Steven R. Oleson and Melissa L. McGuire

Glenn Research Center, Cleveland, Ohio

COMPASS Final Report: Near Earth 

Asteroids Rendezvous and Sample 

Earth Returns (NEARER)

NASA/TM—2009-215825

September 2009

CD–2008–28
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Steven R. Oleson and Melissa L. McGuire

Glenn Research Center, Cleveland, Ohio

COMPASS Final Report: Near Earth 

Asteroids Rendezvous and Sample 

Earth Returns (NEARER)

NASA/TM—2009-215825

September 2009

CD–2008–28

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

Near Earth Asteroids Rendezvous and Sample Earth Returns (NEARER)

sponsored by the 31st International Electric Propulsion Conference

Ann Arbor, Michigan, September 20–24, 2009
Acknowledgments

The Collaborative Modeling for Parametric Assessment of Space Systems (COMPASS) analysis team members include 

Les Balkanyi, Laura Burke, Tony Colozza, John Dankanich, Jon Drexler, Doug Fiehler, James Fincannon, 

John Gyekenyesi, Joe Hemminger, Jeff Juergens, Tom Krivanek, Nelson Morales, T.C. Nguyen, Tom Packard, 

Tom Parkey, Carl Sandifer, Tim Sarver-Verhey, Paul Schmitz, Anita Tenteris, Joe Warner, and Jeff Woytach.

Available from

NASA Center for Aerospace Information

7115 Standard Drive

Hanover, MD 21076–1320

National Technical Information Service

5285 Port Royal Road

Springfi eld, VA 22161

Available electronically at http://gltrs.grc.nasa.gov

Trade names and trademarks are used in this report for identifi cation 

only. Their usage does not constitute an offi cial endorsement, 

either expressed or implied, by the National Aeronautics and 

Space Administration.

Level of Review: This material has been technically reviewed by technical management. 

This report is a formal draft or working 

paper, intended to solicit comments and 

ideas from a technical peer group.

This report contains preliminary fi ndings, 

subject to revision as analysis proceeds.

This report is a preprint of a paper intended for presentation at a conference. 

Because changes may be made before formal publication, this preprint is made available 

with the understanding that it will not be cited or reproduced without 

the permission of the author.
NASA/TM—2009-215825 iii 

Contents 

1.0 EXECUTIVE SUMMARY ............................................................................................................. 1

2.0 STUDY BACKGROUND AND ASSUMPTIONS ......................................................................... 2

2.1 Introduction ......................................................................................................................... 2

2.2 Assumptions........................................................................................................................ 2

2.3 Growth, Contingency and Margin Policy ........................................................................... 3

2.4 Redundancy Assumptions ................................................................................................... 4

2.5 Mission Description ............................................................................................................ 4

2.5.1 Mission Analysis Assumptions .............................................................................. 4

2.5.2 Mission Analysis Analytic Methods ...................................................................... 6

2.5.3 Mission Analysis Event Timeline .......................................................................... 6

2.5.4 Mission Trajectory Details ..................................................................................... 7

2.6 Concept of Operations (CONOPS) ..................................................................................... 7

2.7 Launch Vehicle Details ....................................................................................................... 8

2.7.1 Launch Vehicle Trade-Space Relative Performance ............................................. 8

2.7.2 Atlas 401 (4 m Fairing) Performance .................................................................... 9

2.7.3 Atlas 401 Payload Fairing Details ....................................................................... 10

2.8 Launch Vehicle Packaging ............................................................................................... 11

2.9 Sample Return Capsule (SRC) System Level Summary .................................................. 11

2.9.1 Sample Collection Requirements ......................................................................... 11

2.10 Basic Science Payload Description ................................................................................... 12

2.10.1 Arm Mounted (~1 m) Instruments ....................................................................... 13

2.10.2 Body Mounted Instruments ................................................................................. 13

2.11 Super Science Package Description .................................................................................. 13

3.0 BASELINE DESIGN—CASE 1 ................................................................................................... 15

3.1 Top Level Design (MEL and PEL)................................................................................... 15

3.1.1 Master Equipment List (MEL) ............................................................................ 15

3.1.2 Power Equipment List (PEL) ............................................................................... 16

3.2 Baseline System Level Summary ..................................................................................... 17

3.3 Baseline Design Concept Drawing and Description ......................................................... 18

3.4 Baseline Design Concept Dimensions .............................................................................. 21

4.0 CHALLENGES, LESSONS LEARNED, AREAS FOR FUTURE STUDY ................................ 22

4.1 Future Work ...................................................................................................................... 22

4.2 Lessons Learned ............................................................................................................... 22

5.0 SUBSYSTEM BREAKDOWN ..................................................................................................... 22

5.1 Communications ............................................................................................................... 22

5.1.1 Communications Requirements ........................................................................... 22

5.1.2 Communications Assumptions ............................................................................ 23

5.1.3 Communications Design and MEL ...................................................................... 23

5.1.4 Communications Trades ...................................................................................... 24

5.1.5 Communications Analytical Methods .................................................................. 24

5.1.6 Communications Risk Inputs ............................................................................... 27

5.1.7 Communications Recommendation ..................................................................... 27

5.2 Guidance, Navigation and Control (GN&C) .................................................................... 28

5.2.1 GN&C Requirements ........................................................................................... 28

5.2.2 GN&C Assumptions ............................................................................................ 28

5.2.3 GN&C Design and MEL ..................................................................................... 28

5.2.4 GN&C Trades ...................................................................................................... 29

5.2.5 GN&C Analytical Methods ................................................................................. 29

5.2.6 GN&C Risk Inputs............................................................................................... 30
NASA/TM—2009-215825 iv 

5.3 Command and Data Handling (C&DH) ........................................................................... 30

5.3.1 C&DH Requirements ........................................................................................... 30

5.3.2 C&DH Assumptions ............................................................................................ 30

5.3.3 C&DH Design and MEL ..................................................................................... 31

5.3.4 C&DH Trades ...................................................................................................... 31

5.3.5 C&DH Analytical Methods ................................................................................. 31

5.3.6 C&DH Risk Inputs............................................................................................... 31

5.3.7 C&DH Recommendation ..................................................................................... 31

5.4 Electrical Power System ................................................................................................... 31

5.4.1 Power Requirements ............................................................................................ 31

5.4.2 Power Assumptions ............................................................................................. 31

5.4.3 Power Design and MEL ....................................................................................... 33

5.4.4 Power Trades ....................................................................................................... 35

5.4.5 Power Analytical Methods ................................................................................... 35

5.4.6 Power Risk Inputs ................................................................................................ 35

5.4.7 Power Recommendation ...................................................................................... 35

5.5 Structures and Mechanisms .............................................................................................. 35

5.5.1 Structures and Mechanisms Requirements .......................................................... 35

5.5.2 Structures and Mechanisms Assumptions ........................................................... 35

5.5.3 Structures and Mechanisms Design and MEL ..................................................... 35

5.5.4 Main S/C Bus Design .......................................................................................... 35

5.5.5 SRC Design ......................................................................................................... 37

5.5.6 SRC MEL ............................................................................................................ 38

5.5.7 Structures and Mechanisms Trades ..................................................................... 39

5.5.8 Structures and Mechanisms Analytical Methods ................................................. 39

5.5.9 Structures and Mechanisms Risk Inputs .............................................................. 39

5.5.10 Structures and Mechanisms Recommendation .................................................... 40

5.6 Propulsion and Propellant Management ........................................................................... 40

5.6.1 Propulsion and Propellant Management Requirements ....................................... 40

5.6.2 Propulsion and Propellant Management Assumptions ........................................ 40

5.6.3 Propulsion and Propellant Management Design and MEL .................................. 41

5.6.4 Propulsion and Propellant Management Trades .................................................. 44

5.6.5 Propulsion and Propellant Management Analytical Methods .............................. 44

5.6.6 Propulsion and Propellant Management Risk Inputs ........................................... 48

5.6.7 Propulsion and Propellant Management Recommendation ................................. 48

5.7 Thermal Control ................................................................................................................ 48

5.7.1 Thermal Requirements ......................................................................................... 48

5.7.2 Thermal Assumptions .......................................................................................... 48

5.7.3 Thermal Design and MEL ................................................................................... 49

5.7.4 Thermal Trades .................................................................................................... 50

5.7.5 Thermal Analytical Methods ............................................................................... 50

5.7.6 Thermal Risk Inputs ............................................................................................. 52

5.7.7 Thermal Recommendation ................................................................................... 52

6.0 SOFTWARE COST ESTIMATION ............................................................................................. 53

6.1 Objectives ......................................................................................................................... 53

6.2 Assumptions...................................................................................................................... 53

6.2.1 SW Sizing Assumptions ...................................................................................... 53

6.2.2 S/C Bus Functions ............................................................................................... 53

6.2.3 Payload Functions ................................................................................................ 53

6.3 Software Cost .................................................................................................................... 53

7.0 COST, RISK AND RELIABILITY ............................................................................................... 54
NASA/TM—2009-215825 v 

7.1 Costing .............................................................................................................................. 54

7.1.1 Detailed Cost Breakdown by WBS...................................................................... 55

7.1.2 System Integration Wraps Defined ...................................................................... 56

7.2 Risk Analysis and Reduction ............................................................................................ 57

7.2.1 Assumptions ........................................................................................................ 57

7.2.2 Risk List ............................................................................................................... 57

7.2.3 Risk Summary ..................................................................................................... 57

7.3 Reliability.......................................................................................................................... 57

8.0 TRADE SPACE ITERATIONS .................................................................................................... 58

8.1 Case 1—Baseline .............................................................................................................. 59

8.2 Case 1a .............................................................................................................................. 60

8.3 Case 2 ................................................................................................................................ 61

8.4 Case 3 ................................................................................................................................ 62

8.5 Case 3a .............................................................................................................................. 63

9.0 REFERENCES .............................................................................................................................. 64

APPENDIX A— ACRONYMS AND ABBREVIATIONS ....................................................................... 65

APPENDIX B— RENDERED DESIGN DRAWINGS ............................................................................. 67

List of Figures 

Figure 1.1—NEARER Case 1a Concept Vehicle Design ............................................................................. 1

Figure 2.1—4660 Nereus Body Orbital Details ............................................................................................ 5

Figure 2.2—1996 FG3 Body Orbital Details ................................................................................................ 5

Figure 2.3—Trajectory Main Mission Details .............................................................................................. 6

Figure 2.4—S/C Distance From the Sun and Earth Over Mission Time ...................................................... 7

Figure 2.5—4660 Nereus Trajectory Plot ..................................................................................................... 7

Figure 2.6—Mission Main Events Trajectory Graphic ................................................................................. 8

Figure 2.7—Performance Curves for Launch Vehicles of Interest ............................................................... 9

Figure 2.8—Atlas V 421 Performance Curve. ............................................................................................ 10

Figure 2.9—Atlas 401 4 m LPF Payload Fairing ELV ............................................................................... 10

Figure 2.10—NEA Sample Return S/C—Atlas 401 Fairing Packaging Close View ................................. 11

Figure 2.11—NEA Sample Return S/C—Atlas 401 Fairing Envelop and Packaging View ...................... 11

Figure 2.12—Example Science Sampler Collection Arm .......................................................................... 12

Figure 3.1—NEA Sample Return S/C—Top Components View ............................................................... 19

Figure 3.2—NEA Sample Return S/C—Bottom Components View ......................................................... 19

Figure 3.3—NEA Sample Return S/C—Deployed View ........................................................................... 20

Figure 3.4—NEA Sample Return S/C—Stowed View ............................................................................... 20

Figure 3.5—NEA Sample Return S/C—Footprint Dimensions ................................................................. 21

Figure 3.6—NEA Sample Return S/C—Deployed Dimensions ................................................................. 21

Figure 5.1—NEA Sample Return S/C—Communications Instruments ..................................................... 24

Figure 5.2—Worst Case Nereus Link Budget Analysis ............................................................................. 25

Figure 5.3—Case1 Nereus Link Budget Analysis ...................................................................................... 26

Figure 5.4—Case 2 Nereus Link Budget Analysis ..................................................................................... 26

Figure 5.5—Case 3 Nereus Link Budget Analysis ..................................................................................... 27

Figure 5.6—Case 4 Nereus Link Budget Analysis ..................................................................................... 27

Figure 5.7—Mapping Orbit Delta V Assumptions ..................................................................................... 28

Figure 5.8—GN&C Subsystem Block Diagram ......................................................................................... 29

Figure 5.9—Solar Pressure Torque Geometry ............................................................................................ 30

Figure 5.10—Thrust as a Function of Gimbal Angle ................................................................................. 30

Figure 5.11—Orion-Derived UltraFlex SA Design .................................................................................... 32

Figure 5.12—Example Li-Ion Battery, Built for the Mars Phoenix Lander (Similar Batteries are Planned 

for Orion) ....................................................................................................................................... 32
NASA/TM—2009-215825 vi 

Figure 5.13—Power Available From SA Over Mission Time ................................................................... 33

Figure 5.14—Power System Schematic ...................................................................................................... 34

Figure 5.15—Main S/C Bus Analytical Design .......................................................................................... 36

Figure 5.16—SRC concept design .............................................................................................................. 37

Figure 5.17—Sample Collection Arm ........................................................................................................ 38

Figure 5.18—Example Graphic of Main Structure ..................................................................................... 39

Figure 5.19—Main Propulsion System ....................................................................................................... 41

Figure 5.20—Notional Vehicle Landing Propulsion System ..................................................................... 41

Figure 5.21—RCS Propulsion system schematic ....................................................................................... 42

Figure 5.22—PM 1 Performance Acceptance Test ..................................................................................... 45

Figure 5.23—Gimbal Performance Test and CAD Illustration .................................................................. 46

Figure 5.24—NEXT Thruster PPU schematic ............................................................................................ 46

Figure 5.25—NEXT Propellant Management System................................................................................ 47

Figure 5.26—Hydrazine Tank (Left), Xe Tank (Right) .............................................................................. 48

Figure 5.27—Schemaic of the Louver Prototype ....................................................................................... 51

Figure 5.28—MLI ....................................................................................................................................... 51

Figure 6.1—Typical RAD750 Processor and Software Environment ........................................................ 53

Figure B.1—Rendered Baseline Case 1—Deployed View ......................................................................... 67

Figure B.2—Rendered Baseline Case 1—Close View of Main Body ........................................................ 67

Figure B.3—Transparent Baseline Case 1—Main Bus View 1 .................................................................. 68

Figure B.4—Transparent Baseline Case 1—Main Bus View 2 .................................................................. 68

List of Tables 

Table 1.1—Mission and Spacecraft (S/C) Summary—Baseline Case 1 ...................................................... 2

Table 2.1—Study Assumptions .................................................................................................................... 3

Table 2.2—Percent Mass Growth Allowance ............................................................................................... 4

Table 2.3—Atlas V family Performance ...................................................................................................... 9

Table 2.4—Science Package Portion of the MEL ...................................................................................... 12

Table 2.5—APL Science Payload ............................................................................................................... 13

Table 2.5—NEARER Baseline Science Payload Details (Adapted From Applied Physics Laboratory 

(APL) Science Payload) ................................................................................................................. 14

Table 2.6—Super Science Package Portion of the MEL ............................................................................ 14

Table 3.1—Master Equipment List—Case 1 .............................................................................................. 15

Table 3.2—System Integration Summary—Case 1 .................................................................................... 16

Table 3.3—Power Estimations and Waste Heat Over Mission Modes....................................................... 17

Table 5.1—Communications System Requirements .................................................................................. 23

Table 5.2—Communications MEL for Baseline (Case 1) .......................................................................... 24

Table 5.3—Link Budget Calculations for Communications Scenarios ...................................................... 25

Table 5.4—GN&C MEL (Case 1) .............................................................................................................. 29

Table 5.5—C&DH MEL for Baseline (Case 1) .......................................................................................... 31

Table 5.6—Electrical Power System MEL for Baseline (Case 1) .............................................................. 34

Table 5.7—Structures and Mechanical Systems MEL for Baseline (Case 1) ............................................ 36

Table 5.8—SRC MEL (Case 1) .................................................................................................................. 38

Table 5.9—NEXT Thruster Options Technology Assumptions ................................................................. 40

Table 5.10—Propulsion and Propellant System MEL for Baseline (Case 1) ............................................. 43

Table 5.11—Thermal MEL for Baseline Case 1 ........................................................................................ 49

Table 5.12—Thermal System Data Exchange ............................................................................................ 50

Table 5.13—Thermal System Radiator Sizing Assumptions ..................................................................... 50

Table 5.14—Thermal System Tank Insulation Sizing Assumptions .......................................................... 52

Table 5.15—Thermal System Tank Insulation Sizing Assumptions .......................................................... 52

Table 5.16—Thermal System PMAD Cooling Sizing Assumptions .......................................................... 52
NASA/TM—2009-215825 vii 

Table 6.1—Software Costing estimate ....................................................................................................... 54

Table 7.1—NEXT 1+1, Added Science, Case 1a, LCC Cost (FY09 $M) .................................................. 54

Table 8.1—Spacecraft Case Comparison ................................................................................................... 58

Table 8.2—Spacecraft Total Mass Comparison ......................................................................................... 58

Table 8.3—Case 1—System Summary ...................................................................................................... 59

Table 8.4—Case 1—Sample Return Craft Summary ................................................................................. 59

Table 8.5—Case 1a—System Summary ..................................................................................................... 60

Table 8.6—Case 1a—Sample Return Craft Summary ............................................................................... 60

Table 8.7—Case 2—System Summary ...................................................................................................... 61

Table 8.8—Case 2—Sample Return Craft Summary ................................................................................. 61

Table 8.9—Case 3—System Summary ...................................................................................................... 62

Table 8.10—Case 3—Sample Return Craft Summary ............................................................................... 62

Table 8.11—Case 3a—System Summary ................................................................................................... 63

Table 8.12—Case 3a—Sample Return Craft Summary ............................................................................. 63

 
NASA/TM—2009-215825 viii 

STUDY PARTICIPANTS 

Near Earth Asteroid Sample Return (NEARER) Design Session 

Subsystem Name Center Email 

Customer In Space Propulsion GRC 

COMPASS Team 

COMPASS Team Lead Steve Oleson GRC Steven.R.Oleson@nasa.gov

System Integration, MEL and 

Final Report Documentation Melissa McGuire GRC Melissa.L.Mcguire@nasa.gov

Ground Systems, Launch 

Vehicle Integration Jeff Woytach GRC Jeffrey.M.Woytach@nasa.gov

Mission Design, ACS, GN&C 

and Operations Doug Fiehler GRC Douglas.I.Fiehler@nasa.gov

Mission Design John Dankanich GRC john.dankanich@nasa.gov

Final documentation and 

editing Les Balkanyi GRC Leslie.R.Balkanyi@nasa.gov

Operations and Timeline Carl Sandifer GRC Carl.E.Sandifer@nasa.gov

Mission Design and 

Operations Laura Burke GRC Laura.M.Burke@nasa.gov

Propulsion Tim Sarver–Verhey GRC Timothy.R.Verhey@nasa.gov

Structures and Mechanisms Nelson Morales GRC Nelson.Morales-1@nasa.gov

Structures and Mechanisms Tom Krivanek GRC Thomas.M.Krivanek@nasa.gov

Mechanical Systems John Gyekenyesi GRC john.z.gyekenyesi@nasa.gov

Thermal Tony Colozza GRC Anthony.J.Colozza@nasa.gov

Power Josh Freeh GRC joshua.e.freeh@nasa.gov

Configuration Tom Packard GRC Thomas.W.Packard@nasa.gov

C&DH, Software TC Nguyen GRC Thanh.C.Nguyen@nasa.gov

Communications Joe Warner GRC Joseph.D.Warner@nasa.gov

Cost Tom Parkey GRC Thomas.J.Parkey@nasa.gov

Cost Jon Drexler GRC Jonathan.A.Drexler@nasa.gov

Risk/Reliability Anita Tenteris GRC Anita.D.Tenteris@nasa.gov
NASA/TM—2009-215825 1 

1.0 EXECUTIVE SUMMARY 

In this study, the Collaborative Modeling for Parametric Assessment of Space Systems (COMPASS) team 

completed a design for a multi-asteroid (Nereus and 1996 FG3) sample return capable spacecraft for the 

NASA In-Space Propulsion Office. The objective of the study was to support technology development and 

assess the relative benefits of different electric propulsion systems on asteroid sample return design. The 

design uses a single, heritage Orion solar array (SA) (~6.5 kW at 1 AU) to power a single NASA 

Evolutionary Xenon Thruster ((NEXT) a spare NEXT is carried) to propel a lander to two near Earth

asteroids. After landing and gathering science samples, the Solar Electric Propulsion (SEP) vehicle 

spirals back to Earth where it drops off the first sample’s return capsule and performs an Earth flyby to 

assist the craft in rendezvousing with a second asteroid, which is then sampled. The second sample is 

returned in a similar fashion. The vehicle, dubbed Near Earth Asteroids Rendezvous and Sample Earth 

Returns (NEARER), easily fits in an Atlas 401 launcher and its cost estimates put the mission in the New 

Frontier’s (NF's) class mission. 

Table 1.1 collects the details of the subsystems at a top level in the baseline design (case 1). 

