown. These objects are prioritized in the SENTRY 

system, according to the quality of their orbital data and their potential to closely 

approach Earth’s orbit. 

The JPL system automatically updates the orbits of about 40 NEOs per day and closeapproach tables are generated and posted on the Web (http://neo.jpl.nasa.gov/cgibin/neo_ca). Approximately five risk analyses are run each day. Each run provides 

10,000 multiple solutions out to 2105. A parallel process in Pisa, Italy, and significantly 

non-zero Earth impact cases are manually checked between JPL and Pisa before the riskanalysis data are posted on the respective Web sites. Since its inception in 2002, about 

400 objects have appeared on the SENTRY risk page. For recently discovered objects of 

unusual interest, the MPC, JPL, and Pisa will often alert observers that additional future 

or precovery data are needed. 

J.2.3. Harvard/Smithsonian Minor Planet Center

The Smithsonian Astrophysical Observatory, in coordination with the International 

Astronomical Union (IAU) through a Memorandum of Agreement (MOA), operates the 

Minor Planet Center (MPC). Given its association with IAU, it has an international 
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charter. Since 1978, the MPC has served as the international clearinghouse for all 

asteroid, comet, and satellite astrometric and positional measurements obtained 

worldwide. 

The MPC processes and organizes data, identifies objects, computes orbits, assigns 

tentative names, and disseminates information on a daily basis. For objects of special 

interest, the center solicits follow-up observations and requests archival data searches. 

The MPC is responsible for distributing astrometric observations and orbits via the Minor 

Planet Electronic Circulars (issued as necessary, generally at least once per day) and 

related catalogues. 

In addition to distributing complete orbit and astrometric catalogs for all small bodies in 

the solar system, the MPC facilitates follow-up observations of new potential NEOs by 

placing candidate sky-plane ephemeredes and uncertainty maps on the Web via The NEO 

Confirmation Page. The MPC focuses specifically on identification, short-arc orbit 

determination, and dissemination of information pertaining to NEOs. In most cases, NEO 

observations are distributed to the public, free of charge, within 24 hours of receipt. The 

MPC also provides a variety of tools to support the NEO initiative, including skycoverage maps, lists of known NEOs, lists of NEO discoverers, and a page that allows 

users to select a list of known NEOs in need of astrometric follow-up. The MPC also 

maintains a suite of programs to calculate the probability that any object is a new NEO 

based simply upon two sky-plane positions and a magnitude. [93] 

Primary MPC funding in 2003 was $130,000 a year from NASA, despite the Agency's 

increased spending on NEO surveys. Other income from subscriptions and donations is 

insufficient to cover the 80- to 100-hour workweeks currently staffed by director Brian 

Marsden, associate director Gareth Williams, and a few graduate students. Marsden, now 

69, and Williams are critical assets to the current NEO discovery and cataloguing efforts. 

MPC is having difficulties keeping up with the volume of data it receives now. A new 

survey that will generate up to 100 times more data and the same need for follow-up 

observations will likely overwhelm the organization under its current structure. 

J.2.4. Spaceguard Survey Assets and Capabilities

The NASA NEOO program has leveled out to slightly more than $4 million per year. The 

money funds five NEO search teams, who operate nine separate, one-meter-class survey 

telescopes across the southwestern U.S., Hawaii, and Australia. These five teams are 

listed below: 

Spacewatch, operated by the University of Arizona Lunar and Planetary Laboratory, 

provides two telescopes on Kitt Peak, AZ, http://spacewatch.lpl.arizona.edu/. 

Near-Earth Asteroid Tracker (NEAT), operated by JPL, operates detection cameras on a 

U.S. Air Force telescope on Maui, Hawaii, and on a telescope at Palomar, CA, 

http://neat.jpl.nasa.gov/. 

Lincoln Near-Earth Asteroid Research (LINEAR) is operated by MIT/LL under a U.S. 

Air Force contract funded by NASA. Two telescopes operate near Socorro, NM, 

http://www.ll.mit.edu/LINEAR/. 
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Lowell Observatory Near-Earth Object Search (LONEOS) is obviously operated by the 

Lowell Observatory near Flagstaff, Arizona, 

http://asteroid.lowell.edu/asteroid/loneos/loneos.html. 

Catalina Sky Survey, operated by a separate team also at the University of Arizona Lunar 

and Planetary Laboratory, operates two telescopes on Mt. Lemmon, Arizona, and one at 

Siding Spring, Australia, its only Southern Hemisphere asset, 

http://www.lpl.arizona.edu/css/. 

Planetary Radar Systems - Arecibo and Goldstone 

Two planetary radars are capable of observing near-Earth objects. Radar data are 

extremely powerful in reducing orbital uncertainties, as discussed earlier in this report. In 

addition to gathering data on NEOs, both facilities carry out radar investigations of 

planets in the inner solar system and small bodies that orbit as far as the main asteroid 

belt. 

The Goldstone radar is located in southern California’s Mojave Desert. It uses the 70- 

meter antenna of NASA’s Deep Space Network (DSN14), which is currently equipped 

with a 450-kW transmitter. It can receive on this dish or other nearby DSN antennae. 

Because it is steerable, the antenna can reach most of the sky and can follow the oftenrapid apparent motions of NEOs. 

The second radar is located at Arecibo, Puerto Rico. It is owned and managed by the 

National Science Foundation (NSF) and operated by Cornell University under a 

cooperative agreement with NSF. It is equipped with a 305-meter aperture and a 900-kW 

transmitter. Its reach is farther than that of Goldstone, but because it has a fixed antenna, 

it can only look about 20° off its zenith position. 

As of early September 2006, 195 near-Earth asteroids and 11 comets were observed with 

the planetary radars. 

J.2.5. Current Cataloguing and Data Management Infrastructure

To catalog a newly discovered NEO, enough observations must be obtained to determine 

its orbital path about the Sun. Currently, most new objects are detected by one or more of 

the five NASA-funded search projects, operating up to nine different ground-based 

telescopes. See 0. The sunlight reflected from almost all new NEOs is so dim that 

astronomers can search for them only on the darkest nights of the month, usually for 

about two weeks when moonlight does not obscure their light, and, of course, only on 

nights when the sky is relatively clear. 

To determine whether the detected point of light is not just another star but a small object 

moving in the solar system, multiple images must be collected of the same area of sky. 

Enough time must pass between observations to allow the relatively near object to move 

against the background of stationary stars. Currently, at least three images, sometimes 

five, are taken each night of the same area of sky, separated by about 30 minutes on 

average. A computer-automated image processor then compares the images. A NEO will 

appear as a point of light moving a few arc seconds across the rectified, sequentially 

ordered images. This is considered a detection. 
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Observation data, in the form of right ascension and declination position angles, can then 

be taken on the detected moving object. Astronomers can compare the object’s positions 

to the known positions of the background stars. This observational process is called 

“astrometrics.” With a good star catalog, astronomers can take precise observations with 

errors of less than one-half an arc second. All the search projects collect multiple 

observations of the objects each observable night. 

The projects forward these data to the Minor Planet Center, where they are either 

correlated with already-known objects or determined to be previously unknown objects. 

For these new objects, an initial rough orbit can be determined. Most detections are 

objects in the main asteroid belt beyond the orbit of Mars, but a handful each night are 

determined to be NEOs. The MPC publishes the predicted orbital positions, called 

“ephemeredes,” for new NEOs on its NEO Confirmation Page as a way to solicit and 

collect additional observations, often done by talented amateur astronomers. Additional 

observations must be collected within a few nights to sufficiently determine the new 

object’s orbit, or “secure” it, so that it can be tracked into the future. If the object is 

tracked on at least 3 nights over a span of 40 days, it can be cataloged as a new 

“discovery.” 

If the orbit of a newly detected NEO is predicted to come within 0.05 AU (about 5 

million miles) of Earth’s orbital path, it is a possible new PHO and the MPC will send an 

alert to the NEO Program Office at JPL and to the NEODys operated by the University of 

Pisa, Italy (now with a backup site at the University of Valladolid, Spain). If the newly 

determined orbit shows there is a possibility of an Earth impact at some point in the 

future, the MPC also will send an alert of a new “PHO of Interest.” 

Observatories and amateurs, who attempt to collect follow-up observations, receive these 

alerts and often send what data they collect to the MPC. The MPC updates the 

ephemeredes, and if a possible Earth impact still exists, more observations are directly 

solicited from observers who have a good chance of being able to observe the object. 

Meanwhile, all of the MPC data are automatically sent to JPL and the University of Pisa. 

JPL and NEODys perform separate precision orbit determinations and calculate the 

potential of Earth impact. If observations are obtained confirming the hazardous impact 

orbit, then the organizations pass along the information up the management chain at 

NASA until it reaches the Administrator. 

If, at any time, astronomers determine that an object no longer poses a threat, MPC’s 

routine object processing capability catalogues the orbital information. 
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Survey,

Detect,

& Report

Correlate, Determine

Rough Orbit

Possible

New PHO?

Routine

Processing

Publish

Results

Yes

Possible

Impact?

Resolve

Result

Differences

Publish

Results

No

Precision Orbit

and Follow Up 

Observations 

Impact 

Still 

Possible?

Observations and

Update Orbit

Publish/

Update

Results

No

No

Yes

Yes

Iterate 

Survey Systems

MPC*

JPL (or Pisa)

Radar

Alerts

• MPC PHO 

of interest

• MPC

possible 

impact

• JPL reports 

possible 

impact to 

NASA/HQ

• JPL

publishes 

probable 

possibility 

of impact

* Minor Planet Center

Figure 80. Data Flow for Spaceguard Survey 

Four different types of infrastructure keep this system running: search projects detect new 

objects; observatories — some of which are run by amateurs — collect follow-up 

observations; JPL and Pisa precisely determine orbits and potential impacts; and the MPC 

collects and archives all data. The MPC maintains an archive of all observations received, 

including those that have never been correlated with a cataloged object in hopes that 

someday they will – perhaps providing a critical extension to an orbital arc. MPC also 

maintains an archive of all orbital parameters ever calculated on known objects. 