Figure 1.1—NEARER Case 1a Concept Vehicle Design 
NASA/TM—2009-215825 2 

Table 1.1—Mission and Spacecraft (S/C) Summary—Baseline Case 1 

Subsystem area Details Total mass 

with growth

Top level system SEP enables sample returns from both Nereus and 1996 FG3 Asteroids 1352 kg 

(wet with 

growth) 

Mission, 

operations 

7 yr mission: December 6, 2014 launch; First science (Nereus landing) June 16, 2016; First 

sample return February 6, 2018, Second landing (FG3) May 30, 2020; Second sample return 

October 28, 2021 

Attitude Control 

System (ACS) 

Off-the-shelf (OTS) inertial measurement unit (IMU), Star-trackers, Wheels, hydrazine thrusters 

LIDAR assisted precision landing system (landing gear for contingency, up to 1m/s landing 

velocity) 

Solar pressure torque from off-set solar easily countered by canting electric thrusters <1° 

54 kg 

Launch Atlas 401 Launch, C3 = 38.07 km2

/sec2 New Frontiers (NF) Expendable Launch Vehicle (ELV), 

performance to C3 of 1528 kg (1375 kg after 10% launch margin). 

Science Science/collection arm with camera, Two, six-bay sample capsules 

Extensive in-situ science powered by SA 

Wide/narrow field imager, infrared (IR) spectrometer, laser altimeter, IR, gamma-ray, neutron 

spectrometer, Alpha Particle X-ray Spectrometer (APXS), LAMS, thermal conductivity, electrical 

dissipation, ground penetrating radar return samples in two capsules. 

29.3 kg 

payload, 

(61.4 kg 

empty 

capsules) 

Power Single Orion derived Ultra-Flex SA (built for high-g Orion loads), Li-ion batteries for eclipse stays 134 kg 

Propulsion +1 NEXT Ion thrusters (7 m SA), OTS Xe feed and storage system, hydrazine Reaction Control 

System (RCS), 500 kg Xe for NEXT 

Cold gas Xe ‘landing’ system to minimize surface contamination 

209 kg 

Structures and 

mechanisms 

Thrust tube and tubular space frame propellant tanks mount directly to thrust tube 98 kg 

Communications 0.7 m antenna, two axis gimbal hemispheric coverage, 3 to 10 kbps, three omni antennas, 31 kg 

Command & data 

handling (C&DH) 

Two RAD750 processors for fault tolerance, 48 Gbit data storage 37 kg 

Thermal Heat-pipe radiators for cooling electronic components 

Heaters for propellant systems, MLI for S/C 

47 kg 

2.0 STUDY BACKGROUND AND ASSUMPTIONS 

2.1 Introduction 

The focus of this COMPASS study was to design a S/C and mission which samples at least two near Earth 

asteroids and a returns multiple samples from each to the Earth. This study focused on using SEP to enable 

sample returns from the Asteroids Nereus and 1996 FG3. Additional science mapping and in-situ science 

was also sought as a science objective. The design parameters (ELV choice, launch mass, incl. cost) were 

all designed to fit within a NF Class of mission. The trajectory will utilize an Earth flyby to both return first 

sample capsule and boost the S/C to the second asteroid target. 

A number of Trades to be looked at during the course of the design study were: 

 Trade primary SEP systems 

− NEXT 

− BPT4000 

− HiVHAC 

 Trade level of in-situ science 

 Trade number of asteroids sampled (1 or 2) 

2.2 Assumptions 

Summary of study assumptions and requirements are shown in Table 2.1. 
NASA/TM—2009-215825 3 

Table 2.1—Study Assumptions 

Subsystem 

area

Assumptions and 

study requirements

Critical 

trades

Top-level Sampler to orbit and land on two separate Near Earth Asteroids (NEA) 

Return rock samples and soils 

Figures of Merit (FOMs): Returned sample mass, # of samples, variety, science 

data, mission success probability, cost well below NF 

All- SEP, all-chemical, 

chem/SEP split 

System OTS equipment where possible, Technology Readiness Level 6(TRL 6) cutoff 2010, 

2014 launch year, Single fault tolerant 

Mass growth per ANSI/AIAA R-020A-1999 (add growth to make system level 30%) 

Mission summary Integrated SEP system, Earth-asteroid-Earth-asteroid-Earth, visit/orbit/land/ sample 

two asteroids, sample one, Earth flyby to return first payload, then sample the 

second and Earth flyby return, 6.5 yr SEP thrusting, 7 yr round trip 

Atlas 401 launch, C3 = 38.07 km2

/sec2 

7 yr mission: December 6, 2014 launch; First science (Nereus landing) June 16, 

2016; First sample return February 6, 2018, Second landing (FG3) May 30, 2020; 

Second sample return October 28, 2021 

Return of first sample before second landing lowers risks 

GN&C Closed loop ‘docking’ system, wheels for stability, gimbaled EP, 250 m/s secondary, 

13500 m/s primary (SEP), option for extended missions after sample return 

OTS IMU, Star-trackers, Wheels, hydrazine thrusters 

LIDAR assisted precision landing system (landing gear for contingency, up to 1 m/s 

landing velocity) 

Solar pressure torque from off-set solar easily countered by canting electric 

thrusters <1° 

SEP or chemical trajectories, 

landing or docking or hovering, 

sample collection scheme, 

cold gas for proximity ops 

Launch Vehicle Atlas 401 C3 38.07 km2

/sec2, 1528 kg 

Adapter: 4 m LPF 

Launch loads: Axial 11± 1 g, Lateral .4 ± 1.6g 

Atlas V, trade adaptors 

Propulsion Primary: 1+1 NEXT (7 m SA), 2+1 4.5 kW BPT-4000, four off the shelf Xe tanks 

Secondary: blow-down hydrazine RCS system, 1 lbf thrusters 

Terminal landing: 500 kg Xe cold gas (reduce contamination) 

Trade: 1+1 7 kW ion, 2+2 

4.5 kW BPT-4000, serial PPUs 

or cross-strapped 

Power Single Orion derived Ultra-flex SA (built for high-g Orion loads), Li-ion batteries for 

eclipse stays. 6500 W power to propulsion system (with 400 W housekeeping) 

Batteries for Asteroid and Moons eclipse, Sampling landing (> 9 hr) 

Array type, dual gimbals, cell 

type, battery options, use of 

SA to allow long stay times on 

moons 

Avionics/ 

Communications 

Science run from central controller (and one spare 0.7 m antenna, two axis gimbal 

hemispheric coverage, 3 to 10 kbps, three omni antennas, two RAD750 processors 

for fault tolerance, 48 Gbit data storage 

Computer type, X band or Ka 

band 

Thermal & 

environment 

Body mounted radiator (main loads 350 Wth (PPUs), 100 W (transmitters)). Heatpipe radiators for cooling electronic components 

Heaters for propellant systems, MLI for S/C 

Tank heaters, 0.6 to 1.7 AU thermal environment 

Deep space radiation level at 1.7 AU 

Mechanisms Science arm/camera/sampler, two-axis 0.3 m antenna, thruster gimbals ± 12 °, 

docking legs, sample capsule (2.9 km/s entry velocity capability), parachute impact 

suppression 

Landing legs, sample capsule, 

sampler arm, foam only impact 

suppression, harpoons 

Structures Primary: Rectangular, 3- by 3-m, truss, Al-Li; Secondary: 4% of stage components; 

Thrust tube and tubular space frame propellant tanks mount directly to thrust tube 

Developing model, need 

launch loads 

Science Science/collection arm with camera. Two, six-bay Sample Capsules (for total of 12 

samples) 

Extensive in-situ science powered by SA 

Wide/narrow field imager, IR spectrometer, laser altimeter, IR, gamma-ray, neutron 

spectrometer, APXS, LAMS, thermal conductivity, electrical dissipation, ground 

penetrating radar 

2.3 Growth, Contingency and Margin Policy 

Mass Growth: The COMPASS team uses the ANSI/AIAA R-020A-1999, Recommended Practice for 

Mass Properties Control for Satellites, Missiles, and Launch Vehicles (ref. 1). Table 2.2 shows the Percent 

Mass Growth separated into a matrix specified by level of design maturity and specific subsystem. Mass 

Growth Allowance (MGA) is defined as the predicted change to the basic mass of an item based on an 

assessment of the design maturity and fabrication status of the item, and an estimate of the in-scope design 

changes that may still occur. 
NASA/TM—2009-215825 4 

The percent growth factors are applied to each subsystem, after which the total system growth of the design 

is calculated. An additional growth is carried at the system level in order to add up to a total system growth 

of 30% of the dry mass of the system. Note that growth in propellant is either carried in the propellant 

calculation itself or in the ΔV used to calculate the propellant required to fly a mission. 

Power Growth: The COMPASS team uses a 30% margin on the bottoms up power requirements in 

modeling the power system. See Sections 3.1.2 and 5.4 for the power system assumptions. 

Table 2.2—Percent Mass Growth Allowance 

Code 

Design Maturity 

(Basis for Mass 

Determination) 

Percent Mass Growth Allowance 

Electrical/Electronic 

Components 

Structure 

Thermal Control 

Propulsion 

Batteries 

Wire Harnesses 

Mechanisms 

Instrumentation 

0-5 kg 5-15 kg >15 kg

E Estimated 

(preliminary sketches) 30 20 15 18 18 18 20 50 18 50 

L 

Layout 

(or major modification of 

existing hardware) 

25 20 15 12 12 12 15 30 12 30 

P 

Pre-Release Drawings 

(or minor modification of 

existing hardware) 

20 15 10 8 8 8 10 25 8 25 

C Released Drawings 

(calculated values) 10 5 5 4 4 4 5 5 4 5 

X 

Existing Hardware 

(actual mass from another 

program) 

3 3 3 2 2 2 3 3 2 3 

A Actual Mass 

(measured flight hardware) 0 0 0 0 0 0 0 0 0 0 

CFE Customer Furnished 

Equipment 0 0 0 0 0 0 0 0 0 0 

2.4 Redundancy Assumptions 

 Single fault tolerant where possible in the design of the subsystems. 

 Exceptions 

− SA 

− Propellant tanks 

− Radiators (design can be modified at ~11 kg penalty) 

− Sampling arm 

2.5 Mission Description 

This mission returned samples from two asteroids to the surface of the Earth. The asteroids 4660 Nereus 

and 1996 FG3 were chosen as scientifically desirable asteroids for this mission. 

2.5.1 Mission Analysis Assumptions 

Earth asteroids named after the first of their type discovered in 1862 (Apollo). Their heliocentric orbital 

semi major axes are greater than that of Earth, and their perihelion distance is greater than 1 AU (i.e., the 
NASA/TM—2009-215825 5 

distance the Earth is from the sun). The orbits of the Apollo asteroids cross the orbits of both that of 

Asteroid and Earth. Because Nereus’ orbit frequently comes very close to Earth, it is very accessible from 

Earth as well as a potential threat to the Earth. Due to its small size (approximately 1 km diameter) and 

hence smaller mass, its ∆V for rendezvous is smaller than the ∆V for rendezvous with our Moon. Nereus 

has a roughly 15 hr rotation. The asteroid Nereus trajectory details, orbital elements and assumptions are 

shown in Figure 2.1. 

Figure 2.1—4660 Nereus Body Orbital Details 

Figure 2.2—1996 FG3 Body Orbital Details 

 
NASA/TM—2009-215825 6 

The second asteroid chosen, 1996 FG3, is another Apollo, near Earth asteroid. Initial readings show that 

this asteroid is the dominant part of a binary asteroid system. 1996 FG3’s period of rotation has been 

determined to be about 3.5 hr and the orbital period of its satellite has been determined at 16 hr. The 

average bulk density of the asteroid has been estimated at 1.4 g cm3 and that the surface has a rubble pile 

structure. 1996 FG3 trajectory details, orbital elements and assumptions are shown in Figure 2.2. 

2.5.2 Mission Analysis Analytic Methods 

The trajectory design for this mission was optimized using the Mission Analysis Low-thrust Trajectory 

Optimization (MALTO) tool. The baseline mission launches to a C3 of 38.07 km2

/s2 and performs a 

rendezvous with Nereus in June of 2016, stays at the target for two months for sample collection operations 

and then departs for Earth to drop the first sample return capsule (SRC) with a constrained entry velocity 

and position. After the sample is released, the S/C completes the Earth flyby and arrives at the second 

target, 1996 FG3, in May of 2020. Following the two months at 1996 FG3, the S/C then departs and targets 

Earth with constrained entry conditions with an arrival V∞ of 6.8 km/s. 

Mission analysis was performed in an iterative fashion. An initial trajectory to the target was performed 

using MALTO to get the electric propulsion system propellant loading for the missions. With this 

propellant, the bottoms-up estimation of the vehicle mass was completed by the team. Once this bottoms-up 

mass was calculated, the trajectory was rerun in order to provide performance for at least that calculated 

total wet mass. The mission was iterated until the amount of mass pushed by the EP system was greater 

than or equal to the total wet mass of the vehicle. 

2.5.3 Mission Analysis Event Timeline 

Figure 2.3—Trajectory Main Mission Details 
NASA/TM—2009-215825 7 

Figure 2.4—S/C Distance From the Sun and Earth Over Mission Time 

2.5.4 Mission Trajectory Details 

Figure 2.5—4660 Nereus Trajectory Plot 

2.6 Concept of Operations (CONOPS) 

Remote Sensing (30 days) 

 CHAllenging Minisatellite Payload (CHAMP)imaging, radio science to map gravity, LIDAR, 

neutron detector (find hydrogen) 

 High orbit spiral down to 

 Low orbit (5 km) 

 Asteroids mapped to sufficiency for landing near rock outcropping 

Landing (~100 min) 

 One burn descent from 0.5 to 1 km starting altitude 

 Autonomous landing using maps generated from remote sensing phase, LIDAR and CHAMP 

imager 

− Impact landing speeds of 25 cm/s 

 10° incline (max), nearby large boulder (10s to 100s of meters) 

 Seeking erosion of large boulder for sampling 
NASA/TM—2009-215825 8 

Landed Science (Minimum 3.5 hr, Desired 15 hr) 

 CHAMP imaging (panoramic and microscopy), APXS, Neutron detector 

 Sample acquisition arm/CHAMP imaging of sample area for context of sample 

− Images sent to Earth for review and sample selection 

− Sample acquisition program sent from Earth 

− Samples collected and stored (with confirmation from CHAMP) 

Samples Collected 

 Rocks eroded from ejecta blocks (from asteroid core) 

 Soil samples 

Figure 2.6—Mission Main Events Trajectory Graphic 

2.7 Launch Vehicle Details 

2.7.1 Launch Vehicle Trade-Space Relative Performance 

For this mission, several lower performing launch vehicles were looked at as options. Figure 2.7 shows the 

relative performance as a function of C3, of the Delta II, Delta II H and Falcon 9. 
NASA/TM—2009-215825 9 

C3 (km2/sec2) 

Figure 2.7—Performance Curves for Launch Vehicles of Interest 

2.7.2 Atlas 401 (4 m Fairing) Performance 

From the website astronautix.com, the Atlas V family of launch vehicles offers the performance in Table 

2.3. In order to clear up confusion, note that the Atlas 401 is the Atlas V launch vehicle with a 4 m diameter 

fairing. 

Table 2.3—Atlas V family Performance 

Configuration LEO 28° LEO 

polar

Geosynch 

transfer

Geosynch

Atlas V 401 12,500 10,750 5,000 N/A 

Atlas V 501 10,300 9,050 4,100 1,500 

Atlas V 511 12,050 10,200 4,900 1,750 

Atlas V 521 13,950 11,800 6,000 2,200 

Atlas V 531 17,250 14,600 6,900 3,000 

Atlas V 541 18,750 15,850 7,600 3,400 

Atlas V 551 20,050 17,000 8,200 3,750 

The Atlas V launch vehicle system is based on the 3.8-m (12.5-ft) diameter Common Core Booster (CCB) 

powered by a single RD-180 engine. A three-digit naming convention was developed for the Atlas V launch 

vehicle system to identify it’s multiple configuration possibilities, and is indicated as follows: the first digit 

identifies the diameter class (in meters) of the payload fairing (4 or 5 m); the second digit indicates the 

number of solid rocket motors used (zero for Atlas V 400 and zero to five for Atlas V 501); the third digit 

represents the number of Centaur engines. Figure 2.8 shows the performance of the Atlas V (521) versus C3

from the NASA Kennedy Space Center (KSC) launch performance website. Use of the Atlas 401 for NF 

missions allows for a raise of the cost cap by $40M.
NASA/TM—2009-215825 10 

NASA ELV Performance Curve(s) 

High Energy Orbits 

Please note the ground rules and assumptions below 

C3 (km2/sec2) 

Assumptions: 

Atlas V (521) 

This performance does not include the effects of orbital debris compliance, 

which must be evaluated on a mission-specific basis. This could result in a 

significant performance impact for mission in which launch vehicle hardware 

remains in Earth orbit. 

3-sigma mission required margin, plus additional reserves as determined by 

the LSP. 

Launch from SLC–41 at Cape Canaveral Air Force Station (CCAFS). 

Performance values assume harness, logo, reradiating antenna, three 

payload fairing doors.

Payload mass greater than 9000 kg (19,841 lb) may require mission 

unique accommodations. Type B2 payload adapter plus type C2 

spacer. 

5-m Short Payload Fairing 

185 km (100 n mi) minimum park orbit perigee altitude. 

185 km (100 n mi) minimum escape orbit perigee altitude. 

Performance shown is applicable to declinations between 28.5° and 

–28.5°. 

Figure 2.8—Atlas V 421 Performance Curve. 

2.7.3 Atlas 401 Payload Fairing Details 

Figure 2.9—Atlas 401 4 m LPF Payload Fairing ELV 
NASA/TM—2009-215825 11 

2.8 Launch Vehicle Packaging 

Using a side-launch configuration allows for the following packaging and concept design benefits. 

 Use of smaller, 4-m shroud 

 Fixed landing legs 

 Lower center-of-mass for more stable asteroid landings 

 Eases stowage of large SA 

 Long thrust tube attachment of propellant tanks 

Figure 2.10—NEA Sample Return S/C—Atlas 401 Fairing Packaging Close View 

Figure 2.11—NEA Sample Return S/C—Atlas 401 Fairing Envelop and Packaging View 

2.9 Sample Return Capsule (SRC) System Level Summary 

2.9.1 Sample Collection Requirements 

Science return of at least nine total viable samples and one spare. Each sample collection container has 

enough room to hold six samples, for a total of 12 samples returned (six from each asteroid). Given a total 
NASA/TM—2009-215825 12 

of 12 chances to get a sample worth studying, the following requirements were levied on the sample 

collection containers. 

 Desired sample traits—nine total (100 g each) + one spare 

− Density ~1.7 g/cm3

− Portion of ‘large blocks’ 

− ~4.5 cm diameter, ‘golf ball size’ 

 Mechanisms to collect the sample 

− Four degrees-of-freedom (DOF) collection arm (example in Figure 2.12) 

− 1 m reach 

− Scoop type bifurcated shovel 

− Motorized joints 

− Cable end effecter actuation 

− Sample capsule loading/sealing/separation 

− Swing type carousel 

Figure 2.12—Example Science Sampler Collection Arm 

2.10 Basic Science Payload Description 

Table 2.4 is the MEL of the baseline science payload used in all cases but case 1a in the trade space 

examined in this study. An additional few elements are carried in a super science package included in case 1 

and detailed in the trade studies Section 8.0. 

Table 2.4—Science Package Portion of the MEL 

WBS Description QTY

Unit 

Mass CBE Mass Growth Growth Total Mass

Number NEA Sampler (August 2008) - Case 1 (kg) (kg) (%) (kg) (kg)

06 NEA Sampler Spacecraft 1138.27 9.9% 113.05 1251.32

06.1 Science Payload 25.20 16.1% 4.06 29.26

06.1.1 Arm Mounted Science Instruments 1.40 21.4% 0.30 1.70

06.1.1.a Panoramic / microscopic color imager (JPLÕs CHAMP) 1 1.00 1.00 20.0% 0.20 1.20

06.1.1.b Alpha Particle X-ray Spectrometer (U. GuelphÕs APXS) 1 0.4 0.40 25.0% 0.10 0.50

06.1.1.c Misc #3 0 0.0 0.00 0.0% 0.00 0.00

06.1.1.d Misc #4 0 0.0 0.00 0.0% 0.00 0.00

06.1.1.e Misc #5 0 0.0 0.00 0.0% 0.00 0.00

06.1.1.f Misc #6 0 0.0 0.00 0.0% 0.00 0.00

06.1.2 Body Mounted Science Instruments 23.80 15.8% 3.76 27.56

06.1.2.a Approach/Hazard Avoidance/Landing Lidar (OptechÕs CA 1 20.00 20.00 15.0% 3.00 23.00

06.1.2.b Neutron Detector/Gamma Ray Spect. (IKIÕs HEND) 1 3.80 3.80 20.0% 0.76 4.56

06.1.2.c Misc #3 0 0.00 0.00 0.0% 0.00 0.00

06.1.2.d Misc #4 0 0.00 0.00 0.0% 0.00 0.00
NASA/TM—2009-215825 13 

Below in the following two sections are short, bulleted description of the science instruments used in the 

baseline payload. 

2.10.1 Arm Mounted (~1 m) Instruments 

The following instruments were mounted on the sample collection arm. 

 Panoramic/microscopic color imager (NASA Jet Propulsion Laboratory’s (JPLs) CHAMP, used on 

orbit/surface) 

− 1 kg, 30.1- by 13.1- by 9.5-cm, Powerpeak = 7 W, 3 μm/pixel, 0.4 mrad from orbit, 

120 Mb/day, 10 GB internal storage 

 APXS (U. Guelph, used on surface) 

− 0.2 kg sensor head (on arm) 7- by 5-cm diameter, 0.2 kg electronics (on S/C) (20- by 10- by 

1-cm), Powerpeak 2.5 W (30 V), 32 kB/s 

2.10.2 Body Mounted Instruments 

The following instruments are mounted on the body of the S/C. 