The MPC is currently the only non-redundant part of the data analysis and storage system 

that catalogs and tracks NEOs. Additionally, if the search is expanded to smaller NEOs, 

the day-to-day work that the MPC needs to perform will increase as new, more capable, 

observing platforms become available. It is expected that the data throughput will 

increase at least tenfold, and possibly as much as a hundredfold. The MPC will need to 

substantially update its hardware and software. 

It also is possible that the work done at the MPC and at JPL could be done elsewhere. It 

might be consolidated into one central data management center for NEOs. Potential sites 

include a few of the NASA centers, or major universities or science institutes involved in 

small-body planetary science. 

NASA’s Planetary Data System (PDS) is a distributed archive of solar system data 

prepared in a standard format for use primarily by astronomers, mission planners, 

educators, and students. NASA’s Office of Space Science Exploration of the Solar 

System Division sponsors PDS to ensure the long-term use of data, to stimulate research, 

to facilitate data access, and to support correlative data analysis. 

The Small Bodies Node (SBN) is a distributed node, which specializes in data concerning 

asteroids, comets, and interplanetary dust. Its archives consist of data from NASA 

missions, astronomical observations, and laboratory measurements, organized into a 
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238 DRAFT Pre-Decisional Material

structure and format suitable for archival and retrospective research by the scientific 

community. The node data collection and verification activities are spread among several 

institutions specializing in particular bodies (comets, asteroids, or dust), while the main 

archives and user services are collected in a single place. 

The Comet Subnode is located at the University of Maryland, College Park, MD. In 

addition to maintaining the combined archives of the SBN and supporting a Web site, this 

group collects, formats, verifies, and consults on datasets concerned with comet 

observations. It also supports active comet missions and observing campaigns. 

The Asteroid/Dust Subnode is located at the Planetary Science Institute in Tucson, 

Arizona. The subnode collects, formats, verifies, and reviews ground-based and mission 

data pertaining to asteroids, trans-neptunians, small planetary satellites, and 

interplanetary dust. 

Further information is available at http://pdssbn.astro.umd.edu/. 
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DRAFT Pre-Decisional Material 239 

Appendix K. Object Characteristics Useful for Mitigation 

The physical properties of NEOs are extremely diverse, creating an added challenge to 

mitigation planning. Upon discovery, a NEO is an unresolved point of light whose 

motion reveals its orbit. Basic measurements of colors and albedos are needed for at least 

a preliminary assessment of composition and size. Sizes of currently known Earthintersecting objects range from 10 meters to 9 km. 

Additional measurements can reveal whether a NEO may be single or binary (or more), 

with components orbiting one another or joined to a single mass by gravity or mechanical 

means. About 15% are expected to be multiple-body orbiting systems by commonly 

accepted estimates; however, some estimates are as high as more than 50%. [102] [103] 

NEO interiors are likely highly fractured or porous, with little or no tensile strength. The 

likelihood of having some tensile strength is greatest for bodies below about 180 meters 

in diameter. Most NEOs (estimated 85%) likely originated in the main asteroid belt and 

up to 15% may have originated as comets that now appear inactive. 

The diversity of asteroid and comet compositions, as well as the diversity of meteorites, 

suggests NEOs are made of a wide range of materials. Current results for meteorite 

correlations suggest that NEO compositions range from low-density rocky material to 

high-density nickel-iron. Spin rates are typically a few hours, but can be orders of 

magnitude greater or smaller. 

At the smallest scales, NEO surfaces are thought to be boulders strewn with some 

regolith that is generally coarser than the regolith found on the Moon. Impacts generate 

seismic shaking, which force the fine dust to settle into smooth “ponds.” Seismic shaking 

also may erase small craters and fill in cracks or fractures that might be present at the 

surface. [104] 

A more detailed discussion of some of the asteroid characteristics important for 

evaluating deflection alternatives follows. 

K.1. Mass of Potentially Hazardous Asteroids 

The mass of the PHO is the most important quantity necessary to mitigate the threat 

because the mass affects the basic design and sizing of the impactor, explosive, or slow 

push payload, and therefore the mass that must be launched from Earth. The mass is the 

product of the NEO’s volume and density. Because NEO reflectivity can vary from about 

2% to 50%, simply measuring brightness and assuming reflectivity will result in a 

diameter estimate with a factor-of-two uncertainty. This leads to a factor-of-eight 

uncertainty in the volume. Asteroid densities are believed to vary from about 2- to 7-

g/cm3

. Adopting a mean density introduces another factor-of-two uncertainty into the 

mass. Thus, the total uncertainty in the momentum and impact energy, due to the 

individual uncertainties in the diameter and density, can be as large as a factor of 16 if 

only NEO detection-related means are employed. 
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240 DRAFT Pre-Decisional Material

The taxonomic class of an asteroid can be determined from visual photometry. This 

constrains the mineralogy of the object and reduces the uncertainty in the density 

estimate. Size estimates based on visual photometry and infrared radiometry are accurate 

within 10%. These observations reduce the total mass uncertainty to about 50%. [106] 

K.2. Orbits of Hazardous Asteroids 

A very large percentage (more than 99.9%) of PHO detections are expected to be 

asteroids. In conjunction with mass, the orbit of a threatening object is the most important 

information for designing a deflection alternative. This section discusses the statistical 

distribution of asteroid orbits, derived from the JPL potentially hazardous asteroid (PHA) 

database [71], a set that is assumed to be representative of all PHAs 140 meters and 

larger. While the complexity of orbital mechanics makes it difficult to generalize the orbit 

transfer energy (ΔV) and the time required for a set as diverse as this, studying three 

parameters provides some insight into the difficulty of reaching potential threats before 

they reach the Earth. 

K.2.1. Potentially Hazardous Objects Orbit Periods

Orbital period measures the time an object takes to travel around the Sun — the longer 

the orbital period, the farther the mean distance from the Sun. That means it will take the 

object longer to rendezvous or intercept. Figure 81 illustrates the statistical distribution of 

PHA orbital periods. [71] 

0

1

2

3

4

5

6

7

8

9

10

0 1 2 3 4 5 6 7 8 9 10

Period (years)

% PHAs

orbital (yrs)

synodic (yrs)

From JPL Online 

NEO Database Distribution =

2.3 ± 1.1 years 

Figure 81. Statistical Distribution of PHA Orbit Periods 

K.2.2. Orbit Inclination

The angle between the object’s orbit and the plane of Earth’s orbit determines orbit 

inclination. If the angle is large, more energy is required to travel to the object. Figure 81 

illustrates the statistical distribution of PHA orbit inclinations. [71] 
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DRAFT Pre-Decisional Material 241 

0

1

2

3

4

5

6

7

8

9

10

0 10 20 30 40 50 60 70 80 90

Inclination (deg)

% of Objects

NEOs

PHAs

From JPL Online 

NEO Database

PHA Distribution =

12.6 ± 11.1 degrees 

Figure 82. Statistical Distribution of PHA Orbit Inclinations 

K.2.3. Potentially Hazardous Objects Orbit Eccentricities

Figure 81 illustrates the statistical distribution of PHA orbit eccentricities. 

0

1

2

3

4

5

6

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1

Eccentricity

% of Objects

NEOs

PHAs

From JPL Online 

NEO Database

Figure 83. Statistical Distribution of PHA Orbit Eccentricities 

Orbit eccentricity is measured between 0 and 1 and represents how circular an orbit is. 

For an eccentricity of 0, an object stays equidistant from the Sun and travels in a circle. 

Objects with increasing eccentricity travel in increasingly more elliptical orbits. These 

objects travel much faster near the Sun than when they are farther away, and this 
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242 DRAFT Pre-Decisional Material

variation can affect the required orbit transfer ΔV and impact velocities and intercept 

angles. Note that PHA eccentricities vary more than those of NEOs. [71] 

K.3. Density of Potentially Hazardous Asteroids 

The density of the NEO provides additional information on the internal structure of the 

object and helps to define whether the object is “either a porous or volatile-rich rock, or a 

solid rock.” [57] This property also might help determine the response of the PHO to the 

mitigation attempt. For example, a solid body is more likely to withstand a high-velocity 

impact by a kinetic impactor or a blast of a nuclear explosive than a porous body. Also, 

as was evidenced by the Deep Impact mission [105], volatiles within a body could add 

significantly to the effectiveness of some techniques. They may increase the efficiency of 

the momentum transfer of the ejecta. 

Some impulsive techniques may also fragment a PHO, a concern from the perspectives 

that: 1) It would be more difficult to design back-up techniques to deal with several large 

fragments; 2) Fragments could be a greater threat to Earth than the single object [109]; 

and 3) The fragmentation event could create a debris cloud, making verification of the 

effectiveness difficult. Reference [104] notes that the size of the asteroid is important in 

estimating the likelihood of fragmentation. This reference estimates that asteroids larger 

than 1 km in diameter are the weakest from this perspective. 

V

~7%

C 

~9%

S ~ 48%

Q 

~32%

M < 1%

Other ~4%

Figure 84. Approximate Distribution of Asteroid Types 

Figure 84 shows current estimates for the distribution of the asteroid types described in 

Table 49. [101] Note that the density estimates for asteroids can vary by a factor of about 

four due to material composition (grain density), with additional uncertainty in mass due 

to porosity. The table illustrates the most commonly made associations between 

taxonomic classes and meteorites when other detailed information is unavailable. The 

reference stresses that taxonomy is not composition and that this table and figure 

represent only a preliminary estimate, again in the absence of more detailed information. 
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DRAFT Pre-Decisional Material 243 

Table 49. Generalization of Taxonomic Classes and Densities 

Taxonomic 

Class 

Generally Inferred 

Meteorite Association1 

Meteorite 

Densitites2 

C Carbonaceous chondrites 2.2-3.7 

S 

Diverse: Primitive achondrites 

ordinary chondrites differentiated 

olivine/ pyroxene / metal assemblages 

2 – 5 ? 