 Approach/hazard avoidance/landing LIDAR (Optech’s Canadian Asteroid Exploration LIDAR for 

Orbital Topometry–2 (CAMELOT–2) Canadian Space Agency (CSA) contribution) 

− 20 kg, 0.0225 m3

, Powerpeak = 140 W, Data Ratepeak = 500 kbps 

Lastly, a potential contribution to the list of science instruments is the neutron detector/gamma ray 

spectrometer below. Note that it does not appear in the science payload Some of the science instruments 

were book-kept in other subsystems due to their dual use. Specifically, the laser altimeter sensor listed in 

the Remote Sensing section of Table 2.5 lists the science instruments as chosen by the science team at the 

APL. Table 2.5 lists how those instruments were grouped by the science payload planners in a bottoms-up 

science instrument MEL. Color-coding of blue, yellow and aqua group the science elements into like 

sensing instruments categories. These instruments were then regrouped in the baseline science MEL in 

Table 2.4. 

Table 2.4 is book-kept in the Guidance, Navigation and Control subsystem MEL. The items in light blue are 

grouped together on the arm. Items in light green are grouped together in the spectrometers line item in the 

MEL. The shared DPU is book-kept in the C&DH system MEL Table 2.5 or the MEL in Table 2.4 

 Neutron Detector (ND)/Gamma Ray Spectrometer (IKI’s HEND) (possible contribution—used in 

orbit) 

− Mass = 3.8 kg, dimensions: 25 by 15 by 15 cm, Powerpeak =8 W, Data rate: 1 kB/frame 

2.11 Super Science Package Description 

Table 2.5 lists the science instruments as chosen by the science team at the APL. Table 2.5 lists how those 

instruments were grouped by the science payload planners in a bottoms-up science instrument MEL. Colorcoding of blue, yellow and aqua group the science elements into like sensing instruments categories. These 

instruments were then regrouped in the baseline science MEL in Table 2.4. 

Table 2.5—APL Science Payload 

Instrument Heritage/Analog Mass 

(kg)

Power 

(W)

Remote Sensing 

Wide/Narrow Field Imager MESSENGER (MDIS) 3.5 4.2 

IR Spectrometer MESSENGER (MASCS?) 3.1 6.7 

Laser altimeter NEAR 5 15 

Neutron Spectrometer MESSENGER (GRNS) 4 5 

Gamma-Ray Spectrometer MESSENGER (GRNS) 8 16 

Surface Sensing 

APXS MER 2 2 

LAMS 4 6 

Microscopic Imager 1 0.6 
NASA/TM—2009-215825 14 

Instrument Heritage/Analog Mass 

(kg)

Power 

(W)

Thermal Conductivity Rosetta et al. 0.5 0.5 

Electrical Dissipation 1.5 3 

Ground Penetrating Radar 8 30 

 Total: 40.6 kg 89.0 W 

Support 

System 

Robotic Arm 6 8 

Shared DPU MESSENGER 3.6 6 

Total: 50.2 kg 103.0 W

Some of the science instruments were book-kept in other subsystems due to their dual use. Specifically, the 

laser altimeter sensor listed in the Remote Sensing section of Table 2.5 lists the science instruments as 

chosen by the science team at the APL. Table 2.5 lists how those instruments were grouped by the science 

payload planners in a bottoms-up science instrument MEL. Color-coding of blue, yellow and aqua group 

the science elements into like sensing instruments categories. These instruments were then regrouped in the 

baseline science MEL in Table 2.4. 

Table 2.4 is book-kept in the Guidance, Navigation and Control subsystem MEL. The items in light blue are 

grouped together on the arm. Items in light green are grouped together in the spectrometers line item in the 

MEL. The shared DPU is book-kept in the C&DH system MEL Table 2.5 or the MEL in Table 2.4 

Table 2.5—NEARER Baseline Science Payload Details 

(Adapted From Applied Physics Laboratory (APL) Science Payload) 

Table 2.7 is the MEL used in the super science case 1a described in more detail in the trade space 

section of this report.

Table 2.6—Super Science Package Portion of the MEL 

NEARER Science Payload Heritge/Analog Mass (kg) Power (W)

Arm Mounted Science Instruments 5.4

Panoramic / microscopic color imager (JPL’s CHAMP) CHAMP 1

Alpha Particle X-ray Spectrometer (U. Guelph’s APXS) APXS 0.4

LAMS 4

Misc #4

Misc #5

Misc #6

Body Mounted Science Instruments 48.6

Approach/Hazard Avoidance/Landing Lidar (Optech’s CAMELOT-2) CAMELOT-2 20

Spectrometers: IR, Neutron, Gamma-Ray (fm Trojan Lander) MESSENGER 15.1

IR Spectrometer MESSENGER 3.1 6.7

Neutron Spectrometer MESSENGER 4 5

Gamma-Ray Spectrometer MESSENGER 8 16

Wide/Narrow Field Imager MESSENGER 3.5 4.2

Thermal Conductivity (Rosetta), Electrical Dissipation, Ground Penetrating Radar 10

Thermal Conductivity Rosetta et al. 0.5 0.5

Electrical Dissipation 1.5 3

Ground Penetrating Radar 8 30

Total 54

WBS Description QTY

Unit 

Mass CBE Mass Growth Growth Total Mass

Number NEA Sampler (Sept. 2008) - Case 1a (kg) (kg) (%) (kg) (kg)

06 NEA Sampler Spacecraft 1281.47 9.5% 122.06 1403.53

06.1 Science Payload 54.00 11.7% 6.32 60.32

06.1.1 Arm Mounted Science Instruments 5.40 5.6% 0.30 5.70

06.1.1.a Panoramic / microscopic color imager (JPLĶs CHAMP)1 1.00 1.00 20.0% 0.20 1.20

06.1.1.b Alpha Particle X-ray Spectrometer (U. GuelphĶs APXS) 1 0.4 0.40 25.0% 0.10 0.50

06.1.1.c LAMS 1 4.0 4.00 0.0% 0.00 4.00

06.1.1.d Misc #4 0 0.0 0.00 0.0% 0.00 0.00

06.1.1.e Misc #5 0 0.0 0.00 0.0% 0.00 0.00

06.1.1.f Misc #6 0 0.0 0.00 0.0% 0.00 0.00

06.1.2 Body Mounted Science Instruments 48.60 12.4% 6.02 54.62

06.1.2.a Approach/Hazard Avoidance/Landing Lidar (OptechĶs CAMELOT-2) 1 20.00 20.00 15.0% 3.00 23.00

06.1.2.b Spectrometers: IR, Neutron, Gamma-Ray (fm Trojan Lander) 1 15.10 15.10 20.0% 3.02 18.12

06.1.2.c Wide/Narrow Field Imager 1 3.50 3.50 0.0% 0.00 3.50

06.1.2.d Thermal Conductivity (Rosetta), Electrical Dissipation, Ground Penetrating Radar 1 10.00 10.00 0.0% 0.00 10.00
NASA/TM—2009-215825 15 

3.0 BASELINE DESIGN—CASE 1 

3.1 Top Level Design (MEL and PEL) 

3.1.1 Master Equipment List (MEL) 

Table 3.1 lists the MEL of the design for only the top level masses. The total growth on the dry mass of the 

S/C is then rolled up to find a total growth mass and growth percentage. The Growth column is where each 

subsystem lists the recommended growth factor on each line items following the AIAA WGA schedule 

outlined in Table 2.4 in Section 2.4. The MEL takes all of the items and racks them up into totals and 

calculates a total CBE mass, a Total mass and a total Growth Mass. 

Table 3.1—Master Equipment List—Case 1 

The MEL (Table 3.1) captures the bottoms up estimation of CBE and growth percentage line item by item 

for each subsystem. Table 3.3 wraps up those total masses, CBE and total mass after applied growth 

percentage. In order to meet the total of 30% at the system level, an allocation is necessary for system level 

growth. This additional system level mass is assumed as part of the inert mass that is flown along the 

required trajectory. Therefore, the additional system level growth mass impacts the total propellant loading 

for the mission design. 

Using the low thrust trajectory tool MALTO, and a starting guess of delivered target mass set to the bottoms 

up mass of the Case 1, the performance to the C3 of 38.07 km/s is calculated from a look up table to be 1580 

kg (see 2.5.2 for mission analysis assumptions, and 2.7.1 for Launch Vehicle assumptions). A 10% margin 

is taken off of this performance, leaving 1375 kg available launch performance. This mass of 1375 is what 

the SEP system flies as the wet mass of the S/C, and the xenon propellant loading is sized to push the total 

mass of 1375 kg. Launch vehicle margin is the mass that is not flown with the spacecraft. 

Performance and margin on the launch vehicle are calculated as follows and shown on Table 3.2. 

Available ELV performanc kg1375%)10*(e

Margin kg152%10*

ELV Perfomance to 3 kg1528

−= =

==

==

AA

A

AC

The total bottoms-up wet mass as shown in Table 3.1 of the system before the additional system mass is 

carried is 1251 kg (alsosee the row marked Estimated S/C Wet Mass and look in the column for Total Mass 

in Table 3.2). The total bottoms-up growth in this mass is 113 kg, or 16% of the total dry mass. In order to 

meet the 30% dry mass growth requirement at the system level, an additional 101 kg system mass is carried 

as shown in Table 3.2. This brings the total wet mass with growth to 1352 kg. An additional 23 kg is 

available in this bottom’s up modeling over the 10% launch vehicle mass and can be added to the 1352 wet 

WBS Description QTY

Unit 

Mass CBE Mass Growth Growth Total Mass

Number NEA Sampler (August 2008) - Case 1 (kg) (kg) (%) (kg) (kg)

06 NEA Sampler Spacecraft 1138.27 9.9% 113.05 1251.32

06.1 Science Payload 25.20 16.1% 4.06 29.26

06.2 Lander Spacecraft 1019.77 9.0% 92.07 1111.85

06.2.1 Attitude Determination and Control 44.28 21.1% 9.36 53.64

06.2.2 Command and Data Handling 29.30 27.2% 7.96 37.26

06.2.3 Communications and Tracking 24.20 27.1% 6.55 30.75

06.2.4 Electrical Power Subsystem 115.50 15.7% 18.19 133.69

06.2.5 Thermal Control (Non-Propellant) 40.11 18.0% 7.22 47.33

06.2.6 Propulsion 192.10 8.3% 16.00 208.09

06.2.7 Propellant 425.41 0.0% 0.00 425.41

06.2.8 Structures and Mechanisms 148.88 18.0% 26.80 175.68

06.3 Sample Return Craft 93.30 18.1% 16.91 110.21

06.3.1 Electrical Power Subsystem 6.00 20.0% 1.20 7.20

06.3.2 Thermal Control (Non-Propellant) 4.60 18.0% 0.83 5.43

06.3.3 Structures and Mechanisms 82.70 18.0% 14.89 97.59
NASA/TM—2009-215825 16 

mass to bring the system level growth up from 30%. Each SRC’s mass with 30% growth applied is 61 kg. 

Each SRC can return up to 1kg worth of sample material. 

Table 3.2—System Integration Summary—Case 1 

3.1.2 Power Equipment List (PEL) 

Power mission operations modes assumed: Peak, nominal, and standby. The power system data was 

provided by science, GN&C, avionics, communications, and thermal subsystems. The power required by 

the power subsystem (conversion boxes, line losses) was kept internally to the power design and is not 

shown in the PEL. Peak power was assumed for propulsion subsystem during SEP Thrusting and the 

asteroid science mapping mission events. The panoramic microscopic color imager (JPL’s CHAMP) and 

LIDAR were used during approach and landing phase. The APXS (U. Guelph’s) and Neutron 

Detector/Gamma Ray Spectrometer (IKI’s HEND) were used during Landed Science Phase. Reuse of the 

SEP solar power for science and operations was assumed. Battery size was based on asteroid eclipse times: 

~8 hr maximum. Growth of 30% was assumed for power needs (except for Electric Propulsion). 

The power required for nominal loads, based on mission modes of operation, is shown in Table 3.3. The 

waste heat used to size the thermal system is shown in the bottom of Table 3.3. 

Spacecraft Master Equipment List Rack-up (Mass)

COMPASS S/C 

Design 

WBS Main Subsystems CBE Mass (kg)

Growth 

(kg) Total Mass (kg)

Aggregate 

Growth (%)

01 Asteroids Sampler Spacecraft 1138.3 113.0 1251.3

06.1 Science Payload 25.2 4.1 29.3 16%

06.2 Asteroids Sampler Lander 1019.8 92.1 1111.8

06.2.1 Attitude Determination and Control 44.3 18.7 53.6 42%

06.2.2 Command and Data Handling 29.3 8.0 37.3 27%

06.2.3 Communications and Tracking 24.2 6.5 30.7 27%

06.2.4 Electric Power 115.5 18.2 133.7 16%

06.2.5 Thermal Control 40.1 7.2 47.3 18%

06.2.6 Propulsion 192.1 16.0 208.1 8%

06.2.7 Propellant 425.1

06.2.8 Structures and Mechanisms 148.9 26.8 175.7 18%

06.3 Sample Return Craft (total, empty) 93.3 16.9 110.2 18%

06.3.1 Electrical Power Subsystem 6.0 1.2 7.2 20%

06.3.2 Thermal Control (Non-Propellant) 4.6 0.8 5.4 18%

06.3.3 Structures and Mechanisms 82.7 14.9 97.6 18%

Estimated Spacecraft Dry Mass 713 113 826.2 16%

Estimated Spacecraft Wet Mass 1138 113 1251.3

System LeveL Growth Calculations Total Growth

Dry Mass Desired System Level Growth 713 214 927.1 30%

Additional Growth (carried at system level) 101 14%

Total Wet Mass with Growth 1138 214 1352.2

Available Launch Performance to C3 (kg) 1375.4

Launch margin available (kg) 23.2

Estimated Spacecraft Inert Mass (for traj.) 808 214 1021.5

Sample Return Craft Total Mass CBE Mass (kg) Growth (kg) Total Mass (kg)

Aggregate 

Growth (%)

Estimated Sample Return Craft Mass 93.3 16.9 110.2 18%

Total with System Level Growth 93 28 121.3 30%

Number of Sample Return Craft 2

Total Mass per Sample Return Craft (empty) 60.6 kg

Total Mass, Sample Returned 1 kg

Total Mass, Sample Return Capsule (Full) 61.6 kg
NASA/TM—2009-215825 17 

Table 3.3—Power Estimations and Waste Heat Over Mission Modes Propulsion, 

WAvionics, WComm., WThermal, WGN&C, WPower, WScience, WCBE Total, W30 % Margin, WTotal, W

Launch 0 24 0 33 31 0 0 88 26.37 114 

S/C separation 16 24 383 33 31 0 0 487 141.27 628 

S/C checkout 16 24 403 33 126 0 221 823 242.13 1065 

SEP Thrusting 6350 24 0 33 88 0 2 6497 44.16 6541 

SEP Coast 16 24 403 33 98 0 2 576 168.06 744 

Communications 16 24 383 33 98 0 2 556 162.06 718 

Nereus Targeting 6350 24 0 33 88 0 2 6497 44.16 6541 

Nereus Science Mapping 16 24 0 33 98 0 221 392 112.83 505 

Nereus Mapping Communications 16 24 383 33 98 0 221 775 227.73 1003 

Nereus Approach and Landing 0 24 383 33 98 0 331 869 260.73 1130 

Nereus Landed Science 16 24 383 33 98 0 221 775 227.73 1003 

Nereus Landed Communications 16 24 403 33 98 0 2 576 168.06 744 

Nereus Take-off 16 24 383 33 98 0 2 556 162.06 718 

Earth Sample Dropoff/Flyby 16 24 383 33 98 0 2 556 162.06 718 

1996FG3 Body Targeting, etc. 16 24 383 33 98 0 2 556 162.06 718 

Waste Heat 

Propulsion, W Avionics, W Comm., W Thermal, W GN&C, W Power, W Science, W CBE Total, W 30 % Margin, W Total, W 

Launch 0.0 1.2 0.0 1.7 1.6 0.0 0.0 4.4 1.3 5.7 

S/C separation 0.8 1.2 19.2 1.7 1.6 0.0 0.0 24.3 7.3 31.6 

S/C checkout 0.8 1.2 20.2 1.7 6.3 0.0 11.0 41.2 12.3 53.5 

SEP Thrusting 443.4 1.2 0.0 1.7 4.4 0.0 0.1 450.7 135.2 586.0

SEP Coast 0.8 1.2 20.2 1.7 4.9 0.0 0.1 28.8 8.6 37.5 

Communications 8.0 12.0 191.5 16.5 49.2 0.0 1.0 278.1 83.4 361.5

Nereus Targeting 443.4 1.2 0.0 1.7 4.4 0.0 0.1 450.7 135.2 586.0

Nereus Science Mapping 0.8 1.2 0.0 1.7 4.9 0.0 11.0 19.6 5.9 25.5 

Nereus Mapping Communications 0.8 1.2 19.2 1.7 4.9 0.0 11.0 38.8 11.6 50.4 

Nereus Approach and Landing 0.0 1.2 19.2 1.7 4.9 0.0 16.5 43.5 13.0 56.5 

Nereus Landed Science 0.8 1.2 19.2 1.7 4.9 0.0 11.0 38.8 11.6 50.4 

Nereus Landed Communications 0.8 1.2 20.2 1.7 4.9 0.0 0.1 28.8 8.6 37.5 

Nereus Take-off 0.8 1.2 19.2 1.7 4.9 0.0 0.1 27.8 8.3 36.2 

Earth Sample Dropoff/Flyby 0.8 1.2 19.2 1.7 4.9 0.0 0.1 27.8 8.3 36.2 

1996FG3 Body Targeting, etc. 0.8 1.2 19.2 1.7 4.9 0.0 0.1 27.8 8.3 36.2 

3.2 Baseline System Level Summary 

 Low center of mass configuration for landing stability 

− 500 kg of Xe in COPV tanks: Four cylindrical OTS COPVs 

− SA (> 5 kW Orion heritage) deployed after launch—used for landed power 

− 5 kW Hall or 7 kW Ion propulsion systems on side of S/C 

− Radiators on top deck 

 Minimize deployables/mechanisms (only science, power, and communications) 

− Science/collection arm (1m with 0.5 m telescoping extension) 

− Sample capsule loading/sealing/separation 

− Single axis gimbal SA 

− Two-axis communications antenna 
NASA/TM—2009-215825 18 

 Power 

− Single Orion derived Ultra-Flex SA (built for high-g Orion loads) 

− Li-ion batteries for eclipse stays 

 Propulsion 

− 1+1 NEXT Ion thrusters (7 m SA) , OTS Xe feed and storage system, OTS hydrazine landing 

system 

− 950 kg Xe BPT-4000 Hall or 500 kg for NEXT 

− Cold Gas Xe ‘landing’ system to minimize surface contamination and hold S/C fixed to surface 

during sample arm operations (as needed) 

 GN&C 

− OTS IMU, Star-trackers, wheels, hydrazine thrusters 

− LIDAR assisted precision landing system 

− Solar pressure torque from off-set solar easily countered by canting electric thrusters <1° 

 Avionics/Communications 

− One 0.7 m antenna, two axis gimbal hemispheric coverage, 3 to 10 kbps (Kilobits per second) , 

three omni antennas 

− Two RAD 750 processors for fault tolerance 

 Thermal 

− Heat-pipe radiators for cooling electronic components 

− Heaters for propellant systems 

− MLI for S/C 

3.3 Baseline Design Concept Drawing and Description 

In order to maintain stability while landed, it is desirable to keep the center of gravity (CG) of the landed 

configuration as close to the surface as possible while maximizing the diameter of the landed “footprint”. 

These goals were accomplished by putting the two cylindrical Xe tanks in horizontally with respect to the 

main bus shelf. Cylindrical tanks generally are not designed to handle large loads in the radial direction 

very well, thus it was decided to launch the lander on its side (relative to the landed configuration) as shown 

previously in Figure 2.10, to ensure the launch loads are incurred while the tanks are in their axial direction. 

This orientation allowed the use of a thrust tube type structural design to handle the high launch loads, 

encapsulate the spherical hydrazine tanks for the RCS system, and provided a structurally sound mounting 

point for the Xe tanks. 

Two small decks were placed next to the Xe tanks to allow mounting of all the internal science, power, 

guidance, avionics, and propulsion components. Those guidance and science instruments needed during 

landing or while on the surface were place on the bottom of the lower deck to allow a clear view of the 

surface, while all other internal components were mounted on the top of the decks. The space frame 

structure was used to mount all components external to the S/C including the thrusters, antennas, ultraflex 

array, radiators, and sample collection and return components, as well as provide good mounting points for 

the stowed array during launch. All components included in the design, except for the SA, are shown in 

Figure 3.1 and Figure 3.2. Note that the top of the lander is used to mount the SA and radiators, allowing 

the array to track the sun while keeping the radiators perpendicular to the sun. The dish and omni antennas 

are also mounted on the top surface. The bottom is dedicated to surface science and landing guidance. One 

side is dedicated to the star trackers, another side is dedicated to the thrusters, the third side is dedicated to 

sample collection and return, while the final side is dedicated to interfacing to the launch vehicle. The 

NEXT thrusters are mounted to a structure that allows for minimal gimbaling by orienting the thrust vector 
NASA/TM—2009-215825 19 

of each thruster through the CG of the vehicle. Figure 3.3 and Figure 3.4 show the deployed and stowed 

configurations of the lander while the overall dimensions of the lander are shown in Section 3.4. 

The vehicle design was based around a Single Ultraflex array. This design concept has the following benefits: 

 Saves mass 

 Keeps array away from asteroid/dust 

 Can be kept deployed to allow long term landings (hours-days) 

 Orion Ultraflex design capable of resisting large forces while deployed 

Figure 3.1—NEA Sample Return S/C—Top Components View 

Figure 3.2—NEA Sample Return S/C—Bottom Components View 
NASA/TM—2009-215825 20 

Figure 3.3—NEA Sample Return S/C—Deployed View 

Figure 3.4—NEA Sample Return S/C—Stowed View 
NASA/TM—2009-215825 21 

3.4 Baseline Design Concept Dimensions 

Figure 3.5—NEA Sample Return S/C—Footprint Dimensions 

Figure 3.6—NEA Sample Return S/C—Deployed Dimensions 
NASA/TM—2009-215825 22 

4.0 CHALLENGES, LESSONS LEARNED, AREAS FOR FUTURE STUDY 

4.1 Future Work 

 Better definition of major challenges 

− Precision landing 

− Securing and unsecuring S/C to surface to allow sampling 

− Hopping for more samples 

− Sample acquisition and storage 

− Asteroid environment (low gravity, dust) 

 Trades to reduce costs 

− Simplify science and collection strategy 

− Utlize more OTS systems 

4.2 Lessons Learned 

Electric Propulsion 

 Allows sample returns from two disparate near Earth asteroids 

 The large CEV derived SA provides sufficient power to allow for long-term landings for science 

collection 

 Reuse of Xe propellant for terminal landing and contingency ‘hopping’ avoids hydrazine 

contamination of surface samples 

 Bringing back samples should save costs of in-situ science instruments and operations 

5.0 SUBSYSTEM BREAKDOWN 

5.1 Communications 

This section describes the telecommunications subsystem of NEARER, dealing specifically with 

communications equipment on board the lander. Major telecommunications subsystem components have 

been chosen for NEARER (Near Earth Asteroids Rendezvous and sample Earth Returns) in response to the 

science mission requirements and design considerations such as anticipated maximum distances, desired 

data rates, on-board power and mass limitations. 

5.1.1 Communications Requirements 

The high level general requirements on the telecommunications subsystem are to provide the best signal 

possible in terms of available on-board electrical power, accuracy, reliability, and quality assurance, with 

constraints on mass, size and costs. Table 5.1 provides some of the important communications subsystem 

requirements. 
NASA/TM—2009-215825 23 

Table 5.1—Communications System Requirements 

Requirements Description 

Data rates 10 bps to 3 kbps for command (1 kbps typical) 

40 bps to 10 kbps for health and status telemetry 

10 kbps or higher for mission/science 

Daily data volume Daily data volume shall be at least 120 Mbits/day 

Data storage A minimum of 10 GB (8x1010 bits) internal storage shall be required 

Frequency Shall support Asteroid Moons Sampler at X-Band and use 8.44/7.75 GHz for uplink and downlink 

Housekeeping and 

overhead 

Housekeeping and any overhead shall include in stated data rates in this Table 

Available on-board power Shall consider different options such as using larger antennas/higher efficiency amplifiers and 

data requirements. 

Equivalent Isotropically 

Radiated Power (EIRP) 

EIRPs shall be as required in order to achieve 10 BER (Bit Error Rate) for given data rates 

The assumed requirement: 7 kbits/sec during 8 hr daily communications. The electronics were assumed to 

be single fault tolerant. 

5.1.2 Communications Assumptions 

Data transmitted back to Earth will go through the DSN 70-m dishes. The DSN consists of facilities in 

California’s Mojave Desert; near Madrid, Spain; and near Canberra, Australia. These stations are spaced 

about 120° apart on the globe—making sure any S/C can be observed constantly as Earth rotates. The data 

sent to the 70 m DSN dishes are transferred to some science ground station. 

The highest data rates will come during the mission phases when the Sampler return craft is on the surface 

of the moons. The sampler S/C will be on the surface of each of the asteroids for a maximum of 8 hr. 

The onboard processing and data buffering capabilities of the avionics system will handle data taken by the 

LIDAR during landing and departure to and from the asteroids. 

Orbital Downlink 

 Diameter 40 km (5×108 m2) 

 m/pixel = 1 

 Image overlap 1.5 

 Views/site 2 

 Effective bits/pix 8 

 Colors 5 

 Total orbital downlink 1011 bits 

− Assumes 60% is imaging data 

 Approximately 60 days of downlink 

Landed Downlink 

 Dominated by microscopic imager 

 Estimated total data ~ 8×109 bits 

 Some realtime data needed for validating surface science operations 

5.1.3 Communications Design and MEL 

The key components of the telecommunications subsystem include a 0.7 m high gain antenna (HGA) 

providing two-axis gimbaled hemispheric coverage, two omni-antennas, and two 85-W radio frequency 

(RF) Traveling Wave Tube Amplifiers (TWTA). The 0.7 m HGA is designed to support data rates from 3 to 

10 kbps and the Omni-antennas for emergency at data rates from 10 to 100 bps. The communications MEL 

is provided in Table 5.2. 
NASA/TM—2009-215825 24 

Table 5.2—Communications MEL for Baseline (Case 1) 

The communications system consisted of a 200 W X-Band Transponder, a single, dual axis 0.7 m HGA, 

and two Omni-antennas as shown in Figure 5.1. 

Figure 5.1—NEA Sample Return S/C—Communications Instruments 

5.1.4 Communications Trades 

None 

5.1.5 Communications Analytical Methods 

The following section contains the calculations for five different ways to quantify link budget analysis for 

this mission. The link budgets for NEARER provide values of RF transmit power of 85 and 200 W and 

antenna gains for X-Band. The first link budget calculation is the worst case with a distance between the 

Earth and Nereus of 2.4 AU. The next four are between Earth at apogee and Nereus at apogee assuming the 

WBS Description QTY

Unit 

Mass CBE Mass Growth Growth Total Mass

Number NEA Sampler (August 2008) - Case 1 (kg) (kg) (%) (kg) (kg)

06 NEA Sampler Spacecraft 1138.27 9.9% 113.05 1251.32

06.1 Science Payload 25.20 16.1% 4.06 29.26

06.2 Lander Spacecraft 1019.77 9.0% 92.07 1111.85

06.2.1 Attitude Determination and Control 44.28 21.1% 9.36 53.64

06.2.2 Command and Data Handling 29.30 27.2% 7.96 37.26

06.2.3 Communications and Tracking 24.20 27.1% 6.55 30.75

06.2.3.a X High Gain Antenna 23.40 27.5% 6.43 29.83

06.2.3.a.a Transmitter/Receiver 2 2.90 5.80 30.0% 1.74 7.54

06.2.3.a.b Power Amp 2 2.60 5.20 30.0% 1.56 6.76

06.2.3.a.c Switch Unit 1 4.40 4.40 15.0% 0.66 5.06

06.2.3.a.d Antenna 1 1.50 1.50 30.0% 0.45 1.95

06.2.3.a.e Band Pass Filter 0 0.00 0.00 0.0% 0.00 0.00

06.2.3.a.f Band Reject Filter 0 0.00 0.00 0.0% 0.00 0.00

06.2.3.a.g Sensor 0 0.00 0.00 0.0% 0.00 0.00

06.2.3.a.h Cabling 1 2.00 2.00 50.0% 1.00 3.00

06.2.3.a.i Diplexer 2 0.40 0.80 15.0% 0.12 0.92

06.2.3.a.j Coupler 1 0.40 0.40 15.0% 0.06 0.46

06.2.3.a.k Gimbal 1 2.30 2.30 30.0% 0.69 2.99

06.2.3.a.l Misc#2 1 1.00 1.00 15.0% 0.15 1.15

06.2.3.b Omni Antenna 0.60 3.0% 0.02 0.62

06.2.3.b.a Transponder 0 4.00 0.00 10.0% 0.00 0.00

06.2.3.b.b RF Assembly 0 0.20 0.00 3.0% 0.00 0.00

06.2.3.b.c Processing Module 0 0.00 0.00 0.0% 0.00 0.00

06.2.3.b.d Antenna 2 0.30 0.60 3.0% 0.02 0.62

06.2.3.b.e Misc#1 0 0.00 0.00 0.0% 0.00 0.00

06.2.3.b.f Misc#2 0 0.00 0.00 0.0% 0.00 0.00

06.2.3.b.g Misc#3 0 0.00 0.00 0.0% 0.00 0.00

06.2.3.b.h Misc#4 0 0.00 0.00 0.0% 0.00 0.00

06.2.3.b.i Misc#5 0 0.00 0.00 0.0% 0.00 0.00

06.2.3.c Communications Instrumentation 0.20 50.0% 0.10 0.30

06.2.3.c.a Coaxial Cable 2 0.10 0.20 50.0% 0.10 0.30
NASA/TM—2009-215825 25 

apogees are collinear with the center of the Sun. Table 5.3 summarizes the findings. Note: All analysis is 

done assuming 8 Gbits of data to be transmitted. 

Table 5.3—Link Budget Calculations for Communications Scenarios 

Case Power, 

W

Earth antenna size, 

m

Data rates, 

kbps

Transmission,

hr

Worst 200 70 5.9 376.7 

Case 1 200 70 212.0 10.5 

Case 2 200 34 48.0 46.3 

Case 3 85 70 90.0 24.7 

Case 4 85 34 21.0 105.9 

The number hours on Earth between the beginning of the transmission to the end of transmission will 

depend on the percentage of time the S/C are attached to the asteroid, can see Earth and the probability the 

Earth stations used to receive the transmission is available. The product of these numbers when divided into 

the hours of transmission will give one the total time necessary to transmit the data to Earth. 

The link budgets for NEARER provide values of RF transmit power at most 200 W and antenna gains for 

X-band. Worst case is for 2.4 AU Earth - Nereus distance (an overestimate - should be 2.0 AU), provides 

nearly the 7 kbps rate needed. Further trades of DSN time, antenna sizes (ground and S/C), and transmitter 

power need to be made. 

Figure 5.2—Worst Case Nereus Link Budget Analysis 
NASA/TM—2009-215825 26 

Figure 5.3—Case1 Nereus Link Budget Analysis 

Figure 5.4—Case 2 Nereus Link Budget Analysis 
NASA/TM—2009-215825 27 

Figure 5.5—Case 3 Nereus Link Budget Analysis 

Figure 5.6—Case 4 Nereus Link Budget Analysis 

5.1.6 Communications Risk Inputs 

None submitted 

5.1.7 Communications Recommendation 

See design 
NASA/TM—2009-215825 28 

5.2 Guidance, Navigation and Control (GN&C) 

5.2.1 GN&C Requirements 

The GN&C subsystem shall provide full 6-DOF control of the vehicle from launch through end of mission. 

This includes stabilization of the vehicle after launch vehicle separation, attitude control throughout the 

cruise, commanding and controlling all slews, and performing all automated landings, hops, and ascent 

maneuvers. 

5.2.2 GN&C Assumptions 

Much of the work in modeling the GN&C system for the Near Earth Asteroid Sample Return mission was 

built upon the system designed to perform this function in the Asteroid Sample Return mission completed 

by COMPASS. 

The ΔV values and Xe propellant allocations used are summarized in the table in Figure 5.7. 

Figure 5.7—Mapping Orbit Delta V Assumptions 

5.2.3 GN&C Design and MEL 

The GN&C subsystem hardware is made up of: 

 Four reaction wheels (Valley Forge VF MR 14.0, 14 Nms reaction wheel, 

http://www.vfct.com/aerospace/wheels/small-wheels) 

 Two Star Trackers (Adcole) 

 One internally redundant IMU (Northrop Grumman HRG) 

 Sun sensors to aide in Earth acquisition (Adcole Sun Sensors, New Horizons (NH) Heritage, two 

electronics boxes and three sensor heads, each) 

 GN&C software run on main C&DH computers 

 Utilizes LIDAR in science instrument subsystem for precision landing 

The detailed mass accounting for the GN&C subsystem can be seen in Table 5.4. The block diagram can be 

seen in Figure 5.8. 
NASA/TM—2009-215825 29 

Table 5.4—GN&C MEL (Case 1) 

Figure 5.8—GN&C Subsystem Block Diagram 

5.2.4 GN&C Trades 

No specific trades were completed. A feasible design was constructed and minor changes were made to help 

close the mission. 

5.2.5 GN&C Analytical Methods 

The GN&C subsystem mainly utilized the Rocket Equation to calculate propellant masses for the mapping, 

landing, hop, and ascent burns. Again, because of the short duration of the study, no in-depth dynamic 

analyses were completed. 

5.2.5.1 Solar Pressure Torques 

 Solar pressure torques are significant because of single array configuration 

 3.7 m offset between centers of mass and pressure (Figure 5.9) 

 8.5×10–4 Nm of torque at 1 AU 

 Assuming 15 Nm of momentum storage per axis, wheels would saturate approximately every 5 hr, 

with decreasing frequency as distance to sun increases 

− Despinning the wheels with the EP system with a 1° gimbal angle would take between 1.4 and 

10.4 hr depending on current thrust level 

 Preferably, off-pointing of the thruster by between 0.28° and 2.2° can counter the solar pressure 

torque with minimal thrust loss 

− Thrust loss of less than 0.1% 

WBS Description QTY

Unit 

Mass CBE Mass Growth Growth Total Mass

Number NEA Sampler (August 2008) - Case 1 (kg) (kg) (%) (kg) (kg)

06 NEA Sampler Spacecraft 1138.27 9.9% 113.05 1251.32

06.1 Science Payload 25.20 16.1% 4.06 29.26

06.2 Lander Spacecraft 1019.77 9.0% 92.07 1111.85

06.2.1 Attitude Determination and Control 44.28 21.1% 9.36 53.64

06.2.1.a Guidance, Navigation, & Control 44.28 21.1% 9.36 53.64

06.2.1.a.a Sun Sensors 6 1.00 6.00 20.0% 1.20 7.20

06.2.1.a.b Reaction Wheels 4 5.00 20.00 20.0% 4.00 24.00

06.2.1.a.c Star Trackers 2 3.19 6.38 20.0% 1.28 7.66

06.2.1.a.d IMU 1 6.90 6.90 20.0% 1.38 8.28

06.2.1.a.e Laser Altimeter (from Science Payload) 1 5.00 5.00 30.0% 1.50 6.50

06.2.1.a.f LIDAR 0 0.00 0.00 0.0% 0.00 0.00

06.2.1.a.g Misc#3 0 0.00 0.00 0.0% 0.00 0.00

Star

Trackers

C&DH Subsystem

Includes GN&C Software

IMU IMUIMUReaction

Wheels

Propulsion

Subsystem

Sun

Sensors

GN&C Subsystem

LIDAR

Laser

Altitmeter
NASA/TM—2009-215825 30 

Figure 5.9—Solar Pressure Torque Geometry 

Figure 5.10—Thrust as a Function of Gimbal Angle 

5.2.6 GN&C Risk Inputs 

None submitted 

5.3 Command and Data Handling (C&DH) 

5.3.1 C&DH Requirements 

 Storage: Be able to store all 60 days of an asteroid encounter: >12 Gbit 

 Transmission: 8 hr per day 

 ~7 kbits/sec needed 

5.3.2 C&DH Assumptions 

Assumed ~200 Mbits/day of data for all science instruments/housekeeping from combined science 

packages. 
NASA/TM—2009-215825 31 

5.3.3 C&DH Design and MEL 

Table 5.5—C&DH MEL for Baseline (Case 1) 

5.3.4 C&DH Trades 

None. 

5.3.5 C&DH Analytical Methods 

Design approach will be based on the New Horizons avionics systems, but using RAD750 processor instead 

of RAD6000. GN&C and C&DH processors are to be combined into one. Avionics will be similar to this, 

but greatly modified for RF communications and flight and command & telemetry computers. 

5.3.6 C&DH Risk Inputs 

None submitted 

5.3.7 C&DH Recommendation 

See Design 

5.4 Electrical Power System 

5.4.1 Power Requirements 

The power system shall provide sufficient power for the S/C system throughout the mission including Earth 

orbit, transfer to Nereus orbits, Nereus orbit and surface operations, transfer to 1996 FG3, 1996 FG3 orbit 

and surface operations, and sample return to Earth. In addition, the power system must be sized to supply 

sufficient power to the EP system throughout the mission. 

5.4.2 Power Assumptions 

Two SA sizes are included in different design iterations of the mission: one completely based on the Orion 

SA being developed under Constellation, and a second array with a similar design but larger diameter. An 

example of such an array is shown in Figure 5.11. The former is assumed to be OTS hardware without 

significant development costs. The larger diameter SA would incur some additional design and 

development costs but would provide more capability to the S/C. 

WBS Description QTY

Unit 

Mass CBE Mass Growth Growth Total Mass

Number NEA Sampler (August 2008) - Case 1 (kg) (kg) (%) (kg) (kg)

06 NEA Sampler Spacecraft 1138.27 9.9% 113.05 1251.32

06.1 Science Payload 25.20 16.1% 4.06 29.26

06.2 Lander Spacecraft 1019.77 9.0% 92.07 1111.85

06.2.1 Attitude Determination and Control 44.28 21.1% 9.36 53.64

06.2.2 Command and Data Handling 29.30 27.2% 7.96 37.26

06.2.2.a Command & Data Handling 21.00 18.9% 3.96 24.96

06.2.2.a.a Flight Computer 2 2.00 4.00 20.0% 0.80 4.80

06.2.2.a.b Command and Telemetry Computer 0 0.00 0.00 0.0% 0.00 0.00

06.2.2.a.c Data Interface Unit 2 1.00 2.00 30.0% 0.60 2.60

06.2.2.a.d Data Bus Operations Amplifier 0 0.00 0.00 0.0% 0.00 0.00

06.2.2.a.e Operations Recorder 2 1.10 2.20 30.0% 0.66 2.86

06.2.2.a.f Command and Control Harness (data) 1 4.00 4.00 30.0% 1.20 5.20

06.2.2.a.g Shared DPU (From APL Science Instruments) 0 0.00 0.00 0.0% 0.00 0.00

06.2.2.a.h Avionics enclosure 1 8.80 8.80 8.0% 0.70 9.50

06.2.2.a.i Misc #3 0 0.00 0.00 0.0% 0.00 0.00

06.2.2.b Instrumentation & Wiring 8.30 48.2% 4.00 12.30

06.2.2.b.a Operational Instrumentation, sensors 1 0.30 0.30 0.0% 0.00 0.30

06.2.2.b.b Data Cabling 1 8.00 8.00 50.0% 4.00 12.00

06.2.2.b.c Misc #1 0 0.00 0.00 0.0% 0.00 0.00

06.2.2.b.d Misc #2 0 0.00 0.00 0.0% 0.00 0.00
NASA/TM—2009-215825 32 

Figure 5.11—Orion-Derived UltraFlex SA Design 

The Orion SA is assumed to be about 5.5 m diameter and capable of producing 5 kWe of net power at 

30 Vdc to the S/C at beginning of life (BOL) at 1 AU distance from the sun. The net power includes losses 

from the array wiring and electronics, the gimbal assembly, and the main electronics unit. The mass of the 

SA wing is assumed to be 50 kg. 

The larger diameter SA net power and mass values are scaled based on a constant power and mass per area 

from the Orion array above. 

The dust environment on the Apollo asteroids is assumed to not have significant affect on the SA due the 

lack of an atmosphere. However, the array is expected to have some degraded performance after each 

landing. Consequently, the array is oversized. 

The batteries and electronics are also based on Orion technology. The battery chemistry is Li-ion type 

technology with a battery system-level specific energy of 120 W-hr/kg. An older example is shown in 

Figure 5.12. 

Figure 5.12—Example Li-Ion Battery, Built for the Mars Phoenix Lander 

(Similar Batteries are Planned for Orion) 

The power system electronics are based on the Main Bus Switching Unit (MBSU) approach of Orion where 

a single box contains electronics cards for SA regulation and control, battery charge control, and overall bus 

current handling. An overall charge efficiency of 80% was assumed to account for the losses in recharging 

the battery from the SA. 
NASA/TM—2009-215825 33 

Power Assumptions 

 5 kW BOL, 1 AU 

 2 kW BOL, 2.5 AU 

 320 Wnet during eclipse on Moons’ surface 

 500 Wnet in daylight on Moons’ surface 

Figure 5.13 shows the change of available power from the SAs available over mission time as the S/C 

travels farther from the sun, and then orbits with the NEA. 

Figure 5.13—Power Available From SA Over Mission Time 

5.4.3 Power Design and MEL 

The circular UltraFlex SA was chosen for two primary reasons. First, it provides OTS array technology that 

is lightweight and close to the required net power level for the mission. Second, it is more structurally 

sound when deployed, which may be necessary for the multiple landings on the moons’ surfaces. Leaving 

the array deployed during the landings seems to be less risky than retracting the array for each landing. 

Better understanding of the dust environment may affect this decision in the future. 

The solar distances of the overall mission made a larger SA attractive, so a second UltraFlex design at 7 m 

diameter was also included in another iteration. This size array should have similar characteristics to the 

5.5 m array with possibly some lower amount of allowable acceleration or higher structural requirement or 

both. 

The Orion-based array has not been finalized and there is a chance of the diameter changing. An increase in 

diameter would most likely be beneficial and welcome, whereas a decrease would most likely require a 

redesigned array for this mission due to the 5.5 m array seeming to provide what is the lowest possible 

power level acceptable for this mission. 

Since the projected timeline of this mission is similar to that of Orion, the solar cells were assumed to be in 

the same range of efficiency as Orion, 29% at the cell-level. If the launch date is later, beyond 2020, more 

advanced cells may be more realistic, with efficiencies around 30 to 32% and proportional power increases. 