M Iron or Stony Meteorites 7 - 8 

Q Ordinary chondrites 3.5 – 3.9 

V Basaltic achondrites 3.1 – 3.8 

 1 Caveat emptor. Subject to ambiguities 

 2 These are grain densities (g cm-3) assuming zero porosity 

If interception of a threatening PHO occurs very close to Earth, disruption of the objects 

into small fragments may be a more desirable outcome than no action at all. However, the 

cloud of small fragments would impart the same amount of energy as the original object 

to Earth and the atmosphere, and as Reference [110] notes, such a cloud could cause 

“extreme damage” to orbiting satellites and the climate. 

K.4. Taxonomic Classes and Asteroid Types 

The composition of asteroids shown in Table 49 is drawn from recovered meteorites, 

providing an indirect link between spectral class, taxonomic class, and meteorite data. 

Since these classes do not describe the porosity, number of diverse bodies, elasticity, 

surface structure, or ratio of constituent materials, they are only of moderate use for 

deflection alternatives that require more information than mass and orbit. 

Figure 85. Infrared Signatures of Asteroid Types 
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244 DRAFT Pre-Decisional Material

It is possible that a different taxonomy from those used by scientists may prove useful for 

developing deflection alternatives, but such a new structure is not suggested here. Instead, 

it is assumed that the scientific asteroid types are representative of the types of 

information useful for differentiating mitigation approaches. 

Infrared measurements may distinguish asteroid types by using the signatures seen in and 

correlating those to meteoroid samples. They are shown in Figure 85. 

K.5. Porosity and Efficient Transfer of Momentum 

Figure 86 shows that the momentum transferred per unit mass of the impactor can vary 

three orders of magnitude depending on the porosity of the targeted PHO. The ejecta 

efficiency in the chart is related to the impact efficiency coefficient (β), developed in the 

discussion of the kinetic impact alternative. [29] 

Figure 86. Momentum Exchange Efficiency Depends on Porosity 

K.6. Rotation Rates of Near-Earth Asteroids 

Rotation rate also may be important for some deflection techniques, primarily those that 

require a soft landing, thrusting for long periods, or long-term operation in very close 

proximity to the PHO. In addition, the rotation rate may help estimate the likelihood that 

the object is a binary or a rubble pile. For impulsive techniques, a high spin rate may 

make it difficult to target a favorable location for imparting momentum change. 
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DRAFT Pre-Decisional Material 245 

Absolute Magnitude

10

Approximate Diameter, D (m)

12 16 20 24 28

Rotation Period, P (hours)

0.01

0.1

1

10

100

1000

10000 1000 100

Mechanically

Bound

(“solid”)

Gravity

Bound

(“rubble”)

N=388

Apparent Monolith-toRubble Pile Transition

at D~180m, P~2.2hr

Absolute Magnitude

10

Approximate Diameter, D (m)

12 16 20 24 28

Rotation Period, P (hours)

0.01

0.1

1

10

100

1000

10000 1000 100

Mechanically

Bound

(“solid”)

Gravity

Bound

(“rubble”)

N=388

Apparent Monolith-toRubble Pile Transition

at D~180m, P~2.2hr

Figure 87. Rotation Rates as compared with Diameter - Fast Rotation Barrier 

Figure 87, developed from Reference [107], shows the distribution of 388 observed 

objects according to rotation rates and diameter. This figure shows two thresholds that 

separate asteroids into small mechanically bound “solid” objects and larger 

gravitationally bound “rubble piles.” 

If representative of the general population of asteroids of this size, Figure 87 shows that a 

large percentage of asteroids 180 meters and smaller have rotation rates of under 2 hours. 

Due to the power law relationship between the number of NEOs and diameter, there are 

expected to be 100 times more 140-meter objects than those with diameters of 1 km. If 

proven, this assertion indicates that any of the deflection alternatives that require 

maintaining contact with the target object will operate in a very dynamic and changing 

environment. Operations, such as communications, power generation, and maintaining 

thrust direction, will be very challenging, if not infeasible for some alternatives. 

On almost all asteroids smaller than about 180 meters, the spin rate is so high that a 

surface package would not remain bound except at the poles. The asteroid can still be a 

rubble pile, due to simple frictional effects, but will be rotating so fast (period faster than 

a couple of hours) that a landed package would simply drift into orbit unless somehow 

tethered. Thus, landing surface packages on a PHO up to about 200 meters in diameter 

requires that the rotation rate be well characterized and that it be slower than a critical 

rate, defined as: 

ωcrit = (GM/r)0.5 (14) 
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246 DRAFT Pre-Decisional Material

where r is the radius and M is the asteroid’s mass. Asteroids larger than about 1 km 

almost all rotate slower than this rate, so simple surface anchoring is possible. 

One of the benefits of a fast rotation is that it makes the objects potentially easier to 

disrupt. A relatively small cratering charge on the equator of a rapidly rotating asteroid 

might send a lot of mass to escape velocity because all it has to do is shake the material 

loose, not accelerate it. This has not been studied in any detail. But if a powerful seismic 

event (blast event) shakes the rubble loose, then dynamical friction instead of static 

friction applies, and the asteroid will go into a landslide-type reconfiguration of mass. 

Whatever is moving faster than escape velocity (e.g. the equatorial regions) might be 

permanently lost. 
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Appendix L. Launch Capability for Characterization and Mitigation 

L.1. Launch Capability (C3) of Current and Planned Launch Systems 

The critical parameter needed to match launch capability with the ability to intercept (less 

so for rendezvous) an asteroid with a certain payload by a certain time (flight time) is the 

C3. It is equal to twice the specific (per unit mass) orbital energy, and has units of km2

/s2. 

Figure 88 shows the launch C3, as compared with payload capability for a number of past 

and current launch systems. Note how quickly payload decreases as launch C3 

requirements increase. Deflection options for impulsive techniques have generally been 

sized at C3 = 25 km2

/s2. 

0

20

40

60

80

100

1000 10000 100000

Payload Mass Launched (kg)

C3 (km2/s2)

Ares V

Atlas IIAS

Shuttle / IUS

Atlas V 551

Ariane V

Delta IV H

Delta IV H/STAR 48

Figure 88. C3 of Current and Planned Launch Vehicles 

L.2. Launch Capability Required 

To ascertain both feasible mission designs as well as warning-to-deflection timelines, one 

needs to determine the minimum launch C3 and transit time required to travel to a PHA 

of interest. The C3 computation was performed for two scenarios. 

1. The first sets the launch date and allows the asteroid to vary in its orbit. This 

might represent a mission scenario where the spacecraft is required to the leave 

the Earth by a fixed date. This might be the case for precursor characterization 

missions or deflections of an imminent threat. 

2. The second scenario allows both the launch date and the arrival date to vary. This 

might represent a precursor mission that would characterize a PHA well in 

advance of its projected time of closest approach with the Earth. For this type of 
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2006 Near-Earth Object Survey and Deflection Study 

248 DRAFT Pre-Decisional Material

precursor mission, there would be more flexibility in the launch and arrival dates 

and the mission could wait until a more favorable relative geometry was achieved. 

For each of these, the minimum launch C3 was computed for three conditions: the C3 

required to launch the spacecraft onto an intercept course with the PHA without regard to 

the arrival conditions; rendezvous with the PHA (minimization of the sum of the launch 

and arrival speed C3); and minimization of the arrival speed without concern for the 

launch C3. Since the goal was to produce a statistical summary of the C3 requirements, 

the minimum C3 for each PHA was found over the search space. In addition, a 

cumulative histogram was generated to show what percentage of the PHA population 

could be reached for each of the scenarios as a function of C3. 

L.3. Launch Energy (C3) for Varied and Fixed Launch or Arrival Dates 

The results of the Apophis case for a fixed launch date are shown in Figure 11 for the 

launch-to-intercept condition. The spikes in the plot reflect the times when the Lambert 

solution switched from going the “long way around” the transfer orbit to Apophis to the 

“short way” and back. For each PHA, data such as Figure 89 would be produced and the 

minimum launch C3 would be saved. The launch-to-intercept, rendezvous, and minimum 

arrival speed C3 would not necessarily be the same, as different transfer trajectories could 

produce different optimal launch C3. 

0

1000

2000

3000

012345

Time of Flight (years)

C3 (km /s ) 2 2

Figure 89. Required launch C3 for Apophis for a Fixed Launch Date 

Figure 90 and Figure 91 both show the percentage of the PHAs reachable for a given 

launch C3. In Figure 90 the launch date is fixed to intercept the PHA within 5 years. The 

C3 values computed for the population are relatively high. For example, the Ares V lunar 

cargo launch vehicle has a planned capability of launching a payload of 55,000 kg to a 

C3 of 25 km2

/s2. This indicates it could only intercept about 40% of the PHAs should the 

launch date be a constraint on the mission (i.e., a short warning situation). For rendezvous 

cases such as required by slow push techniques, 27% would be reachable. 
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DRAFT Pre-Decisional Material 249 

In Figure 91, both the launch date and the arrival date at the PHA vary to permit an 

optimal transfer. Much lower transfer velocities are possible if transit time is less of a 

factor. 