Traditional rectangular folding SAs were also evaluated but did not make sense due to their lower power 

per mass and lower structural integrity when deployed. Their cost may be lower than the UltraFlex SA, but 

the other factors appear to be more important here. 
NASA/TM—2009-215825 34 

The SA power requirements are sized primarily from the EP needs. EP requires an array that is oversized 

for the remainder of the mission, with potential exception for the surface operations. Due to the dust 

environment and Sun-pointing challenges on the moons’ surfaces, excess array power for both ‘daytime’ 

power and recharging the batteries for ‘nighttime’ operations is probably worthwhile. 

The challenging sun-pointing requirements mentioned above are also why a two-axis gimbal is specified. 

This additional DOF may allow more potential landing sites that might have significant slopes or 

challenging lines-of-sight to the sun. 

The sample return S/C includes a small mass of batteries and power electronics as well. The same specific 

energy assumption was made here, though a primary battery may be a good choice since it might reduce the 

mass slightly, but no recharge capability is available. 

Figure 5.14—Power System Schematic 

Table 5.6—Electrical Power System MEL for Baseline (Case 1) 

WBS Description QTY

Unit 

Mass CBE Mass Growth Growth Total Mass

Number NEA Sampler (August 2008) - Case 1 (kg) (kg) (%) (kg) (kg)

06 NEA Sampler Spacecraft 1138.27 9.9% 113.05 1251.32

06.1 Science Payload 25.20 16.1% 4.06 29.26

06.2 Lander Spacecraft 1019.77 9.0% 92.07 1111.85

06.2.1 Attitude Determination and Control 44.28 21.1% 9.36 53.64

06.2.2 Command and Data Handling 29.30 27.2% 7.96 37.26

06.2.3 Communications and Tracking 24.20 27.1% 6.55 30.75

06.2.4 Electrical Power Subsystem 115.50 15.7% 18.19 133.69

06.2.4.a Solar Arrays 78.00 10.8% 8.44 86.44

06.2.4.a.a Solar Array Mass (cells and structure only) 1 70.00 70.00 10.0% 7.00 77.00

06.2.4.a.b Solar Array Gimbal Assembly 1 8.00 8.00 18.0% 1.44 9.44

06.2.4.a.c Misc#1 0 0.00 0.00 0.0% 0.00 0.00

06.2.4.a.d Misc#2 0 0.00 0.00 0.0% 0.00 0.00

06.2.4.b Power Management & Distribution 15.00 15.0% 2.25 17.25

06.2.4.b.a Main Bus Switching Unit 1 15.00 15.00 15.0% 2.25 17.25

06.2.4.b.b Battery Charge Control Unit 0 0.00 0.00 0.0% 0.00 0.00

06.2.4.b.c DC Switchgear/Shunt Regulator 0 0.00 0.00 0.0% 0.00 0.00

06.2.4.b.d Misc#2 0 0.00 0.00 0.0% 0.00 0.00

06.2.4.c Power Cable and Harness Subsystem (C and HS) 10.00 50.0% 5.00 15.00

06.2.4.c.a Spacecraft Bus Harness 1 10.00 10.00 50.0% 5.00 15.00

06.2.4.c.b PMAD Harness 0 0.00 0.00 0.0% 0.00 0.00

06.2.4.c.c Electric Propulsion Harnes 0 0.00 0.00 0.0% 0.00 0.00

06.2.4.c.d RPS to Spacecraft Harness 0 0.00 0.00 0.0% 0.00 0.00

06.2.4.c.e Power Cabling 0 0.00 0.00 0.0% 0.00 0.00

06.2.4.c.f Misc#1 0 0.00 0.00 0.0% 0.00 0.00

06.2.4.c.g Misc#2 0 0.00 0.00 0.0% 0.00 0.00

06.2.4.d Battery System 12.50 20.0% 2.50 15.00

06.2.4.d.a Battery Assembly-Primary 1 12.50 12.50 20.0% 2.50 15.00

06.2.4.d.b Secondary Battery Subsystem 0 0.00 0.00 0.0% 0.00 0.00
NASA/TM—2009-215825 35 

5.4.4 Power Trades 

None performed. 

5.4.5 Power Analytical Methods 

The power system analysis is based on current approaches and assumptions from the Orion S/C system. 

5.4.6 Power Risk Inputs 

None submitted 

5.4.7 Power Recommendation 

The recommended power system based on current understanding is a single UltraFlex SA, sized to meet the 

overall mission requirements; and Li-ion batteries for both the main S/C and the sample return S/C. 

5.5 Structures and Mechanisms 

5.5.1 Structures and Mechanisms Requirements 

The intent is to provide the necessary hardware for the science research instrumentation, avionics, 

communications, propulsion and power. Structure must be able to withstand applied loads from launch 

vehicle and provide minimum deflections, sufficient stiffness, and vibration damping to perform the 

mission and survive the round trip trajectory. The design of the structure will strive to minimize weight 

(mass) in order to optimize performance of the S/C and fit on the launch vehicle. Physically, the structure of 

the S/C must allow is to fit within confines of launch vehicle and its payload fairing. The structure must be 

stiff and strong enough to accommodate landing and takeoff from low g terrestrial bodies. 

5.5.2 Structures and Mechanisms Assumptions 

The structural design used the following baseline assumptions in the design of the main bus and SRC. 

 Material: Aluminum 

 Space frame with tubular members 

 Composite sandwich structure shelf 

 Welded and threaded fastener assembly 

5.5.3 Structures and Mechanisms Design and MEL 

Design Description 

 Thrust tube and tubular space frame in square configuration 

 Shelf of composite sandwich architecture with honeycomb core to mount hardware 

 Thin sheets to enclose structure and provide shear stiffness 

 Spring struts to support landing hardware 

5.5.4 Main S/C Bus Design 

The main S/C bus is modeled as a tubular space frame in polygonal configuration shown in Figure 5.15. 

The material is a shelf of composite sandwich architecture with honeycomb core to mount hardware. Thin 

sheets were used to enclose structure. Struts were added to support landing hardware. 

Table 5.7 shows both the lander S/C and the sample return craft as they were reported in the master S/C 

MEL. Note that there are two SRCs in this MEL, with quantity = 2. The installation calculation is done 

using 4% of the dry CVE mass of each subsystem. 
NASA/TM—2009-215825 36 

Figure 5.15—Main S/C Bus Analytical Design 

Table 5.7—Structures and Mechanical Systems MEL for Baseline (Case 1) 

WBS Description QTY

Unit 

Mass CBE Mass Growth Growth Total Mass

Number NEA Sampler (August 2008) - Case 1 (kg) (kg) (%) (kg) (kg)

06 NEA Sampler Spacecraft 1138.27 9.9% 113.05 1251.32

06.1 Science Payload 25.20 16.1% 4.06 29.26

06.2 Lander Spacecraft 1019.77 9.0% 92.07 1111.85

06.2.1 Attitude Determination and Control 44.28 21.1% 9.36 53.64

06.2.2 Command and Data Handling 29.30 27.2% 7.96 37.26

06.2.3 Communications and Tracking 24.20 27.1% 6.55 30.75

06.2.4 Electrical Power Subsystem 115.50 15.7% 18.19 133.69

06.2.5 Thermal Control (Non-Propellant) 40.11 18.0% 7.22 47.33

06.2.6 Propulsion 192.10 8.3% 16.00 208.09

06.2.7 Propellant 425.41 0.0% 0.00 425.41

06.2.8 Structures and Mechanisms 148.88 18.0% 26.80 175.68

06.2.8.a Structures 121.52 18.0% 21.87 143.40

06.2.8.a.a Primary Structures 102.07 18.0% 18.37 120.45

06.2.8.a.a.a Main Bus Structure 1 102.07 102.07 18.0% 18.37 120.45

06.2.8.a.a.b Misc#1 0 0.00 0.00 0.0% 0.00 0.00

06.2.8.a.a.c Misc#2 0 0.00 0.00 0.0% 0.00 0.00

06.2.8.a.b Secondary Structures 19.45 18.0% 3.50 22.95

06.2.8.a.b.a Balance Mass 1 0.00 0.00 18.0% 0.00 0.00

06.2.8.a.b.b Tank Supports and Bracketry 1 9.09 9.09 18.0% 1.64 10.73

06.2.8.a.b.c Landing Gear Structure 1 8.15 8.15 18.0% 1.47 9.62

06.2.8.a.b.d Solar Array Boom 1 0.61 0.61 18.0% 0.11 0.72

06.2.8.a.b.e Misc#3 1 0.70 0.70 18.0% 0.13 0.83

06.2.8.a.b.f Misc#4 1 0.89 0.89 18.0% 0.16 1.05

06.2.8.b Mechanisms 27.35 18.0% 4.92 32.28

06.2.8.b.a Solar Array Mechanisms 0.00 0.0% 0.00 0.00

06.2.8.b.b Thruster Mechanisms 0.00 0.0% 0.00 0.00

06.2.8.b.c Communications Mechanisms 0.00 0.0% 0.00 0.00

06.2.8.b.d Thermal Mechanisms 0.00 0.0% 0.00 0.00

06.2.8.b.e Adaptors and Separation 0.00 0.0% 0.00 0.00

06.2.8.b.f Additional Mechanisms 9.84 18.0% 1.77 11.61

06.2.8.b.f.a Sample Arm & Mechanism 1 7.50 7.50 18.0% 1.35 8.85

06.2.8.b.f.b Landing Gear Displacement Mech. 1 2.34 2.34 18.0% 0.42 2.76

06.2.8.b.g Installations 17.51 18.0% 3.15 20.67

06.2.8.b.g.a Science Payload Installation 1 1.17 1.17 18.0% 0.21 1.38

06.2.8.b.g.b C&DH Installation 1 1.17 1.17 18.0% 0.21 1.38

06.2.8.b.g.c Communications and Tracking Installation 1 0.97 0.97 18.0% 0.17 1.14

06.2.8.b.g.d GN&C Installation 1 0.37 0.37 18.0% 0.07 0.44

06.2.8.b.g.e Electrical Power Installation 1 4.62 4.62 18.0% 0.83 5.45

06.2.8.b.g.f Therrmal Control Installation 1 1.60 1.60 18.0% 0.29 1.89

06.2.8.b.g.g Electric Propulsion Installation 1 7.61 7.61 18.0% 1.37 8.97

06.2.8.b.g.h Misc #1 0 0.00 0.00 0.0% 0.00 0.00
NASA/TM—2009-215825 37 

SRC Design 

The SRC design (see Figure 5.16) will be constant across the three design cases. The design is loosely 

based upon the sample return capsule. Each case will use the same SRC to contain the samples from the 

asteroids for return to Earth. It is designed to withstand the launch and Earth return inertial loading as well 

as the extreme aerothermal re-entry and deceleration loads. The SRC has an overall diameter of 0.7 m and 

is 0.4 m tall with a 45° cone heat shield. The heat shield translates to expose a sample containment carousel 

structure. The heat shield joint is closed by engaging three latching mechanisms equally spaced around the 

joint circumference. The heat shield is constructed of Phenolic Impregnated Carbon Ablator (PICA) to 

withstand temperatures of up to 3600 °C. PICA is a modern Thermal protection systems (TPS) material and 

has the advantages of low density (much lighter than carbon phenolic) coupled with efficient ablative 

capability at high heat flux. 

The SRC is spin stabilized at 14 rpm for control of orientation during the aerodynamic deceleration with an 

estimated re-entry velocity approaching 2.9 km/s. The deceleration chutes are activated by a 3g 

accelerometer switch for the drogue and main chute deployment with timer delays. A battery powered UHF 

beacon transponder is also carried to assist in the recovery operation. 

Figure 5.16—SRC concept design 

The SRC is designed for the return of a total of 1 kg of samples in two separate six-sample carousel (for a 

total of 12 between the two carousels), with each container 5 cm deep with a 5 cm diameter. The entire 

sample carousel is 10 cm tall and 0.25 m in diameter. The temperature design limits are –4 to 40 °C for 

samples with a landing recovery shock mitigated to limit it to 4 g’s of impact by incorporating crushable 

honeycomb foam beneath the heat shield.

The sample retrieval mechanism is a four DOF collection arm with a 1 m reach and a 0.5 m telescoping 

extension. It incorporates a scoop type bi-furcated shovel and uses motorized joints. Other mechanisms 

include three latching and sealing devices and the SRC spin and separate mechanism with an electrical cablesevering guillotine. A Xenon cold gas system is pulsed to stabilized the lander during retrieval arm operations. 

5.5.5 SRC Design 

 Overall diameter 1 m, 45° cone angle 

 Main SRC 0.4 m tall 

 Carousel rotates horizontally for access to the surface and to the science instruments. 

 Sealed side door 

 Re-entry velocity capability—2.9 km/s entry (below the required 11.5 km/s for this mission) 

5 cm

0.25 m

10 cm

0.25 m

0.7 m

3 shell 

latching . 

clamps

ball screw drive

translates heat shield

guide pins

Parachute canister
NASA/TM—2009-215825 38 

− PICA heat shield (3600 °C) 

− Deceleration chutes 

− Beacon for tracking and pickup 

− Battery power 

 12 sample carousels (5 cm diameter by 5 cm deep) 

 Sample carousel 0.4 m diameter., 8 cm tall 

 Temperature limits of –4 to 40 °C for samples 

 Sample return shock mitigation—limit to 4 g’s 

− Crushable honeycomb 

− Foam 

Figure 5.17—Sample Collection Arm 

5.5.6 SRC MEL 

Table 5.8 racks up the masses and subsystems in the SRC located on the NEARER S/C. Note that in this 

baseline case (Case 1) in this study, there are two SRCs. In order to enter two carousels, the MEL line 

elements have quantity (QTY) of 2 in those places where appropriate. For example, there are 2 battery 

subsystems in power, and 2 sample canisters in structures and mechanisms, etc. 

Table 5.8—SRC MEL (Case 1) 

WBS Description QTY

Unit 

Mass CBE Mass Growth Growth Total Mass

Number NEA Sampler (August 2008) - Case 1 (kg) (kg) (%) (kg) (kg)

06 NEA Sampler Spacecraft 1138.27 9.9% 113.05 1251.32

06.1 Science Payload 25.20 16.1% 4.06 29.26

06.2 Lander Spacecraft 1019.77 9.0% 92.07 1111.85

06.2.1 Attitude Determination and Control 44.28 21.1% 9.36 53.64

06.2.2 Command and Data Handling 29.30 27.2% 7.96 37.26

06.2.3 Communications and Tracking 24.20 27.1% 6.55 30.75

06.2.4 Electrical Power Subsystem 115.50 15.7% 18.19 133.69

06.2.5 Thermal Control (Non-Propellant) 40.11 18.0% 7.22 47.33

06.2.6 Propulsion 192.10 8.3% 16.00 208.09

06.2.7 Propellant 425.41 0.0% 0.00 425.41

06.2.8 Structures and Mechanisms 148.88 18.0% 26.80 175.68

06.3 Sample Return Craft 93.30 18.1% 16.91 110.21

06.3.1 Electrical Power Subsystem 6.00 20.0% 1.20 7.20

06.3.1.a Battery Subsystem 2 3 6.00 20.0% 1.20 7.20

06.3.1.b Misc #2 0 0 0.00 0.0% 0.00 0.00

06.3.1.c Misc #3 0 0 0.00 0.0% 0.00 0.00

06.3.1.d Misc #4 0 0 0.00 0.0% 0.00 0.00

06.3.2 Thermal Control (Non-Propellant) 4.60 18.0% 0.83 5.43

06.3.3 Structures and Mechanisms 82.70 18.0% 14.89 97.59

06.3.3.a Sample Canisters w/Carousel 2 3.35 6.70 18.00% 1.21 7.91

06.3.3.b Sample Arm & Mechanism 0 7.50 0.00 18.00% 0.00 0.00

06.3.3.c Primary SRC Structure 2 12.80 25.60 18.00% 4.61 30.21

06.3.3.d S/C SRC Separation Mechanism 2 3.90 7.80 18.00% 1.40 9.20

06.3.3.e Re-Entry Aero Heat Shield 2 13.30 26.60 18.00% 4.79 31.39

06.3.3.f Recovery Parachutes 2 3.50 7.00 18.00% 1.26 8.26

06.3.3.g Shock Absorbing Material 2 4.50 9.00 18.00% 1.62 10.62
NASA/TM—2009-215825 39 

5.5.7 Structures and Mechanisms Trades 

Analysis and Trades 

 Sizing of space frame to accommodate requirements for 

 Antenna 

 Instrumentation 

 Landing gear 

 Fit within confines of launch vehicle 

5.5.8 Structures and Mechanisms Analytical Methods 

Sizing of the space frame to accommodate requirements for the antenna and instrumentation while fitting 

within confines of launch vehicle. 

Figure 5.18—Example Graphic of Main Structure 

Preliminary Structural Analysis 

 Provided conditions 

− Maximum axial acceleration from launch vehicle: 6g 

− Maximum lateral acceleration from launch vehicle: 2g 

− Approach velocity: 10 in./s (0.25 m/s) 

− Maximum allowable pressure on moon surface: 0.5 psi (3.4 kPa) 

 Thrust tube wall thickness of 0.125 in. (3.2 mm) with a 20 in. (508 mm) OD 

− Max. bending stress of 52 ksi (357 MPa) 

− Assumed lateral load through CG of lander 

 Landing gear displacement 

− Four 8 in. (203 mm) diameter pads 

− ~2200 lb (1000 kg) 

− Constant stiffness springs 

− Needed displacement: 5.5 in. (140 mm) 

5.5.9 Structures and Mechanisms Risk Inputs 

Potential impact with foreign object or due to nearby operations. 
NASA/TM—2009-215825 40 

5.5.10 Structures and Mechanisms Recommendation 

Finite Element Analysis (FEA) to determine stresses and displacements along with a modal analysis for 

vibrations. 

5.6 Propulsion and Propellant Management 

5.6.1 Propulsion and Propellant Management Requirements 

The S/C propulsion subsystem was required for three propulsion operations: 

1. Electric propulsion: Orbit transfer and insertion 

2. Chemical propulsion: Reaction/attitude control 

3. Chemical propulsion: Surface landing 

Three cases were examined to quantify benefits of alternate electric propulsion technologies. 

5.6.2 Propulsion and Propellant Management Assumptions 

Because an objective in this study was to determine mission benefits of electric propulsion, the electric 

propulsion subsystems used were either commercially available or systems currently under advanced 

development at NASA GRC. The development status of the technology has been accounted for in the Cost 

Analysis. Electric thruster performance used for mission analysis is based on demonstrated operation. 

All chemical thrusters used in the design as well as the propellant management components and propellant 

tanks for the electric and chemical propulsion propellants were commercially available devices from 

operating manufactures. The current technology parameters for the various EP thrusters are shown the 

Table 5.9. 

Table 5.9—NEXT Thruster Options Technology Assumptions 
NASA/TM—2009-215825 41 

5.6.3 Propulsion and Propellant Management Design and MEL 

Main Electric Propulsion Subsystem 

 Two NASA/GRC NEXT ion thrusters—1 operating, 1 spare 

 Gimbals on each thruster for thrust vector control 

 Two Power Processing Units individually mated to the thrusters (no cross-strapping) 

 Two COPV Ti-lined high-pressure cylindrical storage tanks for the Xe propellant (nominal) 

 Xe distribution system incorporates VACCO (Vacuum and Air Components Company of 

America—www.vacco.com)—developed pressure and flow control devices 

Figure 5.19—Main Propulsion System 

Vehicle Landing Propulsion System 

 Further revision is required to validate current mass estimates and propellant requirements 

 System based on an existing Moog cold-gas thrusters was assembled 

− Used Xe gas from main EP storage tanks 

Figure 5.20—Notional Vehicle Landing Propulsion System 

Reaction Control Propulsion System 

The propulsion subsystem is comprised of 

 16 - 1 lbf mono-prop thrusters placed around S/C body for reaction control 

− Aerojet MR-111 thrusters operating on hydrazine 

• Isp = 229 sec 
NASA/TM—2009-215825 42 

• Thrust = 4.4 N 

− Thrusters require power for operation of catalytic bed 

 Fuel stored in Ti metallic tank 

− Two spherical tank 

− Blow down pressurization with gHe 

− Propellant distribution system used design similar to systems developed for the Constellation 

program 

• Including fault tolerance configuration 

− Multiple tank and line heaters are included in mass model to prevent propellant and pressurant 

from freezing 

• Additionally, insulation included for same elements 

− Instrumentation - nominal suite of temperature and pressure sensors 

Figure 5.21—RCS Propulsion system schematic 
NASA/TM—2009-215825 43 

Table 5.10—Propulsion and Propellant System MEL for Baseline (Case 1) 

WBS Description QTY

Unit 

Mass CBE Mass Growth Growth Total Mass

Number NEA Sampler (August 2008) - Case 1 (kg) (kg) (%) (kg) (kg)

06 NEA Sampler Spacecraft 1138.27 9.9% 113.05 1251.32

06.1 Science Payload 25.20 16.1% 4.06 29.26

06.2 Lander Spacecraft 1019.77 9.0% 92.07 1111.85

06.2.1 Attitude Determination and Control 44.28 21.1% 9.36 53.64

06.2.2 Command and Data Handling 29.30 27.2% 7.96 37.26

06.2.3 Communications and Tracking 24.20 27.1% 6.55 30.75

06.2.4 Electrical Power Subsystem 115.50 15.7% 18.19 133.69

06.2.5 Thermal Control (Non-Propellant) 40.11 18.0% 7.22 47.33

06.2.6 Propulsion 192.10 8.3% 16.00 208.09

06.2.6.a Propulsion Hardware (EP) 38.20 8.0% 3.06 41.26

06.2.6.a.a Primary EP Thrusters 2 13.10 26.20 8.0% 2.10 28.30

06.2.6.a.b EPS Power Processing and Control 0 0.00 0.00 0.0% 0.00 0.00

06.2.6.a.c EPS Structure 12.00 8.0% 0.96 12.96

06.2.6.a.c.a EP Thruster Pod 0 0.00 0.00 0.0% 0.00 0.00

06.2.6.a.c.b EP Thruster Boom 0 0.00 0.00 0.0% 0.00 0.00

06.2.6.a.c.c Gimbal 2 6.00 12.00 8.0% 0.96 12.96

06.2.6.a.d EPS Thermal Control Subsystem 0.00 0.0% 0.00 0.00

06.2.6.a.d.a EPS Multi-Layer Insulation 0 0.00 0.00 0.0% 0.00 0.00

06.2.6.a.d.b EPS Heaters and Sensors 0 0.00 0.00 0.0% 0.00 0.00

06.2.6.a.d.c Misc #1 0 0.00 0.00 0.0% 0.00 0.00

06.2.6.b Propellant Management (EP) 45.03 11.7% 5.28 50.31

06.2.6.b.a Xe propellant tank(s) 2 12.72 25.44 2.0% 0.51 25.95

06.2.6.b.b High Pressure Feed System 1 15.68 15.68 18.0% 2.82 18.50

06.2.6.b.c Low Pressure Feed System 0 0.00 0.00 0.0% 0.00 0.00

06.2.6.b.d Residual Xe Propellant (non deterministic) 0 0.00 0.00 0.0% 0.00 0.00

06.2.6.b.e Temperature sensors 1 3.90 3.90 50.0% 1.95 5.85

06.2.6.b.f Propulsion Tank heaters 0 0.00 0.00 0.0% 0.00 0.00

06.2.6.b.g Propulsion Line heaters 0 0.00 0.00 0.0% 0.00 0.00

06.2.6.b.h Misc#1 0 0.00 0.00 0.0% 0.00 0.00

06.2.6.b.i Misc#2 0 0.00 0.00 0.0% 0.00 0.00

06.2.6.b.j Misc#3 0 0.00 0.00 0.0% 0.00 0.00

06.2.6.b.k Misc#4 0 0.00 0.00 0.0% 0.00 0.00

06.2.6.c Power Processing Unit (PPU) 69.00 8.0% 5.52 74.52

06.2.6.c.a PPU Mass 2 34.50 69.00 8.0% 5.52 74.52

06.2.6.c.b Cabling 0 0.00 0.00 0.0% 0.00 0.00

06.2.6.c.c Misc#1 0 0.00 0.00 0.0% 0.00 0.00

06.2.6.c.d Misc#2 0 0.00 0.00 0.0% 0.00 0.00

06.2.6.c.e Misc#3 0 0.00 0.00 0.0% 0.00 0.00

06.2.6.d Propulsion Hardware (Chemical) 0.52 8.0% 0.04 0.56

06.2.6.d.a Main Engine 0.52 8.0% 0.04 0.56

06.2.6.d.a.a Main Engine 4 0.13 0.52 8.0% 0.04 0.56

06.2.6.d.a.b Main Engine Gimbal 0 0.00 0.00 0.0% 0.00 0.00

06.2.6.d.a.c Misc#1 0 0.00 0.00 0.0% 0.00 0.00

06.2.6.e Propellant Management (Chemical) 0.00 0.0% 0.00 0.00

06.2.6.f Reaction Control System Hardware 7.37 2.0% 0.15 7.51

06.2.6.f.a RCS Thruster Subassembly 4 1.84 7.37 2.0% 0.15 7.51

06.2.6.f.b Misc#1 0 0.00 0.00 0.0% 0.00 0.00

06.2.6.f.c Misc#2 0 0.00 0.00 0.0% 0.00 0.00

06.2.6.f.d Misc#3 0 0.00 0.00 0.0% 0.00 0.00

06.2.6.g RCS Propellant Management 31.99 6.1% 1.95 33.93

06.2.6.g.a Fuel Tanks 2 7.81 15.62 2.0% 0.31 15.93

06.2.6.g.b Fuel Lines 0 0.00 0.00 0.0% 0.00 0.00

06.2.6.g.c Pressurization System - tanks, panels, lines 1 2.04 2.04 10.0% 0.20 2.24

06.2.6.g.d Feed System - regulators, valves, etc 1 14.33 14.33 10.0% 1.43 15.76

06.2.7 Propellant 425.41 0.0% 0.00 425.41

06.2.7.a Propellant (EP) 359.53 0.0% 0.00 359.53

06.2.7.a.a Primary EP Propellant Used 1 331.06 331.06 0.0% 0.00 331.06

06.2.7.a.b Primary EP Propellant Residulals (Unused) 1 28.47 28.47 0.0% 0.00 28.47

06.2.7.a.c Primary EP Propellant Performance Margin (Unused) 0 0.00 0.00 0.0% 0.00 0.00

06.2.7.b Propellant (Chemical) 0.00 0.0% 0.00 0.00

06.2.7.c RCS Propellant 65.58 0.0% 0.00 65.58

06.2.7.c.a RCS Used 1 63.98 63.98 0.0% 0.00 63.98

06.2.7.c.b RCS Residuals 1 1.60 1.60 0.0% 0.00 1.60

06.2.7.d RCS Pressurant 1 0.31 0.31 0.0% 0.00 0.31

06.2.8 Structures and Mechanisms 148.88 18.0% 26.80 175.68

06.3 Sample Return Craft 93.30 18.1% 16.91 110.21
NASA/TM—2009-215825 44 

5.6.4 Propulsion and Propellant Management Trades 

The primary trades of this study were done on thruster system choice. For each of these thruster types, the 

impact on thruster choice will be on the total amount of Xe propellant needed to complete the mission. The 

three types of electric propulsion systems traded out in this study were the following: 

 NEXT (1+1)—Case 1, 1a, 3a 

− 425 kg Xe for mission 

− 7 kW 

− Larger footprint, more susceptible to contamination 

 BPT-4000 (2+1)—Case 2 

− 720 kg Xe for mission 

− 6 kW 

− Smaller footprint, less susceptible contamination 

 HiVHAC (2+1)—Case 3

− 620 kg Xe for mission 

− 7 kW 

− Smaller footprint, less susceptible contamination 

5.6.5 Propulsion and Propellant Management Analytical Methods 

Because the propulsion subsystems were assembled from existing components where possible, the analysis 

performed consisted primarily of maintaining a mass roll up for the various subassemblies. The first 

propulsion operation was performed with a variety of electric propulsion thruster options. A commercial 

high power Hall Thruster was used to establish a baseline vehicle for mission analysis and mass assessment. 

Subsequent trades were performed with two different EP thruster technologies under development at NASA 

GRC. The performance and physical characteristics of these thrusters were obtained directly from their 

development groups. Additionally, propellant management systems (PMS) were used based on breadboard 

systems also currently being developed at GRC through the In Space Program office. Real data for the PMS 

were used where available. Otherwise, it was obtained from development reports. 

The vehicle’s attitude and RCS was comprised of technically mature components with flight history. This 

propulsion system was a mass roll up of physical characteristics obtained from hardware providers. 

For the ‘delicate’ landing operation, a notional cold-gas thruster-based propulsion system was developed. 

These cold-gas thrusters use the Xe propellant from the electric propulsion subsystem to provide the very 

small and controllable thrust levels required for landing on an asteroid (see Main Engine line in table 5.11). 

Commercially available nitrogen-based cold gas thrusters were used for mass estimates while the propellant 

management subsystem for the landing thruster pod was sized similarly to the other PMS elements. 

The primary analysis that was actively performed was to determine the propellant tanks sizes based on 

propellant conditions over the mission duration. The tank requirements were determined using propellant 

density and storage pressure through Hoop Stress Analysis. These requirements were then used to select the 

best match from the PSI and Arde, Inc storage tank catalogs. Thermal control elements (heaters, insulation) 

were then added based on surface area of tanks and propellant lines 

Once the storage tank(s) were selected, the helium pressurization requirements were determined. A 

conventional He pressurization system configuration was used, based on our experience with previous 

lander and Orion Service Module studies. 
NASA/TM—2009-215825 45 

5.6.5.2 The NEXT Thruster Characteristics 

 0.54 to 6.9 kW thruster input power 

 Ring-cusp electron bombardment discharge chamber 

 36 cm beam diameter, 2-grid ion optics 

 Beam current at 6.9 kW: 3.52 A 

 Maximum specific impulse > 4170 sec 

 Maximum thrust > 236 mN 

 Peak efficiency > 70% 

 Xe throughput > 300 kg, (450 kg is the qualification level) 

− Analysis-based capability >450 kg 

 Thruster Mass is 12.7 kg (13.5 kg with cable harnesses) 

Shown in Figure 5.22 is the NEXT thruster in the Prototype Model Thruster (PM1) in Performance 

Acceptance Test. 

Figure 5.22—PM 1 Performance Acceptance Test 

Gimbal Overview 

 Breadboard gimbal 

− Designed and fabricated by Swales Aerospace 

− Flight-like design using JPL-approved materials with certifications 

 Stepper motors have space-rated option 

− Mass < 6 kg 

− Two-axis range of motion: ±19°, ±17° 

 Successful functional testing with PM1 engine 

 Gimbal passed two qual-level vibration tests and low-level shock tests with minor issues (fastener 

backout) 

 Good baseline—few if any modifications needed to move into qual program 

− Need to perform torque margin tests with harness and propellant line routing 
NASA/TM—2009-215825 46 

Figure 5.23—Gimbal Performance Test and CAD Illustration 

NEXT Power Processing Unit (PPU) 

 EM PPU build by L3 Communications Corporation (L3) ETI 

 Modular beam supply and improved packaging provides performance and predictability benefits 

over NSTAR approach 

 Digital Control Interface Unit (DCIU) to be integrated in next development phase 

Figure 5.24—NEXT Thruster PPU schematic 
NASA/TM—2009-215825 47 

5.6.5.3 NEXT Propellant Management System (PMS) 

Figure 5.25—NEXT Propellant Management System 

 All PMS assemblies are complete 

− Two HPA’s, one Flight-like 

− Three LPA’s, one Flight-like 

− Nonflight assemblies are identical except for use of lower cost equivalent parts 

 All assemblies have completed functional tests 

 Flight-like LPA and HPA successfully completed qual-level vibration testing and post-vibe 

functionals 

 Qual-level thermal/vacuum testing is pending 

 Xe Storage tanks based on COTS unit from ATK-PSI Inc. 

− Carbon Overwrapped tank with Ti liner 

− Derived from model no. 80465-1 

• Size: 0.42 m by 0.75 m L (16.5 in. by 29.6 in. L) 

• Manufacturers expected operating pressure (MEOP)/burst pressure = 198.2/310.3 bar 

(2875/4500 psig) 

− Minor size changes to match propellant load 

 Hydrazine storage tank based on COTS unit design from ATK-PSC Inc. 

− Ti metallic tanks w/polymer diaphragm for blow-down pressurization 

• Size: 0.47 m dia. (18.6 in. diameter) 

− Derived from ATK/PSI Model No. 80439-1 
NASA/TM—2009-215825 48 

 

Figure 5.26—Hydrazine Tank (Left), Xe Tank (Right) 

5.6.6 Propulsion and Propellant Management Risk Inputs 

No Risks gathered in this study. 

5.6.7 Propulsion and Propellant Management Recommendation 

Design is recommended. 

5.7 Thermal Control 

Objective: To provide spreadsheet based models capable of estimating the mass and power requirements of 

the various thermal systems. The thermal modeling provides power and mass estimates for the various 

aspects of the vehicle thermal control system based on a number of inputs related to the vehicle geometry, 

flight environment and component size. The system consists of the following elements 

 Electric heaters 

 MLI 

 Thermal paint 

 Radiator with louvers 

 Thermal Control System (sensors, switches, data acquisition) 

5.7.1 Thermal Requirements 

The thermal requirements for the mission were to provide a means of cooling and heating of the S/C 

equipment during transit to and operation on the Asteroid’s surface in order to remain within their 

maximum and minimum temperature requirements. 

The maximum heat load to be rejected by the thermal system was 586 W, and the desired operating 

temperature for the electronics was 300 and 250 K for S/C structure. The S/C was required to survive and 

operate through any nighttime or shadow periods; therefore a heating system was also required. 

5.7.2 Thermal Assumptions 

The assumptions utilized in the analysis and sizing of the thermal system were based on the operational 

environment, both in transit to the asteroids and operation on the asteroid’s surface. The following 

assumptions were utilized to size the thermal system. 

 Moon surface operation: Day and night 

 Radiator designed to see deep space with minimal view factors to the asteroid surface. 
NASA/TM—2009-215825 49 

 The maximum angle of the radiator to the Sun was 25°. 

 The radiator temperature was 320 K. 

 A redundant radiator was used to account for vehicle orientation on the surface and to increase 

overall reliability. 

 MLI was used to insulate the S/C to minimize heat transfer to and from the surroundings. 

 Electric heaters and the radiator louvers were used to maintain the desired internal temperature of 

the S/C 

5.7.3 Thermal Design and MEL 

The thermal system is used to remove excess heat from the electronics and other components of the system 

as well as provide heating to thermally sensitive components throughout shadow or nighttime periods. 

Excess heat is collected from a series of aluminum cold plates located throughout the interior of the S/C. 

These cold plates have heat pipes integrated into them. The heat pipes transfer heat from the cold plates to 

the radiator, which radiates the excess heat to space. The portions of the heat pipes that extend from the 

communications box and are integrated to the radiator are protected with a micro meteor shield. The system 

utilize a louver system on the radiators to regulate the internal temperature and to insulate the radiators 

during the asteroid nighttime. 

Two radiators were used to provide redundancy and margin as well as account for the unknown landing 

orientation of the S/C. This added margin insures against unforeseen heat loads, degradation of the radiator 

due to asteroid dust buildup and increased view factor toward any other thermally hot body not accounted 

for in the analysis. 

Table 5.11—Thermal MEL for Baseline Case 1 

WBS Description QTY

Unit 

Mass CBE Mass Growth Growth Total Mass

Number NEA Sampler (August 2008) - Case 1 (kg) (kg) (%) (kg) (kg)

06 NEA Sampler Spacecraft 1138.27 9.9% 113.05 1251.32

06.1 Science Payload 25.20 16.1% 4.06 29.26

06.2 Lander Spacecraft 1019.77 9.0% 92.07 1111.85

06.2.1 Attitude Determination and Control 44.28 21.1% 9.36 53.64

06.2.2 Command and Data Handling 29.30 27.2% 7.96 37.26

06.2.3 Communications and Tracking 24.20 27.1% 6.55 30.75

06.2.4 Electrical Power Subsystem 115.50 15.7% 18.19 133.69

06.2.5 Thermal Control (Non-Propellant) 40.11 18.0% 7.22 47.33

06.2.5.a Active Thermal Control 6.65 18.0% 1.20 7.85

06.2.5.a.a Heaters 20 0.25 5.00 18.0% 0.90 5.90

06.2.5.a.b Thermal Control/Heaters Circuit 2 0.20 0.40 18.0% 0.07 0.47

06.2.5.a.c Data Acquisition 1 1.00 1.00 18.0% 0.18 1.18

06.2.5.a.d Thermocouples 25 0.01 0.25 18.0% 0.05 0.30

06.2.5.a.e Misc#1 1 0.00 0.00 18.0% 0.00 0.00

06.2.5.a.f Misc#2 1 0.00 0.00 18.0% 0.00 0.00

06.2.5.b Passive Thermal Control 26.87 18.0% 4.84 31.71

06.2.5.b.a Heat Sinks 2 3.46 6.93 18.0% 1.25 8.17

06.2.5.b.b Heat Pipes 1 2.93 2.93 18.0% 0.53 3.46

06.2.5.b.c Radiators 1 10.06 10.06 18.0% 1.81 11.87

06.2.5.b.d MLI (Multi Layer Insulation) 1 3.84 3.84 18.0% 0.69 4.53

06.2.5.b.e Temperature sensors 50 0.01 0.50 18.0% 0.09 0.59

06.2.5.b.f Phase Change Devices 1 0.00 0.00 18.0% 0.00 0.00

06.2.5.b.g Thermal Coatings/Paint 1 0.95 0.95 18.0% 0.17 1.12

06.2.5.b.h Micro Meteor shielding 1 0.00 0.00 18.0% 0.00 0.00

06.2.5.b.i Spacecraft RTG MLI 1 0.00 0.00 18.0% 0.00 0.00

06.2.5.b.j Spacecraft Engine MLI 1 1.66 1.66 18.0% 0.30 1.96

06.2.5.c Semi-Passive Thermal Control 6.59 18.0% 1.19 7.78

06.2.5.c.a Louvers 1 5.79 5.79 18.0% 1.04 6.84

06.2.5.c.b Thermal Switches 4 0.20 0.80 18.0% 0.14 0.94

06.2.5.c.c Misc#1 0 0.00 0.00 18.0% 0.00 0.00

06.2.5.c.d Misc#2 0 0.00 0.00 18.0% 0.00 0.00

06.2.5.c.e Misc#3 0 0.00 0.00 18.0% 0.00 0.00

06.2.6 Propulsion 192.10 8.3% 16.00 208.09

06.2.7 Propellant 425.41 0.0% 0.00 425.41

06.2.8 Structures and Mechanisms 148.88 18.0% 26.80 175.68

06.3 Sample Return Craft 93.30 18.1% 16.91 110.21
NASA/TM—2009-215825 50 

5.7.4 Thermal Trades 

No significant design trades were made between components of the thermal control system. 

5.7.5 Thermal Analytical Methods 

The analysis performed to size the thermal system is based on first principle heat transfer from the S/C to 

the surroundings. This analysis takes into account the design and layout of the thermal system and the 

thermal environment to which heat is being rejected to or insulated from. 

Environmental Models 

Solar Intensity Based on S/C Location components were sized for worst case operating conditions, Heat 

Rejection: Near Earth, Minimum Temperature: near Earth Asteroid Orbital Location 

Systems Modeled 

 Micro meteor shielding on radiator 

 Radiator panels 

 Thermal control of propellant lines and tanks 

 S/C insulation 

 Avionics, and Power Management and Distribution (PMAD) cooling 

Table 5.12—Thermal System Data Exchange 

Input Output

S/C dimensions (length, diameter) Heat pipe length and mass 

Power management and electronics dimensions Cold plate size and mass 

Waste heat load to be rejected 

Distance from the sun and S/C orientation 

View factor to the SAs and their temperature 

Radiator size and mass 

S/C insulation mass and thickness 

Thermal system components mass 

Propellant tank dimensions and operating temperature Propellant tanks insulation mass and heater power level 

Propellant line lengths and operating temperature Propellant line insulation mass and heater power level 

Radiator Sizing 

The radiator panel area has been modeled along with an estimate of its mass. The model was based on first 

principles analysis of the area needed to reject the identified heat load to space. From the area, a series of 

scaling equations were used to determine the mass of the radiator within the asteroid environment. Asteroid 

orbit 1 AU thermal environment was used to size the radiator. 

Table 5.13—Thermal System Radiator Sizing Assumptions 

Variable Value

Radiator solar absorptivity ................................................... 0.14

Radiator emissivity ............................................................. 0.84

Radiator Sun angle ............................................................... 70°

Radiator operating temperature ......................................... 320 k

Total radiator dissipation power .................................... 656.5 W

View Factor to SA ................................................................ 0.10

View Factor to Earth ............................................................ 0.10

Louvers are active or passive devices that regulate the amount of heat rejected by the radiator. Active 

controlled louvers use temperature sensors and actuators to control the louver position. Passive controlled 

louvers commonly use a bimetallic spring that opens and closes the louver based on temperature. The 

louver specific mass is 4.5 kg/m2
NASA/TM—2009-215825 51 

Figure 5.27—Schematic of the Louver Prototype 

Thermal Analysis Propellant Lines and Tanks 

Power requirements and mass have been modeled. This modeling included propellant tank MLI and heaters 

and propellant line insulation and heaters. 