0

20

40

60

80

100

0 50 100 150 200

C3 (km /s )

% of Objects

Ares V

(CaLV)

Example

Launch to Intercept

Launch to Rendezvous

Minimum Arrival Speed

2 2

Figure 90. Launch C3 for Fixed Launch Date 

0

20

40

60

80

100

0 20 40 60 80 100 120 140 160 180 200

C3 (km /s )

% of Objects

Ares V

(CaLV)

Example

Launch to Intercept

Launch to Rendezvous

Minimum Arrival Speed

2 2

Figure 91. Launch C3 Launch and/or Arrival Dates Vary 

An additional solution was examined wherein an angular constraint of 15o was placed 

upon the approach angle to the PHA. This constraint would represent the end-point 

requirement for a kinetic impactor that needed to either push or slow the PHA by 

directing its force as close as possible to the PHA’s pre-existing velocity vector. 
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2006 Near-Earth Object Survey and Deflection Study 

250 DRAFT Pre-Decisional Material

Figure 92 shows the result for the launch-to-intercept condition with and without the 

angular constraint for a fixed launch date. For the Ares V, imposing a 15o constraint 

reduces the PHAs that could be reached from about 40% to about 25%. For rendezvous, 

the percentage of reachable PHAs would drop from 27% to 18%. If other constraints on 

the mission are added, such as line-of-sight communications between the spacecraft and 

the Earth during mission critical events, the impact on the C3 requirements could be even 

more pronounced. Figure 93 shows that many more potential threats are reachable if 

launch and/or arrival dates are allowed to vary. 

0

20

40

60

80

100

0 50 100 150 200

C3 (km /s )

% of Objects

Ares V

(CaLV)

Example

No Intercept Constraint on Launch

15 degree Intercept Constraint on Launch

2 2

Figure 92. Realistic Intercept Constraints – Fixed Launch Date 

0

20

40

60

80

100

0 50 100 150 200

C3 (km /s )

% of Objects

Ares V

(CaLV)

Example

No Intercept Constraint on Launch

2 2

15 degree Intercept Constraint on Launch

Figure 93. Realistic Intercept Constraints – Variable Launch/Arrival Date 
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2006 Near-Earth Object Survey and Deflection Study 

DRAFT Pre-Decisional Material 251 

Similar results as those shown in Figure 90 and Figure 92 are generated (within 10-15% 

of ΔV for a given C3) if the arrival time is fixed but the launch date is allowed to vary. 

As many NASA missions have shown, swing-by trajectories using the Earth or other 

celestial bodies may reduce transfer ΔV in some circumstances. Swing-by trajectories 

also are launch-date dependent and are likely not applicable to short-warning launches, 

but would be studied for the preponderance of threats that will have long warning 

durations. Since flyby trajectories are case dependent and analysis intensive, they were 

considered outside the scope of this study, except to note their value in producing 

minimum ΔV transfers. 

L.4. Launch Capability Required to Intercept Comets 

The launch capability required to intercept comets is discussed in the context of the 

scenario described in Section 6.13.6. 

L.5. Launch Capability Summary and Conclusions 

The conclusions of this section are manifold. C3 requirements are large (compared with 

current and planned launch capability) for characterization or mitigation missions when 

launch or arrival time is unconstrained and a favorable geometry between the Earth and 

the PHO cannot be attained. For realistic launch constraints, fewer than 25% of potential 

threats can be reached with the planned capabilities of the Ares V. 

For characterization missions without a time constraint, or deflection missions with 

sufficiently long warning, C3 values are lower and within current and expected launch 

capabilities. Since the synodic periods of the majority of PHAs is on the order of a few 

years, warning times of several decades will allow for selection of favorable geometries 

and thus avoid the worst case crisis situation. 
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252 DRAFT Pre-Decisional Material

Appendix M. Additional Cost Results 

The following sections display the results of developing the life-cycle cost (LCC) figure 

of merit (FOM). The sections include the LCC for the individual concepts, followed by 

combinations of concepts that collectively form an architecture. 

M.1. Detection and Tracking Survey Element Costs 

Table 50 displays a further breakout of the space-based survey systems in FY06$M. 

Figure 94 shows a breakout of development, launch, and operations costs through 2020. 

Table 50. Space-Based Survey Systems Life-cycle cost Breakout 

Concept D-8A D-8B D-9 D-10 D-11 D-12 D-13 D-14

Aperture 2.0 m 2.0 m 1.0 m 2.0 m 0.5 m 1.0 m 0.5 m 1.0 m

Band Visible Visible Visible Visible IR IR IR IR

Orbit LEO L1/L2 Venus Venus L1/L2 L1/L2 Venus Venus

PM/SE/MA 102 $ 102 $ 91 $ 117 $ 37 $ 57 $ 41 $ 63 $ 

Flight System 476 $ 476 $ 507 $ 582 $ 170 $ 279 $ 200 $ 318 $ 

Payload 274 $ 274 $ 161 $ 276 $ 100 $ 140 $ 101 $ 140 $ 

Pre-launch GDS 112 $ 112 $ 100 $ 128 $ 40 $ 62 $ 45 $ 68 $ 

Development Reserve 394 $ 395 $ 298 $ 406 $ 87 $ 133 $ 88 $ 147 $ 

Total Development 1,357 $ 1,358 $ 1,157 $ 1,508 $ 434 $ 671 $ 474 $ 736 $ 

Launch Vehicle 95 $ 132 $ 132 $ 159 $ 95 $ 107 $ 107 $ 132 $ 

MO&DA + Reserve 329 $ 329 $ 330 $ 328 $ 341 $ 336 $ 340 $ 335 $ 

Total Mission Cost 1,781 $ 1,819 $ 1,619 $ 1,995 $ 870 $ 1,114 $ 921 $ 1,204 $ 

$0.0 $0.5 $1.0 $1.5 $2.0

0.5m IR L1/L2

1.0m IR L1/L2

0.5m IR Venus-Like

Orbit

1.0m IR Venus-Like

Orbit

2.0m Visible LEO

2.0m Visible L1/L2

1.0m Vis Venus-Like

Orbit

2.0m Vis Venus-Like

Orbit

Detection, Tracking, and Cataloguing - Lifecycle Cost ($B 2006)

Development

Launch

Operations

Figure 94. Breakout of Search Element Costs 
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DRAFT Pre-Decisional Material 253 

M.2. Characterization Element Results Costs 

Figure 95 displays the LCC through 2020 in FY06$B for the characterization concepts 

and Figure 96 breaks out the costs for the space-based systems. 

$0.1

$0.1

$0.2

$0.5

$0.1

$0.2

$1.8

$1.8

$2.0

$0.9

$1.0

$1.1

$1.2

$0.5

$0.5

$0.8

$0.7

$1.0

$0.9

0.0 0.5 1.0 1.5 2.0 2.5

Shared DCT

Shared LSST

PS 8

PS 16

Dedicated LSST

Dedicated IRTF

Bistatic Radar 100m

2.0m Visible + Filters LEO

2.0m Visible + Filters L1/L2

2.0m Visible Venus-Like

0.5m Multi-Channel IR L1/L2

0.5m MC IR Venus-Like

1.0m Multi-Channel IR L1/L2

1.0m MC IR Venus-Like

Flyby Electric Propulsion

Flyby Chemcial Propulsion

Flyby (Chemical) w/Impactor

Orbiter

Orbiter with Lander

Orbiter with Penetrator

Characterization - Lifecycle Cost ($B 2006)

Figure 95. Characterization Alternatives Life-cycle cost Results 

$0.0 $0.5 $1.0 $1.5 $2.0 $2.5

2.0m Visible + Filters LEO

2.0m Visible + Filters L1/L2

2.0m Visible Venus-Like Orbit

0.5m Multi-Channel IR L1/L2

0.5m MC IR Venus-Like Orbit

1.0m Multi-Channel IR L1/L2

1.0m MC IR Venus-Like Orbit

Flyby Electric Propulsion

Flyby Chemcial Propulsion

Flyby (Chemical) w/Impactor

Orbiter

Orbiter with Lander

Orbiter with Penetrator

Characterization Elements Lifecycle Cost ($B 2006)

Development

Launch

Operations

Figure 96. Breakout of Space-based Characterization Element Costs 
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2006 Near-Earth Object Survey and Deflection Study 

254 DRAFT Pre-Decisional Material

Table 51 displays a further breakout of the space-based remote sensing Characterization 

systems in FY06$M 

Table 51. Space-Based Remote Characterization Systems LCC Breakout 

Concept C-8A C-8B C-09 C-16 C-17 C-18 C-19

Aperture 2.0 m 2.0 m 2.0 m 0.5 m 0.5 m 1.0 m 1.0 m

Band Visible FW Visible FW Visible FW IR MC IR MC IR MC IR MC

Orbit LEO L1/L2 Venus L1/L2 Venus L1/L2 Venus

PM/SE/MA 109 $ 109 $ 118 $ 40 $ 43 $ 59 $ 64 $ 

Flight System 519 $ 519 $ 584 $ 188 $ 213 $ 291 $ 322 $ 

Payload 277 $ 277 $ 277 $ 102 $ 102 $ 142 $ 143 $ 

Pre-launch GDS 119 $ 119 $ 128 $ 43 $ 47 $ 65 $ 69 $ 

Development Reserve 362 $ 363 $ 307 $ 72 $ 77 $ 125 $ 147 $ 

Total Development 1,386 $ 1,386 $ 1,413 $ 444 $ 483 $ 682 $ 746 $ 

Launch Vehicle 95 $ 132 $ 159 $ 95 $ 132 $ 107 $ 132 $ 

MO&DA + Reserve 329 $ 329 $ 330 $ 342 $ 341 $ 337 $ 335 $ 

Total Mission Cost 1,809 $ 1,847 $ 1,902 $ 880 $ 956 $ 1,126 $ 1,213 $ 

Table 52 shows a summary of space-based in-situ characterization system costs. 

Table 52. Space-Based In-Situ Characterization Systems LCC Breakout 

Concept C-21 C-22 C-23 C-24 C-25 C-26

Approach Flyby Flyby Flyby Orbiter Orbiter Orbiter

Descirption Electric Chemical Impactor Only Lander Penetrator

PM/SE/MA 23 $ 24 $ 47 $ 32 $ 57 $ 51 $ 

Flight System 87 $ 92 $ 267 $ 157 $ 290 $ 257 $ 

Payload 88 $ 88 $ 89 $ 89 $ 144 $ 128 $ 

Pre-launch GDS 19 $ 19 $ 38 $ 26 $ 47 $ 41 $ 

Development Reserve 54 $ 45 $ 120 $ 87 $ 135 $ 116 $ 

Total Development 272 $ 268 $ 561 $ 392 $ 674 $ 593 $ 

Launch Vehicle 117 $ 117 $ 117 $ 117 $ 142 $ 142 $ 

MO&DA + Reserve 75 $ 78 $ 74 $ 198 $ 193 $ 194 $ 

Total Mission Cost 464 $ 463 $ 752 $ 706 $ 1,009 $ 930 $ 

M.3. Detailed Deflection Element Cost Results 

Table 53 contains a breakout of deflection system elements included in the life-cycle 

costs. These estimates include nominal costs for an appropriately sized Delta launch 

vehicle; the costs for Ares V at higher flight rates have not been determined. 

Table 53. Deflection Systems Life-Cycle Cost Breakout 

Concept M-1 M-2 M-3A M-3B M-4 M-5 M-7

Description Nuclear 

Standoff

Nuclear 

Surface

Conventional 

Surface

Conventional 

SurfaceDelay

Kinetic 

Impactor Space Tug Gravity 

Tractor

PM/SE/MA 169 $ 162 $ 161 $ 161 $ 66 $ 486 $ 568 $ 

Flight System 1,166 $ 1,115 $ 1,105 $ 1,105 $ 421 $ 3,320 $ 3,979 $ 

Payload 50 $ 50 $ 50 $ 50 $ 51 $ 174 $ 111 $ 

Pre-launch GDS 192 $ 184 $ 182 $ 182 $ 74 $ 551 $ 645 $ 

Development Reserve 531 $ 487 $ 507 $ 507 $ 134 $ 1,294 $ 1,769 $ 

Total Development 2,108 $ 1,998 $ 2,006 $ 2,006 $ 747 $ 5,825 $ 7,073 $ 

Launch Vehicle 233 $ 233 $ 220 $ 220 $ 132 $ 233 $ 233 $ 

MO&DA + Reserve 112 $ 112 $ 112 $ 112 $ 113 $ 559 $ 557 $ 

Total Mission Cost 2,452 $ 2,342 $ 2,337 $ 2,337 $ 992 $ 6,616 $ 7,862 $ 
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M.4. Characterization Architecture Cost Results 

The figures in the section display the annual funding requirements in RY$M for each of 

the following characterization architectures: 

Option Descriptions 

Option 1 Use Existing Assets + Detection and Tracking Systems 

Option 2 O1 + Dedicated Ground Systems 

Option 3 O1 + Dedicated Space-Based Remote Sensing (L1/L2) 

Option 4 O1 + Dedicated Space-Based Remote Sensing (Venus-Like Orbit) 

Option 5 O1+ O2+ O3 + 2 Flyby Missions 

Option 6 O1 + O2 + O3 + 8 Orbiter Missions 

Option 7 O1 + O2 + O3 + Orbiter Mission at Fixed P(I) Threshold 

Table 54 shows the buildup of average life-cycle cost for each characterization capability 

option. 

Table 54. Life-cycle costs of Characterization Capability Options 

Nominal 

Development 

Time 

(Parallel 

Developments)

Lifecycle Cost 

(Through 2020)

Architecture Months $B 2006

Option 1: Use Existing Assets + Detection and Tracking Systems

Existing Optical + Shared DCT + Shared LSST 72 0.7 $ 

Existing Radars + Shared DCT + Shared LSST 67 0.2 $ 

Option 1 Average 70 0.5 $ 

Option 2: Dedicated Ground Systems

Existing Radars + Shared DCT + Shared LSST + Dedicated PS8 73 0.3 $ 

Existing Radars + Shared DCT + Shared LSST + Dedicated PS16 73 0.4 $ 

Existing Radars + Shared DCT + Shared LSST + Dedicated LSST 80 0.7 $ 

Existing Radars + Shared DCT + Shared LSST + Dedicated IRTF 66 0.3 $ 

Existing Radars + Shared DCT + Shared LSST + Dedicated Bistatic Radar 99 0.4 $ 

Option 2 Average 78 0.4 $ 

Option 3: Option 1 + Dedicated Space-Based Remote Sensing

Existing Radars + Shared DCT + Shared LSST + 2m Vis LEO w/Filters 86 2.0 $ 

Existing Radars + Shared DCT + Shared LSST + 2m Vis L1/L2 w/Filters 86 2.1 $ 

Existing Radars + Shared DCT + Shared LSST + 2m Vis Venus-Like Heliocentric Orbit w/Filters 88 2.3 $ 

Existing Radars + Shared DCT + Shared LSST + 0.5m Multi-Channel IR L1/L2 65 1.1 $ 

Existing Radars + Shared DCT + Shared LSST + 0.5m Multi-Channel IR Venus-Like Heliocentric Orbit 63 1.2 $ 

Existing Radars + Shared DCT + Shared LSST + 1.0m Multi-Channel IR L1/L2 72 1.4 $ 

Existing Radars + Shared DCT + Shared LSST + 1.0m Multi-Channel IR Venus-Like Heliocentric Orbit 71 1.4 $ 

Option 3 Average 76 1.6 $ 

Option 4: Option 1+ Option 2+ Option 3 + 2 Flyby Missions

Existing Radars + Shared DCT + Shared LSST Dedicated IRTF + 2 SEP Flyby Missions 67 2.0 $ 

Existing Radars + Shared DCT + Shared LSST Dedicated IRTF + 2 Flyby Missions 67 2.0 $ 

Option 4 Average 67 $ 2.0

Option 5: Option 1 + Option 2 + Option 3 + 8 Orbiter/Flyby Missions

Existing Radars + Shared DCT + Shared LSST + Dedicated IRTF + 0.5m Multi-Channel IR + 8 Orbiters 67 5.2 $ 

Existing Radars + Shared DCT + Shared LSST + Dedicated IRTF + 0.5m Multi-Channel IR + 8 Flybys 67 3.7 $ 

Existing Radars + Shared DCT + Shared LSST + Dedicated IRTF + 0.5m Multi-Channel IR + 8 Flybys/Impactors 69 4.9 $ 

Option 5 Average 68 4.6 $ 

Option 6: Option 1 + Option 2 + Option 3 + 8 Orbiter/Lander Missions

Existing Radars + Shared DCT + Shared LSST + Dedicated IRTF + 0.5m Multi-Channel IR + 8 Orbiter/Landers 68 6.4 $ 

Existing Radars + Shared DCT + Shared LSST + Dedicated IRTF + 0.5m Multi-Channel IR + 8 Orbiter/Penetrator 69 6.1 $ 

Option 6 Average 69 6.3 $ 
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Appendix N. Roles of Advanced Systems and Technologies 

N.1. Detection, Tracking, Cataloguing, Characterization Systems 

N.1.1. Advanced Follow-up Systems and Technologies

Radar measurements are useful for rapidly improving the accuracy of orbital information 

particularly during a close approach to Earth. Radar also can determine other PHO 

characteristics, including shape. Arecibo and Goldstone are limited to 0.3 AU for 

tracking and 0.1 AU for characterization. Radar signal strength decreases proportionally 

to the distance to the target (R) to the fourth power for radar signals (∝ R4

), making it 

very difficult to significantly extend the signal reach beyond what current systems can 

provide. Furthermore, increasing power beyond current levels will not significantly 

increase performance due the Earth’s atmosphere. It is unlikely that a more capable radar 

will developed, and no technologies have been identified to improve effective radar 

ranging. 

On the contrary, proposed telescopes with extremely large apertures will significantly 

affect both characterization and tracking accuracy. Based on the increase in aperture, the 

amount of light reaching a detector will increase by orders of magnitude. The increase in 

flux will allow more detailed spectroscopic measurements using smaller bandpasses. 

Some of the largest projects are predicting that with adaptive optics they will be 

diffraction limited and have resolution on the order of hundreds of micro-arc seconds, 

offering improved positional accuracy with every observation. In addition, recent 

advances have created ever-larger detectors that have a wide dynamic range, smaller 

pixels, large well depths, low dark current, low noise, fast readout, and very high 

quantum efficiency. These technologies will lead to improvements in follow-up and 

characterization systems. 

Currently, several operational or nearly complete 8-meter or larger telescopes are coming 

on line, as shown in Table 55. Designed for astronomical imaging and/or spectroscopy, 

they, therefore, have small fields of view. 

Table 55. Planned or Existing Large Observatories 

Observatory 

Effective 

Aperture Location(s) 

Gemini 8.1 m Mauna Kea, Hawaii 

Cerro Pachon, Chile 

Very Large Telescope 8.2 m Cerro Paranal, Chile 

Subaru 8.3 m Mauna Kea, Hawaii 

Hobby-Eberly 9.2 m Mount Fowlkes, Texas 

South African Large Telescope 9.8 m Sutherland, South Africa 

Keck 10.0 m Mauna Kea, Hawaii 

Large Binocular Telescope 11.8 m Mount Graham, Arizona 
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Two approaches currently are available for obtaining a larger aperture: using a large 

number of small segments or combining the light from several telescopes. A number of 

telescopes with effective apertures significantly larger than 8 meters are under 

construction or in a significant design stage, and some of these are shown in Table 56. 

None of these systems is designed to have a wide field of view required by a survey 

mission, but they will likely add considerably to the ability to follow-up and characterize 

newly discovered objects. 

Table 56. Proposed Very Large Observatories 

Observatory 

Effective 

Aperture Location(s) 

Gran Telescope Canarias 10.4 m La Palma, Canary Islands 

Keck Interferometer 14.6 m Mauna Kea, Hawaii 

Very Large Telescope 16.4 m Cerro Paranal, Chile 

Giant Magellan Telescope 21.4 m Chile? 

Even as these proposed telescopes are progressing, astronomers are planning the next 

class of ground-based telescopes, collectively considered Extremely Large Telescopes. 

They are listed in Table 57. Their funding and planning remain in flux due to the many 

technical challenges that they must be overcome. 