Figure 5.28—MLI 

The model was based on a first principles analysis of the radiative heat transfer from the tanks and 

propellant lines through the S/C structure to space. The heat loss through the insulation set the power 

requirement for the tank and line heaters. The 1 AU thermal environment was used to calculate the heat 

loss. Assumptions used: 
NASA/TM—2009-215825 52 

Table 5.14—Thermal System Tank Insulation Sizing Assumptions 

Variable Value 

Tank surface emissivity (εt) ................................................................................................... 0.1 

MLI emissivity (εi) ................................................................................................................ 0.07 

MLI material ............................................................................................................................. Al 

MLI material density (ρi) .......................................................................................... 2,770 kg/m3

Internal tank temperature (Ti) ........................................................................................... 300 K 

MLI layer thickness (ti) ................................................................................................ 0.025 mm 

Number of insulation layers (ni) .............................................................................................. 10 

MLI layer spacing (di) ..................................................................................................... 1.0 mm 

Tank immersion heater mass & power level ...................................... 1.02 kg @ up to 1,000 W 

S/C inner wall surface emissivity ......................................................................................... 0.98 

S/C outer wall surface emissivity ......................................................................................... 0.93 

Line foam insulation conductivity ......................................................................... 0.0027 W/m K 

Line foam insulation emissivity ............................................................................................ 0.07 

Propellant line heater specific mass & power ................................ 0.143 kg/m @ up to 39 W/m 

Line foam insulation density ......................................................................................... 56 kg/m3

Thermal Analysis—S/C Insulation 

The mass of the S/C MLI insulation was modeled to determine the mass of the insulation and heat loss. The 

model was based on a first principles analysis of the heat transfer from the S/C through the insulation to 

space. Nighttime asteroid surface thermal environment was used to size the insulation. Two types of heaters 

were considered, Radioisotope Heater Unit (RHU), and electrical heaters. Assumptions used: 

Table 5.15—Thermal System Tank Insulation Sizing Assumptions 

Variable Value 

S/C MLI material ..................................................................................... Al 

S/C MLI material density (ρisc) ................................................ 2,770 kg/m3

MLI layer thickness (ti) ............................................................... 0.025 mm 

Number of insulation layers (ni) ........................................................... 100 

MLI layer spacing (di) ................................................................... 1.0 mm 

S/C Radius (rsc) .......................................................................... 0.825 m

Thermal Analysis—PMAD Cooling 

Thermal control of the electronics and Active Thermal Control System (ATCS) is accomplished through a 

series of cold plates and heat pipes to transfer the excess heat to the radiators. The model for sizing these 

components was based on a first principles analysis of the area needed to reject the identified heat load to 

space. From the sizing, a series of scaling equations were used to determine the mass of the various system 

components. Assumptions used: 

Table 5.16—Thermal System PMAD Cooling Sizing Assumptions 

Variable Value 

Cooling plate & lines material .................................................................. Al 

Cooling plate & lines material density ..................................... 2,770 kg/m3

Number of cooling plates ........................................................................... 4 

Cooling plate lengths .......................................................................... 0.5 m 

Cooling plate widths ........................................................................... 0.5 m 

Cooling plate thickness ...................................................................... 5 mm 

Heat pipe specific mass .............................................................. 0.15 kg/m 

5.7.6 Thermal Risk Inputs 

None 

5.7.7 Thermal Recommendation 

See Design 
NASA/TM—2009-215825 53 

6.0 SOFTWARE COST ESTIMATION 

6.1 Objectives 

6.2 Assumptions 

6.2.1 SW Sizing Assumptions 

 Programming Language is Ada 

 Two RAD750 processors 

 Complex autonomy 

 RTOS on RAD750 compact PCI processor board 

 Ground software not included 

Figure 6.1—Typical RAD750 Processor and Software Environment 

6.2.2 S/C Bus Functions 

 Attitude Determination & Control (ACS): 5500 Single Line of Codes (SLOC) 

− This estimate includes: Sun sensor, Star Tracker, Rate Gyros, Complex Ephemeris, Kinematic 

Integration, Error Determination, Thruster Control, Reaction Wheel Control, Orbit Propagation 

− C&DH: 750 SLOCs 

• Command and telemetry processing 

− On-board autonomy (complex assumed): 4100 SLOCs 

− Fault detection (on-board systems monitoring and correction): 1650 SLOCs 

− Power Management and Thermal Control: 500 SLOCs 

6.2.3 Payload Functions 

 Sensor Processing: 670 SLOCs 

 Data Reduction and Transmission: 200 SLOCs 

6.3 Software Cost 

Software Development 

 Estimated 13370 SLOCs for on-board software 

 Estimate at least 32,000 SLOCs for ground support software development, including 
NASA/TM—2009-215825 54 

 Avionics simulation software 

 Data collection and reduction software 

 Ground station interface and simulation software 

Support Equipment 

 Estimate 4 engineering units (RAD750) 

 RAD750 Compiler and run-time kernel 

 Estimate 10 PCs/Workstations 

 GUI builder, software development tools 

 I/O cards, card cage, and drivers 

Table 6.1—Software Costing estimate 

S/C bus functions Code size, 

K words

Code size by 4, 

K words

SLOC

C&DH (C&T processing) 2.8 11.0 750 

AD&C 

(Sun sensor, Star Tracker, rate gyros, complex 

ephemeris, kinematic integration, error determination, 

thruster control, reaction wheel control, orbit propagation) 

20.2 80.8 5500 

Onboard autonomy 

 Complex autonomy 

15.0 60.0 4100 

fault detection 

 Onboard system monitoring 

 Fault correction 

4.0 

2.0 

16.0 

8.0 

1100 

550 

Power management and thermal control 

 Residing in onboard computers assumed 1.8 7.2 500 

Subtotal 45.8 183.2 12500

Total (including payload functions): 13370 SLOCs. 

7.0 COST, RISK AND RELIABILITY 

7.1 Costing 

The following section contains a draft Cost Estimate (all cost in FY09$M). The S/C cost estimates 

represents prime contractor cost. Assumes a proto-flight development. Flight spares are included where 

appropriate. Mission operations costs include 2 yr data analysis post sample return per NF 2009 AO 

(5.1.5.2—Curation of Returned Samples). Launch Services were not included per NF 2009 AO. 

Table 7.1—NEXT 1+1, Added Science, Case 1a, LCC Cost (FY09 $M) 

Near Earth Asteroid Sample Return

NF Mission

All costs in FY09 $M 

NASA Project Office/Technical Oversight 25 5% of all other costs 

Phase A 21 5% of S/C cost and fee 

S/C (without science instruments) 183 50 percentile estimate (DD and FH only) 

Science instruments and SRC 98 Expanded science package (DD and FH only) 

S/C systems integration and wraps 119 Includes integrating S/C and science instruments 

S/C Prime Contractor Fee (10%) 30 Not applied to science instruments costs. 

Mission operations 54 Based on DAWN (includes data analysis) 

Life Cycle Cost Estimate 530 *Does not include Launch Services per 2009 NF AO 

 

NF 2009 Cost Cap 650 NF 2009 AO ('09 NFAO) 

Launch services: 4 m fairing/med performance 40 Per 2009 NFAO (5.9.2—Launch Services) 

Propulsion system: NEXT 15 Per 2009 NFAO (5.9.3—Propulsion Technology Infusion) 

Adjusted cost cap 705 

Reserves based on LCC estimate 175 33% Reserve 
NASA/TM—2009-215825 55 

7.1.1 Detailed Cost Breakdown by WBS 

WBS element Element name 

DDT&E 

total by 

$M

Flight

hardware

by $M

Mfg/DDT&E

total

by $M

06.1.1 Science Payload $48.5 $23.9 $72.4 

06.1.1 Arm Mounted Science Instruments 

06.1.1.a Panoramic / microscopic color imager $12.4 $5.3 $17.7 

06.1.1.b APXS $3.2 $1.4 $4.6 

06.1.1.c LAMS $11.1 $4.7 $15.8 

06.1.2 Body Mounted Science Instruments 

06.1.2.a Approach/Hazard Avoidance/Landing Lidar $7.2 $6.2 $13.4 

06.1.2.b Neutron Detector/Gamma Ray Spect. $1.4 $0.6 $2.1 

06.1.2.c wide narrow field imager $6.0 $2.6 $8.5 

06.1.2.d Ground penetrating radar $7.2 $3.1 $10.3 

06.2.1 Attitude Determination and Control $17.1 $13.1 $30.2 

06.2.1.a.a Sun Sensors $1.2 $2.9 $4.2 

06.2.1.a.b Reaction Wheels $2.2 $2.4 $4.5 

06.2.1.a.c Star Trackers $1.4 $1.7 $3.1 

06.2.1.a.d IMU $5.6 $2.8 $8.4 

06.2.1.a.e Laser Altimeter (from Science Payload) $6.7 $3.2 $9.9 

06.2.2 Command & Data Handling $16.6 $4.6 $21.2 

06.2.2.a.a Flight Computer $2.3 $2.4 $4.7 

06.2.2.a.c Data Interface Unit $0.5 $0.3 $0.8 

06.2.2.a.e Operations Recorder $0.1 $0.1 $0.2 

06.2.2.a.f Command and Control Harness (data) $6.2 $0.6 $6.8 

06.2.2.a.g Shared DPU (From APL Science Instruments) $2.3 $1.2 $3.5 

06.2.2.b Instrumentation & Wiring $0.0 $0.0 $0.0 

 Flight Software/Firmware $5.1 $0.0 $5.1 

06.2.3 Communications and Tracking $17.2 $7.6 $24.8 

06.2.3.a X/Ka High Gain Antenna 

06.2.3.a.a Transmitter/Receiver $2.1 $1.2 $3.3 

06.2.3.a.b Power Amp $2.6 $1.7 $4.3 

06.2.3.a.c Switch Unit $2.5 $1.0 $3.5 

06.2.3.a.d Antenna $0.9 $0.4 $1.2 

06.2.3.a.h Cabling $1.3 $0.2 $1.5 

06.2.3.a.i Diplexer $0.5 $0.4 $1.0 

06.2.3.a.j Coupler $1.0 $0.5 $1.5 

06.2.3.a.k Misc#1 $1.9 $0.7 $2.7 

06.2.3.a.l Misc#2 $0.3 $0.0 $0.3 

06.2.3.b Ka-band Antenna 

06.2.3.b.a Transponder $2.4 $0.0 $2.4 

06.2.3.b.b RF Assembly $0.1 $0.0 $0.1 

06.2.3.b.d Antenna $0.9 $1.1 $2.0 

06.2.3.c.a Coaxial Cable $0.7 $0.3 $1.0 

06.2.4 Electrical Power Subsystem $18.3 $9.9 $28.2 

06.2.4.a Solar Arrays $0.0 $0.0 $0.0 

06.2.4.a.a Solar Array Mass (cells and structure only) $10.1 $6.4 $16.6 

06.2.4.a.b Solar Array Gimbals $1.5 $0.8 $2.2 

06.2.4.b.a Power management/control electronics $3.5 $1.5 $5.0 

06.2.4.c Power Cable and Harness Subsystem $1.4 $1.0 $2.3 

06.2.4.d Battery System $1.8 $0.3 $2.1 

06.2.5 Thermal Control (Non-Propellant) $6.4 $1.2 $7.6 

06.2.5.a Active Thermal Control $0.5 $0.4 $0.9 

06.2.5.b Passive Thermal Control $5.9 $0.7 $6.6 

06.2.6 Propulsion $36.8 $15.3 $52.1 

06.2.6.a.a Primary EP Thrusters $8.5 $4.3 $12.8 

06.2.6.a.c.c Gimbal $2.6 $1.7 $4.2 

06.2.6.b.a Xe propellant tank(s) $0.3 $0.3 $0.7 

06.2.6.b.b High Pressure Feed System $5.4 $1.4 $6.8 

06.2.6.b.e Temperature sensors $0.1 $0.0 $0.1 

06.2.6.c Power Processing Unit (PPU) 

06.2.6.c.a PPU Mass $13.0 $5.5 $18.5 

06.2.6.d.a.a Main Engine $0.0 $0.0 $0.1 

06.2.6.f Reaction Control System Hardware 

06.2.6.f.a RCS Thruster Subassembly $0.2 $0.1 $0.3 

06.2.6.g RCS Propellant Management 

06.2.6.g.a Fuel Tanks $0.3 $0.3 $0.5 

06.2.6.g.c Feed and Pressurizations Systems $6.4 $1.7 $8.1 

06.2.7 Propellant $0.0 $0.0 $0.0 
NASA/TM—2009-215825 56 

WBS element Element name 

DDT&E 

total by 

$M

Flight

hardware

by $M

Mfg/DDT&E

total

by $M

06.2.8 Structures and Mechanisms $32.3 $11.6 $43.9 

06.2.8.a Structures $10.6 $5.1 $15.7 

06.2.8.b.f.a Arm $7.9 $3.4 $11.3 

06.2.8.b.f.b Mechanisms $0.9 $0.5 $1.4 

06.3 Sample Return Craft $12.8 $2.7 $15.5 

Subtotal $193.3 $87.1 $280.4 

SYSTEMS INTEGRATION $98.0 $20.9 $118.9 

1.6.2.2 IACO $8.2 $3.4 $11.6 

1.6.2.2 STO $9.6 $0.0 $9.6 

1.6.2.2 GSE Hardware $17.1 $0.0 $17.1 

1.6.2.1 SE&I $29.8 $12.2 $42.0 

1.6.2.1 PM $18.0 $5.3 $23.2 

1.6.2.2 LOOS $15.3 $0.0 $15.3 

TOTAL PRIME COST $291.3 $108.0 $399.3

7.1.2 System Integration Wraps Defined 

The Integration, Assembly and Checkout (IACO) element contains all labor and material required to 

physically integrate (assemble) the various subsystems into a total system. Final assembly, including 

attachment, and the design and manufacture of installation hardware, final factory acceptance operations, 

packaging/crating, and shipment are included. IACO charged to DDT&E represents those costs incurred for 

the integration, assembly, and checkout of major test articles. IACO charged to the flight unit includes those 

same functions applied to the actual flight unit. 

This item excludes the engineering effort required to establish the integration, assembly, and checkout 

procedures necessary for this effort. These engineering efforts are covered under systems engineering and 

integration. 

The System Test Operations (STO) element includes development testing and the test effort and test 

materials required for qualification and physical integration of all test and qualification units. Also included 

is the design and fabrication of test fixtures. 

Specifically included are tests on all STO to determine operational characteristics and compatibility with 

the overall system and its intended operational parameters. Such tests include operational tests, design 

verification tests, and reliability tests. Also included are the tests on systems and integrated systems to 

verify acceptability for required mission performance. These tests are conducted on hardware that has been 

produced, inspected, and assembled by established methods meeting all final design requirements. Further, 

system compatibility tests are included, as well as, functions associated with test planning and scheduling, 

data reduction, and report preparation. 

Functional elements associated with Ground Support Equipment (GSE) include the labor and materials 

required to design, develop, manufacture, procure, assemble, test, checkout, and deliver the equipment 

necessary for system level final assembly and checkout. Specifically, the equipment utilized for integrated 

and/or electrical checkout, handling and protection, transportation, and calibration, and items such as 

component conversion kits, work stands, equipment racks, trailers, staging cryogenic equipment, and many 

other miscellaneous types of equipment are included. 

Specifically excluded is the equipment designed to support only the mission operational phase. 

The functions included in the Systems Engineering and Integration (SE&I) element encompass: (1) the 

system engineering effort to transform an operational need into a description of system requirements and/or 

a preferred system configuration; (2) the logistics engineering effort to define, optimize, and integrate 

logistics support considerations to ensure the development and production of a supportable and cost 

effective system; and (3) the planning, monitoring, measuring, evaluating, and directing of the overall 

technical program. Specific functions include those for control and direction of engineering activities, 

cost/performance trade-offs, engineering change support and planning studies, technology utilization, and 
NASA/TM—2009-215825 57 

the engineering required for safety, reliability, and quality control and assurance. Also included is the effort 

for system optimization, configuration requirements analyses, and the submittal and maintenance of 

Interface Control Documents (ICDs). 

Excluded from the SE&I element are those functions which are identifiable to subsystem SE&I. 

Elements included in the Program Management (PM) function consist of the effort and material required 

for the fundamental management direction and decision-making to ensure that a product is developed, 

produced, and delivered per requirements 

Specifically included are direct charges for program administration, planning and control, scheduling and 

budgeting, contracts administration, and the management functions associated with engineering, 

manufacturing, support, quality assurance, configuration and project control, and documentation. 

The PM element sums all of the effort required for planning, organizing, directing, coordinating, and 

controlling the project to help ensure that overall objectives are accomplished. This element also includes 

the effort required to coordinate, gather, and disseminate information. 

Excluded from the PM element are those functions commonly charged to subsystem level activities. 

7.2 Risk Analysis and Reduction 

7.2.1 Assumptions 

7.2.2 Risk List 

7.2.3 Risk Summary 

Risks 

 Dust and debris impact on 

 Flight system 

 Instruments 

 Gridded Ion thrusters 

 SA 

Mitigation 

 Single Fault Tolerant Design for most flight systems 

 Mission Scenario 

− Spiral down and Map Asteroid 1 

− Sample Asteroid 1 

− Perform Asteroid 1 sample return at Earth 

− Repeat for second asteroid 

− Ensures first asteroid sample not at risk during second asteroid sampling 

 SA away from surface 

These risks, with proper pro-active planning can be mitigated early to avoid becoming problems late in the 

development life cycle. 

7.3 Reliability 

None performed for this study. 
NASA/TM—2009-215825 58 

8.0 TRADE SPACE ITERATIONS 

Two different trade options were explored in the five cases run on the multiple small body sample return 

study. The first option looked at using different thruster technologies. Cases 1, 2, and 3 used the three 

different thruster technologies NEXT, BPT-4000 and HiVAC respectively. A secondary set of trades was 

run on the science payload and the SRC to explore the ceiling and basement of science capabilities of this 

mission. In cases 1a and 3a, a slightly larger single SRC was returned to the Earth from the Near Earth 

Asteroids, rather than two SRC. Both cases used the NEXT thruster but case 1a delivered a larger science 

payload while case 3a delivered the standard baseline science package used in cases 1, 2 and 3. 

The baselined case 1a and case 3a returned a single sample capsule of a slightly larger size than cases 1, 2 

and 3. 

Table 8.1—Spacecraft Case Comparison 

Case 1 Case 1a Case 2 Case 3 Case 3a

Launch 

vehicle 

Atlas 401 Atlas 401 Atlas 521 Atlas 401 Atlas 401 

Thruster NEXT NEXT BPT-4000 HiVAC NEXT 

Science 

payload 

Baseline science 

package 

Super science 

package 

Baseline science 

package 

Baseline Science 

package 

Baseline science 

package 

SRC 2 full size 1 slightly larger size 2 full size 2 full size 1 slightly larger size 

Table 8.2—Spacecraft Total Mass Comparison 

Case 1 Case 1a Case 2 Case 3 Case 3a

WBS Main Subsystems

Total Mass 

(kg)

Total Mass 

(kg)

Total Mass 

(kg)

Total Mass 

(kg)

Total Mass 

(kg)

01 Moon Sampler Spacecraft 1251.3 1262.8 1561.8 1416.8 993.8

06.1 Science Payload 29.3 60.3 29.3 29.3 29.3

06.2 Moon Sampler Lander 1111.8 1141.0 1422.3 1277.3 903.1

06.2.1 Attitude Determination and Control 53.6 53.6 53.6 53.6 53.6

06.2.2 Command and Data Handling 37.3 37.3 37.3 37.3 37.3

06.2.3 Communications and Tracking 30.7 30.7 30.7 30.7 30.7

06.2.4 Electric Power 133.7 133.7 133.7 133.7 133.7

06.2.5 Thermal Control 47.3 47.3 47.3 47.3 47.3

06.2.6 Propulsion 208.1 209.3 217.2 179.5 154.3

06.2.7 Propellant 425.1 451.8 726.9 620.5 272.8

06.2.8 Structures and Mechanisms 175.7 176.9 175.2 174.3 173.2

06.3 Sample Return Craft (total, empty) 110.2 61.4 110.2 110.2 61.4

06.3.1 Electrical Power Subsystem 7.2 7.2 7.2 7.2 7.2

06.3.2 Thermal Control (Non-Propellant) 5.4 5.4 5.4 5.4 5.4

06.3.3 Structures and Mechanisms 97.6 48.8 97.6 97.6 48.8

Estimated Spacecraft Dry Mass 826.2 810.9 834.9 796.3 721.0

Estimated Spacecraft Wet Mass 1251.3 1262.8 1561.8 1416.8 993.8

System LeveL Growth Calculations

Dry Mass w/ Desired System Level Growth 927.1 913.7 937.5 884.9 799.5

Additional Growth (carried at system level) 100.9 102.8 102.6 88.6 78.5

Total Wet Mass with Growth 1352.2 1365.5 1664.4 1505.4 1072.3

Available Launch Performance to C3 (kg) 1375.4 1421.9 1686.4 1600.0 1200.0

Launch margin available (kg) 23.2 56.4 22.0 94.6 127.7

Estimated Spacecraft Inert Mass (for traj.) 1021.5 1009.8 1058.6 996.0 856.0

000 0

Sample Return Craft Total Mass

Total Mass 

(kg)

Total Mass 

(kg)

Total Mass 

(kg)

Total Mass 

(kg)

Total Mass 

(kg)

Estimated Sample Return Craft Mass 110.2 61.4 110.2 110.2 61.4

Total with System Level Growth 121.3 67.5 121.3 121.3 67.5

Number of Sample Return Craft 2122 1

Total Mass per Sample Return Craft (empty) 60.6 67.5 60.6 60.6 67.5

Total Mass, Sample Returned 1111 1

Total Mass, Sample Return Capsule (Full) 61.6 68.5 61.6 61.6 68.5

Spacecraft Total Mass Comparison
NASA/TM—2009-215825 59 

8.1 Case 1—Baseline 

Case 1 using the NEXT thruster system for electric propulsion was chosen as the baseline for this study 

report and is documented in detail in the subsystem sections. 