Table 57. Postulated Extremely Large Observatories 

Proposed Observatory 

Effective 

Aperture Proposed Location 

Atacama Telescope Project 25 m Cerro Chajnantor, Chile 

Thirty Meter Telescope 30 m To be determined 

Maximum Aperture Telescope 30-50 m To be determined 

Euro50 50 m La Palma, Canary Islands 

Overwhelmingly Large Telescope 100 m To be determined 

The Atacama Telescope Project is intended to observe very long wavelengths, ranging 

from 100 microns into the sub-millimeter. Longer wavelengths should lead to improved 

characterization of asteroid composition and structure. The Thirty Meter Telescope 

project resulted from the merger of several 30-meter class projects each of which had run 

into technical or funding issues. 

Using an infrared bandpass for the follow-up and characterization of PHOs is desirable. 

PHOs tend to have low albedo and will radiate in the infrared; there are twice as many 

solar photons available in that band as contrasted with the visible. In addition, solar phase 

angle dominates the apparent visible brightness of solar system objects, but plays little 

part in the infrared. 

In summary, the ability to follow-up observations of asteroids to more accurately 

determine position, composition, existence of companions, rotation, and of asteroids 

using optical and infrared means can be expected to greatly increase over the foreseeable 

future. 
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N.1.2. Advanced Survey Systems and Technologies

Increasing the rate of detection of PHOs required for surveys relies primarily on two 

variables: the amount of light reaching the detector (telescope aperture) and observation 

geometry. Therefore, few technologies make a significant difference in NEO surveys 

other than increasing the amount of light available by increasing size of the aperture. 

The size of near infrared detectors has progressed from 58 x 62 pixels to 2048 x 2048 

pixels in a little over a decade; optical detectors are commercially available in much 

larger sizes, over 10k x 10k. An area of very active research today is to continue 

increasing the spatial extent of detectors with an immediate industry goal of creating a 

monolithic IR detector of 4096 x 4096 pixels. One of the greatest challenges is 

maintaining high uniformity across such a large area. These larger detectors will lead to 

more capable space-based systems, but at increased cost and spacecraft complexity. 

While the sensors on planned systems such as the James Webb Space Telescope may 

meet some of the sensitivity requirements for PHO detection, these systems are generally 

ill-suited to survey operations which need to cover larger areas of sky. 

Future technology may allow for cheaper and more reliable space missions but will not 

fundamentally change the need to view from different angles for a PHO survey. Aside 

from the other benefit of circumventing atmospheric absorption, the difficulty, cost, and 

risk of space missions, particularly of launching very large optics and radiation damage to 

detectors leaves little room for growth barring a breakthrough in launch vehicles or 

deployable mirrors. The size of the Hubble Space Telescope (2.4 m), which itself is not 

suitable to survey or follow-up operations, appears to be at the practical limit for spacebased optical systems. 

Beyond the current 8-meter class of ground-based telescopes, proposed systems are not 

expected to have a large field of view, except the LSST. The steady increase in telescope 

aperture does not generally correspond to an increase in étendue, which is the limiting 

factor in optical system throughput. Therefore, the field of view for these extremely large 

telescopes will not increase appreciably from that of typical astronomical observatories 

without overcoming difficult optical design challenges. To avoid untenable growth in the 

size of the telescope enclosures; very fast optical systems are required leading to 

formidable difficulties keeping stray light off of the massive optical elements. This, in 

turn, requires short integration when observing in visible light, which partially negates 

the gain in aperture and leads to a reduction in telescope performance. 

In parallel with ground state of the art, there are a number of large aperture (3.5 – 6 

meters) space-based astronomical IR telescopes which are either on-orbit or will be by 

the close of the decade: Spitzer, Space Infrared Telescope for Cosmology and 

Astrophysics (SPICA), the Single Aperture Far-Infrared Observatory (SAFIR), and the 

James Webb Space Telescope (JWST). These telescopes share several key aspects in 

common: they will operate far from the Earth, either near L2 or in an Earth-trailing orbit; 

they are actively cooled, some to as low as 5.5 K to observe nearly the entire infrared 

spectrum, and have narrow fields of view. As detailed in another section of this study, 

detection rate of asteroids can be greatly increased if significantly different viewing 

geometries are available, e.g. a detector located in a Venus-like orbit. Future technology 

will likely allow for cheaper and more reliable space missions but will not fundamentally 

change the need to view from different angles for a NEO survey. Aside from the other 
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benefit of circumventing atmospheric absorption, the difficulty, cost, and risk of space 

missions, particularly of launching very large optics and radiation damage to detectors 

leaves little room for growth barring a breakthrough in launch vehicles or deployable 

mirrors. 

Cooling requirements have been relaxed somewhat for optical detectors but cooling of 

infrared detectors and optics remains necessary so that the instrument does not create its 

own warm photons that would mask the desired incoming light. Factors that directly 

affect the temperature at which a detector must be cooled are the composition of the 

detector (material and doping), the bias, and the cutoff wavelength; longer wavelengths 

require greater cooling. For instance, undoped silicon-based detectors have a bandpass of 

roughly 0.4 – 1.1 microns and can be operated at temperatures from 140o to 300o Kelvin. 

As the operating temperature is increased, the red response suffers and the dark current 

increases exponentially, resulting in noisier frames. Doped silicon, particularly the 

impurity-band conductor arsenic-doped silicon, can reach wavelengths of 26 microns but 

must be cooled to 10 K. If cooling were not an issue, this would be an example of a 

highly appropriate detector for a survey mission as it can be fabricated into fairly large 

arrays. 

Material Typical Bandpass (μm) Operating Temperature (oK) 

Si 0.35 – 1.1 140 – 300 

Doped Si < 26 10 

Ge 1.1 – 2.3 140 – 300 

Doped Ge < 25 25 

InGaAs 1.1 – 1.7 77 – 140 

InSb 2.2 – 5.5 65 – 77 

HgCdTe 5.5 – 12 20 – 300 

In the optical, quantum efficiency can be greater than 80% over much of the visible 

spectrum and into the near infrared; strictly in the infrared, HgCdTe chips routinely 

achieve 60% efficiency over a wider bandpass. Quantum efficiency is a function of 

wavelength and may drop off significantly before reaching either end of the bandpass 

depending on detector doping and fabrication. Since the quantum efficiency is generally 

so high, particularly in the optical, the only practical method of achieving lower threshold 

flux is to increase aperture. These improvements are not expected to lead to revolutionary 

survey system capabilities, but are part of the evolution of optical sensor technologies. 

Two areas where infrared detector research is actively conducted are bolometers and 

quantum-well infrared photodetectors (QWIPS). These are unlikely to provide additional 

capabilities to survey or follow-up systems. Bolometers have a few advantages. They can 

be run at fairly warm temperatures, can be fabricated into fairly large detectors, and have 

very long cutoff wavelengths that extend to the microwave. However, little solar energy 

is reflected back at these long wavelengths. Their poor sensitivity exacerbates this. 

QWIPS has appeared to be a promising detector technology for the infrared, as the 

bandpass is fairly tunable during fabrication. The resulting bandpass is usually quite 

narrow and the detector sensitivity is roughly two orders of magnitude less than silicon, 

for instance. 
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In summary, advanced technologies are not expected to increase sensitivity over the 

currently planned 8-meter ground systems, and no planned systems larger than 8 meters 

are expected to be capable of a field of view useful for next-generation surveys. While 

space-based systems of the size proposed in this study appear feasible, these systems are 

approaching the practical limit of aperture size and, therefore, survey performance. 

N.2. In-Situ Mitigation 

N.2.1. Technologies

As discussed in recent NASA reports, advanced technologies may substantially increase 

the number of alternatives available for asteroid or comet deflection. [26] [27] A 

summary of these advanced technologies follows. 

• High-thrust, high-specific impulse propulsion systems (such as plasma or nuclear) 

would deliver orbit modification systems to target NEOs and may provide the 

necessary efficiency to enable or significantly improve additional deflection 

alternatives. 

• Multi-megawatt to gigawatt-class electrical power systems are required for many 

advanced propulsion and laser applications. 

• Advanced thermal management systems are critical to reject large amounts of 

waste heat developed during the operation of many advanced deflection systems. 

• Reliable, high-power pulsed laser ablation systems with adaptive laser optics, 

precision beam width focusing, and closed-loop control system would provide 

continuous orbit modification capability. Systems also could be potentially used 

as an active ranging system for precision orbit determination. 

• Advanced autonomous or semi-autonomous rendezvous and station-keeping 

capability that engage the NEO at close distances are also necessary for many 

advanced concepts. In addition, formation-flying capability and precise attitude 

control may be needed for interferometry using orbiting detectors. 

• Development of neutron or X-ray nuclear explosives may be considerably more 

effective than conventional devices. They can shed the crust of some types of 

asteroids, thereby providing additional efficiency in the transfer of explosive 

energy to PHO momentum. 

• Lightweight materials for solar sails and inflatable structures might enable highpower systems, space tug alternatives, or allow hazardous objects to be 

“wrapped” in order to change thermal and electrical forces on the objects. 

N.2.2. Suggested Studies

This analysis of the deflection alternatives suggests several areas where additional 

research and technology verification may be warranted to enhance the ability to mitigate 

a possible threat. Key issues raised in this study could be resolved through the following 

additional studies: 

• Understand the timing and key decisions that must be made during the evaluation 

of a potential threat. As noted in this chapter, if a credible threat is identified in 

the next decade or two (while a precision orbit catalog is developed), it is likely 

that a decision to proceed will be based on imperfect information. If this transient 
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problem is a concern, studies of the critical steps in the decision-making process 

would provide a decision tree to be followed in response to a credible threat. 

Perhaps a “war games” simulation would provide some useful insights. 

• Study intercept-trajectory design further. As noted in some of the scenarios 

developed in this study, finding intercept trajectories that are within an available 

launch capability and actually intercepting the object can be a significant 

challenge for missions with short warning times. Studies should be conducted to 

define the set of threat conditions where trajectory options are currently available 

and suggest alternatives for those that are not. 

• Understand the prevalence, characteristics, and hazard of companion bodies. It is 

known that NEOs can have companion bodies or moons, and it is possible that 

these bodies could be of sufficient size that they also threaten Earth. Studies are 

recommended to characterize the orbits of companion bodies. 
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Appendix O. Potential Synergy of Survey Systems with DoD 

A classified report examining the possibility of using NEO survey systems for 

Department of Defense (DoD) applications and applying DoD assets or data to the NEO 

survey mission was prepared for this study. However, all of the fundamental findings of 

this effort are unclassified and are presented in this unclassified appendix. 

While certain planned assets might, in principle, be able to contribute to the NEO 

detection problem, they could never perform this mission by themselves. In addition, the 

extensive NEO detection simulations carried out by JPL in recent months, with particular 

emphasis on Potentially Hazardous Asteroids (PHA), has clearly demonstrated that the 

incremental benefit would be small of almost all existing/planned assets identifying 90% 

of the 140-meter PHAs by 2020. 

One of the most important findings of the NASA 2006 NEO Survey Study is that a key 

asset (such as an LSST-class telescope) trumps all others, except another key asset with a 

different bandpass and/or geometrical perspective — hence, the powerful synergy that is 

found between an LSST-class, optical ground-based telescope, and a space-based IR 

sensor at the L1 point, or the even more powerful synergy between a space-based optical 

sensor at Venus' orbit with a space-based IR one at either the L1 point or in a Venus orbit. 

No currently planned DoD sensor assets analyzed can provide much serendipitous 

support to key assets such as those. 

In the following sections, the potential utility and known or expected disadvantages of 

existing or planned assets are discussed. 

O.1. Contributions of DoD Systems to PHO Survey 

O.1.1. System A

This geographically distributed set of ground-based optical telescopes has been for many 

years one of the mainstays of the Space Surveillance Network for the detection and 

tracking of deep space targets. With its recent upgrade to advanced CCD cameras, it is 

now producing more tracks per day than ever before. However, because of limited 

aperture sizes, and the Concept of Operations (CONOPS) of their standard surveillance 

mode, these sensors are unable to approach the limiting magnitude required for the PHA 

detection problem. Comparably sized optical telescopes with CCD cameras currently 

being used for asteroid detection, such as those of the Catalina Sky Survey and Project 

Spacewatch, have much better limiting magnitudes after several minutes of integration 

time per single look. It would not be possible to significantly improve upon this limit of 

System A. 

O.1.2. System B

The primary mission and resultant CONPOS of System B preclude it from reaching the 

required limiting magnitude. In addition, its extremely fast optical design means that its 

limiting magnitude would be degraded by night-sky brightness even at longer integration 
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times. An analysis of comparable systems has shown that copies of this asset, with a 

different optical design and a different CONOPS, might possibly contribute to the PHO 

detection problem, although not with sufficient performance to meet the congressionally 

mandated 2020 deadline. This contribution would at best be similar to systems such as 

sharing PS4 or LSST. 

O.1.3. System C

This system has been the pathfinder for space-based surveillance systems. However, with 

its small aperture it could not contribute in any meaningful way to the PHO detection 

problem 

O.1.4. System D

This sensor is designed to replace System C and act as a second pathfinder for spacebased surveillance. However, despite its larger aperture and more advanced CCD focal 

planes, it could still only have a moderate PHO detection capability. Further, it is a single 

satellite system with a primary mission duty cycle that would preclude secondary tasking 

for NEO detection, other than purely serendipitous ones at rather bright limiting 

magnitudes. Duplicates of the space segment (with a new ground segment) are 

encompassed by other analyses in this report. 

O.1.5. System E

This concept may be a future, multiple-satellite constellation of sensors, each at least as 

capable as that of System D. The multiple satellites in the constellation might alleviate 

the asset availability issue. Furthermore, there is the possibility of a future technology 

upgrade from Si CCDs, with a bandpass of 0.35 - 1.1 μm, to InGaAs detectors, with an 

increased bandpass of 0.4 to 2 μm, thereby doubling the number of reflected solar 

photons available to the sensor. Nevertheless, these systems ultimately will still be 

aperture-, and thus, stare-time limited. Longer integration times place severe drift and 

jitter demand on the spacecraft and, along with the requirement for multiple looks, also 

on the duty cycle of the sensor constellation. These space systems could potentially be 

modified for the purpose of detection PHOs, but are not very capable in their current 

configurations for this task. 

In general, these space sensor systems would have to interleave PHO tasking with their 

standard mission planning. New search patterns and algorithms also would be required, 

along with appropriate data pathways. While it is possible that an integrated system could 

be developed, it is difficult to predict if the global benefit would warrant the 

demonstrated difficulty of working across classification and national mission boundaries. 

O.1.6. System F

This concept shares many of the attributes of System E, but it is unlikely that its sensors 

will actually be as capable in terms of limiting magnitude. 

O.1.7. Summary

In summary, all current and planned sensor assets studied have technological or 

operational limitations that will constrain their ability to contribute significantly to the 
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NEO search mission. This is even more so when such assets are compared with a groundbased LSST-class telescope or a space-based system. 

O.2. Possible Contributions of NEO Survey Systems to Other Missions 

A fully government-owned and operated LSST-class system would be able to contribute 

to other missions, although it would have concept of operations issues if a NEO search 

were its primary mission. Pan-STARRS1 (PS1) may have to deal with exactly this 

problem. However, the LSST design is not ideal for synoptic space surveillance (slew 

rate will be low and the telescope is too large). System B is a better design for synoptic 

space surveillance of deep space targets than LSST. An LSST could do targeted space 

surveillance (and SOI) for deep space. This would slow down the NEO search rate, 

however, and extend the 90% Survey completion schedule. 

It is very likely that none of the space-based concepts for NEO search can contribute at 

all to other missions — most are simply stationed too far from the Earth (L1 or Venus 

orbit). A large-aperture LEO optical asset might be ideal for these other missions, but is 

sub-optimal for the search for PHOs. Follow-up assets for NEOs have their own unique 

requirements, and generally do not lend themselves to other missions, except perhaps 

Space Object Identification (SOI). For example, given a large-aperture, ground-based IR 

telescope at high-altitude, the ability to share such an asset between NEO follow-up 

observations and SOI missions could work, although prioritization may be an issue. 

In summary, no asset proposed for this study meets any significant fraction of the DoD 

requirements when concept of operations and duty cycle are considered. 

O.3. Sharing of Archived Data 

In general, the DoD systems studied do not archive data sufficiently long enough to 

support activities such as precovery. Systems A and C may archive more data than others, 

but neither of these meets sensitivity requirements for precovery. In addition, assets 

planned for the near future will have the same data archiving limitations. Precovery is 

most effective in astronomy applications tailored for NEO surveys; that is, when large 

field of view photographic plates are used. As both the Survey program systems and 

future DoD systems are fielded, opportunities for sharing data between missions will be 

explored further. If opportunities for cooperation or collaboration are identified, these 

will be pursued as they are now, but advanced planning to facilitate these interactions 

may not be warranted. 
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Appendix P. Potential Synergies of Deflection with Exploration 

One approach to developing technologies and capabilities required for deflection is to 

develop and use deflection-related technologies as part of exploration missions. 

Techniques and hardware required for robotic systems and resource utilization on the 

Moon or landing large payloads on Mars, for example, relate directly to those needed for 

subsurface explosives, mass drivers, or a space tug. Developing confidence in the ability 

to utilize these common technologies reliably will lower the risk of using them for 

deflection missions. 
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Table 58 shows that a number of capabilities, including space transportation, rendezvous, 

autonomous systems, and highly robust solutions, are needed for many PHO 

characterization and deflection systems. They overlap with technologies required to 

explore the solar system. Other technologies, such as nuclear and conventional explosives 

and precision intercept, do not. Therefore, the Deep Impact missions and other 

opportunities might be used to develop confidence in these techniques. Large energy 

sources also appear to be a common need, and the availability of reliable sources of 

power would facilitate the evolution of slow push deflection approaches. 

Given the relatively poor performance of the slow push techniques, it does not appear 

that they provide a conclusive rationale for developing nuclear or other high specific 

impulse propulsion technologies. However, if these technologies were developed for 

Exploration or science, their use could be re-evaluated in the context of their 

demonstrated performance. 

There may be options for system-level demonstrations of the Crew Exploration Vehicle 

(CEV), Ares I (crew launch vehicle), Lunar Surface Access Module (LSAM), and/or 

Ares V (lunar cargo launch vehicle) to contribute to PHO deflection demonstrations. For 

example, in 2012 and 2013 several uncrewed launches of the Ares I will be conducted 

with the CEV’s Crew and Service Modules. These systems have considerable mass, 

precision rendezvous and docking capability, as well as considerable performance margin 

when the Service Module is fully loaded. A similar system launched at a passing PHO 

could provide a significant deflection demonstration. 

In addition, while the Ares V’s Earth Departure Stage does not have rendezvous 

capability, it does have sufficient performance to launch a large demonstration vehicle for 

many of the deflection alternatives. It is possible that the demonstration of parts of the 

lunar system could be directed toward an asteroid deflection to improve the reliability of 

lunar system elements. 
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Table 58. Characterization and Deflection Technology Needs Compared with 

Exploration 
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Appendix Q. Potential Synergies of Deflection with the DoD 

Q.1. Overview 

The U.S. Air Academy SPACECAST 2020 report, which was published in 1994, 

discusses the surveillance of objects that could potentially impact the Earth and examines 

ways to counter threats through various deflection techniques. This Air Force report takes 

the position that asteroids are potential threats that the DoD should evaluate and prepare 

to defend against. The report also discusses benefits of a DoD role in an international 

effort and provides some specific recommendations. 

However, unless significant funding is allocated to developing and testing asteroid 

deflection technique(s), any future asteroid deflection campaign will be a “come as you 

are” affair. It will involve whatever assets and infrastructure are available at the time, 

probably on a global scale. 