Table 8.3—Case 1—System Summary 

Table 8.4—Case 1—Sample Return Craft Summary 

GLIDE

container Mars_Moons_Sampler: NEA_sampler_case1

Spacecraft Master Equipment List Rack-up (Mass)

COMPASS S/C 

Design 

WBS Main Subsystems CBE Mass (kg)

Growth 

(kg) Total Mass (kg)

Aggregate 

Growth (%)

01 Asteroids Sampler Spacecraft 1138.3 113.0 1251.3

06.1 Science Payload 25.2 4.1 29.3 16%

06.2 Asteroids Sampler Lander 1019.8 92.1 1111.8

06.2.1 Attitude Determination and Control 44.3 18.7 53.6 42%

06.2.2 Command and Data Handling 29.3 8.0 37.3 27%

06.2.3 Communications and Tracking 24.2 6.5 30.7 27%

06.2.4 Electric Power 115.5 18.2 133.7 16%

06.2.5 Thermal Control 40.1 7.2 47.3 18%

06.2.6 Propulsion 192.1 16.0 208.1 8%

06.2.7 Propellant 425.1

06.2.8 Structures and Mechanisms 148.9 26.8 175.7 18%

06.3 Sample Return Craft (total, empty) 93.3 16.9 110.2 18%

06.3.1 Electrical Power Subsystem 6.0 1.2 7.2 20%

06.3.2 Thermal Control (Non-Propellant) 4.6 0.8 5.4 18%

06.3.3 Structures and Mechanisms 82.7 14.9 97.6 18%

Estimated Spacecraft Dry Mass 713 113 826.2 16%

Estimated Spacecraft Wet Mass 1138 113 1251.3

System LeveL Growth Calculations Total Growth

Dry Mass Desired System Level Growth 713 214 927.1 30%

Additional Growth (carried at system level) 101 14%

Total Wet Mass with Growth 1138 214 1352.2

Available Launch Performance to C3 (kg) 1375.4

Launch margin available (kg) 23.2

Estimated Spacecraft Inert Mass (for traj.) 808 214 1021.5

Sample Return Craft Total Mass CBE Mass (kg) Growth (kg) Total Mass (kg)

Aggregate 

Growth (%)

Estimated Sample Return Craft Mass 93.3 16.9 110.2 18%

Total with System Level Growth 93 28 121.3 30%

Number of Sample Return Craft 2

Total Mass per Sample Return Craft (empty) 60.6 kg

Total Mass, Sample Returned 1 kg

Total Mass, Sample Return Capsule (Full) 61.6 kg
NASA/TM—2009-215825 60 

8.2 Case 1a 

Case 1a uses the same NEXT thrusters to perform all the interplanetary burns as Case 1. The mission still 

carries a single SRC and takes science from the two asteroids chosen in the mission analysis section. The 

change in this case is that the science payload used on cases 1 through 3, has been increased on Case 1a to 

what is called a super science payload. Note that in this table, the word “Moon” appears. This table was 

built off of the Mars Moon Sampler mission done previously by COMPASS and the words were not 

changed in time to make it into this report. 

Table 8.5—Case 1a—System Summary 

Table 8.6—Case 1a—Sample Return Craft Summary 

GLIDE 

container: Mars_Moons_Sampler: NEA_sampler_case1a

Spacecraft Master Equipment List Rack-up (Mass)

COMPASS S/C 

Design 

WBS Main Subsystems CBE Mass (kg)

Growth 

(kg) Total Mass (kg)

Aggregate 

Growth (%)

01 Moon Sampler Spacecraft 1154.7 108.1 1262.8

06.1 Science Payload 54.0 6.3 60.3 12%

06.2 Moon Sampler Lander 1048.7 92.3 1141.0

06.2.1 Attitude Determination and Control 44.3 18.7 53.6 42%

06.2.2 Command and Data Handling 29.3 8.0 37.3 27%

06.2.3 Communications and Tracking 24.2 6.5 30.7 27%

06.2.4 Electric Power 115.5 18.2 133.7 16%

06.2.5 Thermal Control 40.1 7.2 47.3 18%

06.2.6 Propulsion 193.3 16.0 209.3 8%

06.2.7 Propellant 451.8

06.2.8 Structures and Mechanisms 149.9 27.0 176.9 18%

06.3 Sample Return Craft (total, empty) 52.0 9.5 61.4 18%

06.3.1 Electrical Power Subsystem 6.0 1.2 7.2 20%

06.3.2 Thermal Control (Non-Propellant) 4.6 0.8 5.4 18%

06.3.3 Structures and Mechanisms 41.4 7.4 48.8 18%

Estimated Spacecraft Dry Mass 703 108 810.9 15%

Estimated Spacecraft Wet Mass 1155 108 1262.8

System LeveL Growth Calculations Total Growth

Dry Mass Desired System Level Growth 703 211 913.7 30%

Additional Growth (carried at system level) 103 15%

Total Wet Mass with Growth 1155 211 1365.5

Available Launch Performance to C3 (kg) 1421.9

Launch margin available (kg) 56.4

Estimated Spacecraft Inert Mass (for traj.) 799 211 1009.8

Sample Return Craft Total Mass CBE Mass (kg) Growth (kg) Total Mass (kg)

Aggregate 

Growth (%)

Estimated Sample Return Craft Mass 52.0 9.5 61.4 18%

Total with System Level Growth 52 16 67.5 30%

Number of Sample Return Craft 1

Total Mass per Sample Return Craft (empty) 67.5 kg

Total Mass, Sample Returned 1 kg

Total Mass, Sample Return Capsule (Full) 68.5 kg
NASA/TM—2009-215825 61 

8.3 Case 2 

Case 2 ran the same mission to the same two asteroids using the BPT-4000 thruster model for the electric 

propulsion system. Two SRCs were returned to Earth as in case 1. Note that in this table, the word “Moon” 

appears. This table was built off of the Mars Moon Sampler mission done previously by COMPASS and the 

words were not changed in time to make it into this report. 

Table 8.7—Case 2—System Summary 

Table 8.8—Case 2—Sample Return Craft Summary 

GLIDE 

container: Mars_Moons_Sampler: NEA_sampler_case2

Spacecraft Master Equipment List Rack-up (Mass)

COMPASS S/C 

Design 

WBS Main Subsystems CBE Mass (kg)

Growth 

(kg) Total Mass (kg)

Aggregate 

Growth (%)

01 Moon Sampler Spacecraft 1448.1 113.8 1561.8

06.1 Science Payload 25.2 4.1 29.3 16%

06.2 Moon Sampler Lander 1329.6 92.8 1422.3

06.2.1 Attitude Determination and Control 44.3 18.7 53.6 42%

06.2.2 Command and Data Handling 29.3 8.0 37.3 27%

06.2.3 Communications and Tracking 24.2 6.5 30.7 27%

06.2.4 Electric Power 115.5 18.2 133.7 16%

06.2.5 Thermal Control 40.1 7.2 47.3 18%

06.2.6 Propulsion 200.5 16.8 217.2 8%

06.2.7 Propellant 726.9

06.2.8 Structures and Mechanisms 148.5 26.7 175.2 18%

06.3 Sample Return Craft (total, empty) 93.3 16.9 110.2 18%

06.3.1 Electrical Power Subsystem 6.0 1.2 7.2 20%

06.3.2 Thermal Control (Non-Propellant) 4.6 0.8 5.4 18%

06.3.3 Structures and Mechanisms 82.7 14.9 97.6 18%

Estimated Spacecraft Dry Mass 721 114 834.9 16%

Estimated Spacecraft Wet Mass 1448 114 1561.8

System LeveL Growth Calculations Total Growth

Dry Mass Desired System Level Growth 721 216 937.5 30%

Additional Growth (carried at system level) 103 14%

Total Wet Mass with Growth 1448 216 1664.4

Available Launch Performance to C3 (kg) 1686.4

Launch margin available (kg) 22.0

Estimated Spacecraft Inert Mass (for traj.) 842 216 1058.6

Sample Return Craft Total Mass CBE Mass (kg) Growth (kg) Total Mass (kg)

Aggregate 

Growth (%)

Estimated Sample Return Craft Mass 93.3 16.9 110.2 18%

Total with System Level Growth 93 28 121.3 30%

Number of Sample Return Craft 2

Total Mass per Sample Return Craft (empty) 60.6 kg

Total Mass, Sample Returned 1 kg

Total Mass, Sample Return Capsule (Full) 61.6 kg
NASA/TM—2009-215825 62 

8.4 Case 3 

Case 2 ran the same mission to the same two asteroids using the HiVAC thruster model for the electric 

propulsion system. Two SRCs were returned to Earth as in case 1 and 2. Note that in this table, the word 

“Moon” appears. This table was built off of the Mars Moon Sampler mission done previously by 

COMPASS and the words were not changed in time to make it into this report. 

Table 8.9—Case 3—System Summary 

Table 8.10—Case 3—Sample Return Craft Summary 

GLIDE 

container: Mars_Moons_Sampler: NEA_sampler_case3

Spacecraft Master Equipment List Rack-up (Mass)

COMPASS S/C 

Design 

WBS Main Subsystems CBE Mass (kg)

Growth 

(kg) Total Mass (kg)

Aggregate 

Growth (%)

01 Moon Sampler Spacecraft 1301.2 115.6 1416.8

06.1 Science Payload 25.2 4.1 29.3 16%

06.2 Moon Sampler Lander 1182.7 94.6 1277.3

06.2.1 Attitude Determination and Control 44.3 18.7 53.6 42%

06.2.2 Command and Data Handling 29.3 8.0 37.3 27%

06.2.3 Communications and Tracking 24.2 6.5 30.7 27%

06.2.4 Electric Power 115.5 18.2 133.7 16%

06.2.5 Thermal Control 40.1 7.2 47.3 18%

06.2.6 Propulsion 160.8 18.7 179.5 12%

06.2.7 Propellant 620.5

06.2.8 Structures and Mechanisms 147.7 26.6 174.3 18%

06.3 Sample Return Craft (total, empty) 93.3 16.9 110.2 18%

06.3.1 Electrical Power Subsystem 6.0 1.2 7.2 20%

06.3.2 Thermal Control (Non-Propellant) 4.6 0.8 5.4 18%

06.3.3 Structures and Mechanisms 82.7 14.9 97.6 18%

Estimated Spacecraft Dry Mass 681 116 796.3 17%

Estimated Spacecraft Wet Mass 1301 116 1416.8

System LeveL Growth Calculations Total Growth

Dry Mass Desired System Level Growth 681 204 884.9 30%

Additional Growth (carried at system level) 89 13%

Total Wet Mass with Growth 1301 204 1505.4

Available Launch Performance to C3 (kg) 1600.0

Launch margin available (kg) 94.6

Estimated Spacecraft Inert Mass (for traj.) 792 204 996.0

Sample Return Craft Total Mass CBE Mass (kg) Growth (kg) Total Mass (kg)

Aggregate 

Growth (%)

Estimated Sample Return Craft Mass 93.3 16.9 110.2 18%

Total with System Level Growth 93 28 121.3 30%

Number of Sample Return Craft 2

Total Mass per Sample Return Craft (empty) 60.6 kg

Total Mass, Sample Returned 1 kg

Total Mass, Sample Return Capsule (Full) 61.6 kg
NASA/TM—2009-215825 63 

8.5 Case 3a 

Case 3 went to the same two asteroid targets, but this time only brought along minimal science package 

onboard and only returned one SRC to the Earth. The NEXT thruster system was used for electric 

propulsion. Note that in this table, the word “Moon” appears. This table was built off of the Mars Moon 

Sampler mission done previously by COMPASS and the words were not changed in time to make it into 

this report. 

Table 8.11—Case 3a—System Summary 

Table 8.12—Case 3a—Sample Return Craft Summary 

GLIDE 

container: Mars_Moons_Sampler: NEA_sampler_case3a

Spacecraft Master Equipment List Rack-up (Mass)

COMPASS S/C 

Design 

WBS Main Subsystems CBE Mass (kg)

Growth 

(kg) Total Mass (kg)

Aggregate 

Growth (%)

01 Moon Sampler Spacecraft 887.8 106.0 993.8

06.1 Science Payload 25.2 4.1 29.3 16%

06.2 Moon Sampler Lander 810.6 92.5 903.1

06.2.1 Attitude Determination and Control 44.3 18.7 53.6 42%

06.2.2 Command and Data Handling 29.3 8.0 37.3 27%

06.2.3 Communications and Tracking 24.2 6.5 30.7 27%

06.2.4 Electric Power 115.5 18.2 133.7 16%

06.2.5 Thermal Control 40.1 7.2 47.3 18%

06.2.6 Propulsion 137.5 16.8 154.3 12%

06.2.7 Propellant 272.8

06.2.8 Structures and Mechanisms 146.7 26.4 173.2 18%

06.3 Sample Return Craft (total, empty) 52.0 9.5 61.4 18%

06.3.1 Electrical Power Subsystem 6.0 1.2 7.2 20%

06.3.2 Thermal Control (Non-Propellant) 4.6 0.8 5.4 18%

06.3.3 Structures and Mechanisms 41.4 7.4 48.8 18%

Estimated Spacecraft Dry Mass 615 106 721.0 17%

Estimated Spacecraft Wet Mass 888 106 993.8

System LeveL Growth Calculations Total Growth

Dry Mass Desired System Level Growth 615 184 799.5 30%

Additional Growth (carried at system level) 78 13%

Total Wet Mass with Growth 888 184 1072.3

Available Launch Performance to C3 (kg) 1200.0

Launch margin available (kg) 127.7

Estimated Spacecraft Inert Mass (for traj.) 672 184 856.0

Sample Return Craft Total Mass CBE Mass (kg) Growth (kg) Total Mass (kg)

Aggregate 

Growth (%)

Estimated Sample Return Craft Mass 52.0 9.5 61.4 18%

Total with System Level Growth 52 16 67.5 30%

Number of Sample Return Craft 1

Total Mass per Sample Return Craft (empty) 67.5 kg

Total Mass, Sample Returned 1 kg

Total Mass, Sample Return Capsule (Full) 68.5 kg
NASA/TM—2009-215825 64 

9.0 REFERENCES 

System References 

1. ANSI/AIAA R-020A-1999, Recommended Practice for Mass Properties Control for Satellites, Missiles, and Launch Vehicles 

2. Space Mission Analysis and Design, 3rd edition (Space Technology Library) (Space Technology Library) , Wiley J. Larson 

(Editor), James R. Wertz (Editor), October 1999. 

Mission References 

3. Sauer, C., “A Users Guide to VARITOP: A General Purpose Low-Thrust Trajectory Optimization Program,” Advanced 

Projects Group, Jet Propulsion Laboratory, November 1991. 

4. The Boeing Company, “Delta IV Payload Planner’s Guide,” MDC 00H0043, October 2000. 

5. The Boeing Company, “Delta IV Payload Planner’s Guide Update—April 2002,” MDC 00H0043, April 2002. 

6. International Launch Systems (ILS), “Atlas Launch System Planner’s Guide,” Rev. 9, September 2001. 

7. Morton Thiokol, Inc., “Space Rocket Motors,” June 1987. 

8. COMPASS Final Report: Mars Moons Sampler, April 2008, CD-2008-23. 

9. (175706) 1996 FG3, http://www.johnstonsarchive.net/astro/astmoons/am-175706.html 

10. Mottola, Stefano; Lahulla, Felix, Mutual Eclipse Events in Asteroidal Binary System 1996 FG3: Observations and a Numerical 

Model, Icarus, Volume 146, Issue 2, pp. 556–567 (2000). 

Communications & C&DH 

11. Study Overview of the Near Earth Asteroid Sample Return, An ESA Technology Reference Study, European Space Agency, 

May 2007. 

12. NASA Preferred Reliability Practices, Practice No. PD–Ed–1203. 

Propulsion References 

13. ATK/PSI Propellant Storage Tank Specifications, www.psi-pci.com 

14. Arde, Inc. Propellant Storage Tank Specifications, www.ardeinc.com 

15. Star Motor reference pages from ATK. (http://www.atk.com/starmotors/starmotors_star48v.asp) 

16. Aerojet engines and propellant systems (http://www.aerojet.com/capabilities/S/C.php) 

17. BPT–4000 hall effect thruster. http://www.astronautix.com/engines/bpt4000.htm 

Thermal References 

18. Gilmore, David G. Editor, S/C Thermal Control Handbook: Volume 1 Fundamental Technologies, AIAA, 2002. 

19. Hyder, A.J., Wiley, R.L., Halpert, G., Flood, D.J. And Sabripour, S. S/C Power Technologies, Imperial College Press, 2000. 

20. Larson, W.J. And Wertz, J.R., Space Mission Analysis and Design, Third Edition, Space Technology Library, Microcosm 

Press, 1999. 

21. Bejan, A. and Kraus, A.D., Heat Transfer Handbook, John Wiley & Sons, 2003. 

22. Incopera, F.P. And DeWitt, D.P., Fundamentals of Heat and Mass Transfer, John Wiley and Sons, 1990. 

23. PICA definitions: http://nextbigfuture.com/2009/07/pica-and-pica-x-heat-shield.html 
NASA/TM—2009-215825 65 

APPENDIX A—ACRONYMS AND ABBREVIATIONS 

ACS Attitude Control System 

AO Announcement of Opportunity 

APXS Alpha Particle X-Ray 

Spectrometer 

ARU Array Regulator Unit 

AU Astronomical Unit 

BAE British Aerospace 

BCCU Battery Charge Control Unit 

BOL beginning of life 

C&DH Command and Data Handing 

C3 Launch energy per unit mass 

CBE current best estimate 

CCB Common Core Booster 

CEV Crew Exploration Vehicle 

CG Center of Gravity 

Comm Communications 

COMPASS COllaborative Modeling and 

Parametric Assessment of Space 

Systems 

COPV Composite Overwrapped Pressure 

Vessel 

COTS commercial off-the-shelf 

DCIU Digital Control Interface Unit 

DMR Design for Minimum Risk 

DOF Degree of Freedom 

DSN Deep Space Network 

DTE direct to Earth 

EELV Evolved Expendable Launch 

Vehicle 

ELV Expendable Launch Vehicle 

EM Electro Magnetic 

EP Electric Propulsion 

FEA finite element analysis 

FOM figure of merit 

GSFC NASA Goddard Space Flight 

Center 

GLIDE GLobal Integrated Design 

Environment 

GN&C Guidance, Navigation and Control 

GRC NASA Glenn Research Center 

GSE Ground Support Equipment 

HiVAC High-Voltage Hall Accelerator 

HPA High Pressure Assembly 

HQ NASA Headquarters 

IACO Integration, Assembly and 

Checkout 

IMDC I M Design Center 

IMU Inertial Measuring Unit 

IP Internet protocol 

JPL NASA Jet Propulsion Laboratory 

KSC NASA Kennedy Space Center 

LIDAR Light Detection and Ranging 

LPA Low Pressure Assembly 

LSP Launch Service Program 

LSTO Launch Service Task Order 

MEL Master Equipment List 

MEOP Manufacturers Expected Operating 

Pressure 

MLI multilayer insulation 

MGA Mass Growth Allowance 

MPU Makeup Power Unit 

NASA National Aeronautics and Space 

Administration 

Nav navigation 

NEA Near Earth Asteroid 

NEARER Near Earth Asteroids Rendezvous 

and sample Earth Returns 

NEXT NASA Evolutionary Xe Thruster 

NLS NASA Launch Services 

OTS off-the-shelf 

PEL Power Equipment List 

PICA Phenolic Impregnated Carbon 

Ablator 
NASA/TM—2009-215825 66 

PM Program Management 

PMAD Power Management and 

Distribution 

PMS Propellant Management System 

PM preliminary model 

PPM Propellant and Propellant 

Management 

PPU power processing unit 

RCS Reaction Control System 

RHS radioisotope heater unit 

S/C spacecraft 

SADA SA Drive Assembly 

SE&I Systems Engineering and 

Integration 

SEP Solar Electric Propulsion 

SLOC single line of code 

SN signal-to-noise 

SPACE System Power Analysis for 

Capability Evaluation 

SPU solar power unit 

SRC sample return capsule 

STO System Test Operations 

TDRSS Tracking and Data Relay Satellite 

System 

TPS Thermal Protection System 

TRL Technology Readiness Level 

TWTA Traveling Wave Tube Amplifier 

VACCO Vacuum and Air Components 

Company of America 
NASA/TM—2009-215825 67 

APPENDIX B—RENDERED DESIGN DRAWINGS 

Figure B.1—Rendered Baseline Case 1—Deployed View 

Figure B.2—Rendered Baseline Case 1—Close View of Main Body 
NASA/TM—2009-215825 68 

Figure B.3—Transparent Baseline Case 1—Main Bus View 1 

Figure B.4—Transparent Baseline Case 1—Main Bus View 2 

 
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01-09-2009 

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Technical Memorandum 

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4. TITLE AND SUBTITLE

COMPASS Final Report: Near Earth Asteroids Rendezvous and Sample Earth Returns 

(NEARER) 

5a. CONTRACT NUMBER

5b. GRANT NUMBER

5c. PROGRAM ELEMENT NUMBER

6. AUTHOR(S)

Oleson, Steven, R.; McGuire, Melissa, L. 

5d. PROJECT NUMBER

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5f. WORK UNIT NUMBER

WBS 346620.02.01.01.03.02 

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National Aeronautics and Space Administration 

John H. Glenn Research Center at Lewis Field 

Cleveland, Ohio 44135-3191 

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E-17087 

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NASA 

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Subject Categories: 12, 18, 20, 88, and 91 

Available electronically at http://gltrs.grc.nasa.gov 

This publication is available from the NASA Center for AeroSpace Information, 443-757-5802

13. SUPPLEMENTARY NOTES

14. ABSTRACT

In this study, the Collaborative Modeling for Parametric Assessment of Space Systems (COMPASS) team completed a design for a multiasteroid (Nereus and 1996 FG3) sample return capable spacecraft for the NASA In-Space Propulsion Office. The objective of the study was 

to support technology development and assess the relative benefits of different electric propulsion systems on asteroid sample return design. 

The design uses a single, heritage Orion solar array (SA) (~6.5 kW at 1 AU) to power a single NASA Evolutionary Xenon Thruster 

((NEXT) a spare NEXT is carried) to propel a lander to two near Earth asteroids. After landing and gathering science samples, the Solar 

Electric Propulsion (SEP) vehicle spirals back to Earth where it drops off the first sample’s return capsule and performs an Earth flyby to 

assist the craft in rendezvousing with a second asteroid, which is then sampled. The second sample is returned in a similar fashion. The 

vehicle, dubbed Near Earth Asteroids Rendezvous and Sample Earth Returns (NEARER), easily fits in an Atlas 401 launcher and its cost 

estimates put the mission in the New Frontier’s (NF's) class mission.

15. SUBJECT TERMS

Electric propulsion; Sample return; Mars exploration 

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80 

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