Q.2. Launch Systems 

Launch requirements for specific asteroid deflection approaches are described in the body 

of this report. In general, the minimum launch requirements for asteroid deflection 

include large-lift capacity (6000 kg or more into high-energy, interplanetary trajectories), 

multiple interceptor launches during infrequent and very short windows of opportunities 

(a few weeks at most), and long transit times (up to many years) prior to an autonomous 

terminal guidance phase to collide or rendezvous with the target asteroid. 

The Delta 4 Heavy is the largest launcher available in the U.S. and is operated primarily 

for the U.S. Air Force. NASA plans call for the development of a heavy-lift expendable 

launch vehicle for Moon and Mars exploration. Expected to be available by 2020, the 

vehicle may use RS-68 engines similar to those on the Delta 4 Heavy. 

With several years of warning, the Delta 4 may be able to launch a nuclear explosive that 

could deflect a percentage of the potentially threatening objects, with sizes ranging up to 

several hundred meters in diameter. But the Delta 4 cannot be surge launched in rapid 

succession during short windows, which this report suggests to assure launch success and 

redundant intercept attempts. Also, the Delta 4 has only one launch site, which is subject 

to weather delays. 

Until a more capable (larger and higher surge flight rate) is available, an asteroid 

deflection campaign will require using the heaviest launch vehicles available from many 

nations to ensure overall success. 

Although the DoD’s evolving launch requirements call for eventually developing a 

capability to responsively launch tactical payloads in rapid succession, these systems will 

not provide the heavy-lift capability required for asteroid deflection. Recent studies have 

not been able to find a stronger linkage than currently exists between DoD requirements 

for small “launch on need” capabilities, DoD and NASA requirements for “commercial” 

launch Earth orbit on schedule services, NASA’s requirement for infrequent “launch on 

schedule” interplanetary missions, and a capability to support a possible asteroid 
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deflection (many rapid very heavy interplanetary launches once every 5,000 years on 

average). 

In 1996, The Aerospace Corporation analyzed a wide range of futuristic space 

applications (civil, military, and commercial) and their launch requirements to see if 

dramatic cost reductions might be possible if missions emerged. As part of this study 

Aerospace developed a Space Transportation Economic Index (STEI) — a numerical 

scoring method to compare the feasibility of each futuristic application’s spacelift 

requirements with current spacelift capabilities and cost. 

The index considers weight to orbit, flight rate, required launch cost (for economic 

feasibility of the mission), and manned as compared with unmanned reliability as 

variables in the evaluation, and included many types of commercial, civil, and military 

missions. An STEI score of 1.0 is typical of today’s commercial and military programs. 

Potential space applications with higher STEI scores require significant reduction in 

launch costs to be economically feasible. 

Asteroid deflection (a launch application not considered in the original Aerospace study) 

has among the highest STEI scores recorded. Therefore, it is very unlikely that any future 

military, civil, or commercial application will emerge to justify the development and 

sustainment of a heavy-lift launch fleet with a surge launch capability that could be 

available for an asteroid deflection campaign. 

Launch services are, indeed, likely to be “come as you are” situations. 

Q.3. Deflection Campaign Operations 

Currently, no government agency or international organization has overall responsibility 

for deflecting asteroids to avert the threat of an impending collision, although NASA’s 

charter includes the goal of protecting Earth. Coordinating and de-conflicting deflection 

attempts by different countries will be required. Mission planning and execution 

procedures among missions will need to be established, probably by international 

agreement under United Nations auspices. Oversight by a U.S. military agency would 

likely not be acceptable to the world community, particularly if nuclear technologies 

were required, as is likely. 

If an asteroid threatened the Earth, it can be assumed that the U.S. government would 

suspend many of the treaty and ITAR constraints that currently discourage international 

cooperation. Such constraints would place severe limits on international cooperation. 

Q.4. Guidance Technology 

During the terminal guidance phase, the interceptor must autonomously guide itself 

toward the desired aim point on the surface of the asteroid at very high closing velocities. 

Rendezvous, on the other hand, occurs at very low closing relative velocity and is not 

technically as difficult as interception, provided there is sufficient fuel and maneuver 

authority to continuously correct guidance errors until it lands or achieves orbit. 

A unique aspect of the asteroid intercept problem is that the extended target may have an 

unknown, irregular shape. For missile defense interceptors, a relatively smaller target 

remains essentially a point image right up until impact, simplifying the guidance problem 
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significantly. The asteroid’s image will balloon in size on an optical tracking system’s 

focal plane during the terminal phase. The expanding edge could confuse methods 

developed for defense applications to the point that they think that the object has 

additional relative motion. DoD experience with homing guidance against large extended 

targets is limited and the asteroid impact problem would likely require different 

techniques. 

Additionally, radar and laser ranging can become ambiguous due to the size and terrain 

of an asteroid. This can affect terminal guidance and warhead fusing. The DoD has 

considerable experience with bomb fusing, but not at the high closing velocities seen in 

asteroid deflection. 

For kinetic and surface explosive deflection techniques, the desired aim point is through 

the object’s center of gravity. The center of gravity may be difficult to determine if the 

shape is irregular and made of multiple large masses. Characterization missions might be 

able to determine the shape of the object, but the exact attitude of the object at the time of 

impact also must be known to determine the exact aim point. Again, DoD experience 

with this sort of targeting problem is limited or non-existent, as targets are generally well 

characterized. 

The time delays and high closing velocity during terminal intercept maneuvers preclude 

man-in-the-loop commanding of an asteroid interceptor, as with TV-guided bombs. Many 

of the rendezvous and “slow push” operations will need to be autonomous due to 

communication time delays or the lack of communication altogether if the deflection 

system is out of Earth’s line of sight. 

The DoD has considerable experience with “fire and forget” cruise missiles, but flight 

times for these systems are only a few hours, not the years required for transit time or the 

months of “slow push” operations. NASA experience with interplanetary trajectories is 

likely more applicable to this problem than DoD’s. 

Since significant research and development and testing of sensor suites and guidance 

systems suitable for hitting the center of gravity of an extended, irregularly shaped 

asteroid target will be required, there is little synergy with DoD terminal guidance 

requirements and technologies. 

Q.5. Industrial Base 

Currently, all sectors of the space industry (military, civil and commercial) are struggling. 

The worldwide launch rate has been stagnant or declining for the last several years. In 

addition, fiber optics, the end of the Cold War, emergence of terrorist threats, and lack of 

enthusiastic support for human spaceflight make some less bullish on the future growth 

of space applications. The overall situation is pressuring the commercial and civil space 

sectors to consolidate and eliminate capabilities. Consequently, military space acquisition 

officials are concerned about the future health of the industrial base needed to support 

proposed advanced military programs. 

In light of these prospects, will adequate technologies and sufficient production and 

launch infrastructure exist in the future to support an asteroid deflection campaign? 
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As called for in the SPACECAST 2020 report, DoD cooperation with agencies planning 

for an asteroid deflection should be encouraged. The challenge is to translate public and 

congressional concerns about the potential threat of an asteroid hitting the Earth into a 

reinvigoration of the military and civil industrial base to meet such threats. This could be 

done while avoiding the security and international issues that a cooperative effort may 

create. Even in the event of an identified threat, the issues related to the classification of 

some potentially relevant technologies may make DoD participation difficult, particularly 

if the asteroid strike is not likely to affect the U.S. (which is likely for local or regional 

impact scenarios). 

The need to work with the international space community, as NASA does with the 

International Space Station, poses a severe constraint on DoD’s involvement in the 

asteroid deflection and technology development effort. Military space officials are always 

mindful that foreign governments may want to develop their own military space 

capabilities and see ventures with the U.S. as a way to gain access to U.S. technology. 

Q.6. Conclusion 

The DoD would obviously play important roles in any asteroid deflection campaign, 

especially if nuclear explosives are called for. But many aspects of asteroid deflection 

(surge heavy-lift launch, high-speed terminal guidance against large extended objects, 

autonomous operations over several years, international cooperation, and information 

sharing) have little or no synergy with DoD’s current or future requirements for space 

systems. 
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Appendix R. Role of Nuclear Explosives Technology 

As part of this study, a representative of the National Nuclear Safety Administration 

(NNSA) prepared a classified report on the role of nuclear technologies for the deflection 

of PHOs. An unclassified summary of this report appears in this appendix. 

R.1. Nuclear Explosive Design 

The principal assumption made about nuclear explosives is that their approximate yieldto-weight ratio in the range of tens of kilotons to tens of megatons is well known. It is 

further assumed that these devices are sufficiently well known and that they either exist 

or could be made available within 12 to 18 months. 

Within limits these assumptions are true. The process to construct nuclear explosives in 

the range of yields that have been tested and stockpiled by the U.S. is well known. It is 

probably safe to assume that the same holds true of Russia. There are some points that 

must be taken into consideration, however. They are described in detail in the classified 

report: 

• Current designs for nuclear explosives were not intended for use in deep space. It 

is unknown and unlikely that one could take a stockpile weapon and use it for 

PHO mitigation without any modification. However, nuclear explosives have 

been tested in space and there are no known impediments to their use for PHO 

deflection. The roles of DoD and DoE and the regulatory requirements are not 

clear in the event such an event became necessary. 

• While no nuclear testing would be needed to qualify a physics package for space 

applications, the necessary electromechanical components (arming, fusing, and 

firing) might need some modification and in-space qualification. 

• The U.S. no longer stockpiles multi-megaton warheads. If there were a need for a 

very high-yield device, there is ample design and test information to construct 

such a device given sufficient lead-time (12-18 months). 

• Enhanced Radiation Systems (nuclear explosives designed for generating a large 

fraction of energy in the form of neutrons, for example) have been designed and 

tested by the U.S. Explosives of this type for NEO deflection and mitigation could 

be produced without requiring testing. 

• Specific designs can and would be designed for optimal launch and potential reentry safety. This includes using the most advanced available technologies to 

ensure that the device could not be used as a weapon if lost.