Steven R. Oleson and Melissa L. McGuire

Glenn Research Center, Cleveland, Ohio

COMPASS Final Report: Radioisotope Electric 

Propulsion (REP) Centaur Orbiter 

New Frontiers Mission

NASA/TM—2011-216971

February 2011

CD–2007–16
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Steven R. Oleson and Melissa L. McGuire

Glenn Research Center, Cleveland, Ohio

COMPASS Final Report: Radioisotope Electric 

Propulsion (REP) Centaur Orbiter 

New Frontiers Mission

NASA/TM—2011-216971

February 2011

CD–2007–16

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135
Acknowledgments

The COMPASS team wishes to thank the support of the Engineering Directorate, and specifi cally branch Chief Maria Babula and 

mission analysts of the Mission Design and Analysis Branch for their unending support and high caliber of engineer support. The

team would also like to thank the In Space Propulsion Offi ce for always providing challenging design study ideas.

Available from

NASA Center for Aerospace Information

7115 Standard Drive

Hanover, MD 21076–1320

National Technical Information Service

5301 Shawnee Road

Alexandria, VA 22312

Available electronically at http://www.sti.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.
NASA/TM—2011-216971 iii

Contents

1.0 Executive Summary .............................................................................................................................1

2.0 Study Background and Assumptions ...................................................................................................3

2.1 Introduction.................................................................................................................................3

2.2 Purpose........................................................................................................................................4

2.3 Assumptions and Approach ........................................................................................................4

2.3.1 NASA NF Class Mission Definition............................................................................. 4

2.3.2 Past Design Starting Points........................................................................................... 4

2.4 Growth, Contingency and Margin Policy ...................................................................................6

2.5 Mission Description....................................................................................................................7

2.5.1 Trajectory Baseline—Thereus ...................................................................................... 9

2.5.2 Mission Analysis Analytic Methods........................................................................... 11

2.5.3 Trojan Flyby Analysis................................................................................................. 12

2.5.4 Mission Analysis Event Timeline ............................................................................... 12

2.5.5 Mission Trajectory Details: Six ASRG Case .............................................................. 13

2.6 Launch Vehicle Details.............................................................................................................13

2.7 Science and Science Instruments Overview .............................................................................14

2.7.1 Science CONOPS ....................................................................................................... 16

2.8 System Design Trade Space......................................................................................................17

2.9 Baseline System Design............................................................................................................17

2.10 Internal COMPASS Details......................................................................................................19

2.10.1 GLIDE Study Share .................................................................................................... 19

2.10.2 GLIDE Study Container (Architecture)...................................................................... 19

2.10.3 GLIDE Study Container(s) ......................................................................................... 19

3.0 Baseline Design (Six ASRG) .............................................................................................................19

3.1 Top Level Design (MEL and PEL) for Six ASRG Case ..........................................................19

3.1.1 MEL for Six ASRG Case............................................................................................ 19

3.1.2 Power Equipment List (PEL) for Six ASRG Case...................................................... 20

3.2 System Level Summary ............................................................................................................20

3.3 Design Concept Drawing and Description................................................................................22

4.0 Subsystem Breakdown .......................................................................................................................22

4.1 Attitude Control System (ACS)................................................................................................22

4.1.1 ACS Trades................................................................................................................. 23

4.1.2 ACS Analytical Methods............................................................................................ 23

4.1.3 ACS Recommendation................................................................................................ 24

4.2 Communications.......................................................................................................................24

4.2.1 Communications Requirements.................................................................................. 24

4.2.2 Communications Assumptions.................................................................................... 24

4.2.3 Communications Design and MEL............................................................................. 24

4.2.4 Communications Trades.............................................................................................. 25

4.2.5 Communications Analytical Methods......................................................................... 25

4.2.6 Communications Risk Inputs...................................................................................... 25

4.2.7 Communications Recommendation ............................................................................ 25

4.3 Command and Data Handling (C&DH)—Avionics.................................................................25

4.3.1 Avionics Requirements............................................................................................... 25

4.3.2 Avionics Assumptions ................................................................................................ 26

4.3.3 Avionics Design and MEL.......................................................................................... 26

4.3.4 Avionics Trades .......................................................................................................... 27

4.3.5 Avionics Analytical Methods...................................................................................... 27

4.3.6 Avionics Risk Inputs................................................................................................... 27
NASA/TM—2011-216971 iv

4.3.7 Avionics Concerns, Comments, Recommendations................................................... 27

4.4 Electrical Power System ...........................................................................................................27

4.4.1 Power Requirements................................................................................................... 27

4.4.2 Power Assumptions..................................................................................................... 28

4.4.3 Power Design and MEL.............................................................................................. 28

4.4.4 Power Trades............................................................................................................... 30

4.4.5 Power Analytical Methods.......................................................................................... 30

4.4.6 Power Recommendations for Future Analysis............................................................ 30

4.5 Structures and Mechanisms ......................................................................................................31

4.5.1 Structures and Mechanisms Requirements ................................................................. 31

4.5.2 Structures and Mechanisms Assumptions................................................................... 31

4.5.3 Structures and Mechanisms Design and MEL............................................................ 32

4.5.4 Structures and Mechanisms Trades............................................................................. 32

4.5.5 Structures and Mechanisms Analytical Methods........................................................ 33

4.5.6 Structures and Mechanisms Risk Inputs ..................................................................... 33

4.5.7 Structures and Mechanisms Recommendation ........................................................... 33

4.6 Propulsion and Propellant Management ...................................................................................33

4.6.1 Propulsion and Propellant Management Requirements.............................................. 33

4.6.2 Propulsion and Propellant Management Assumptions................................................ 33

4.6.3 Propulsion and Propellant Management Design Trades............................................. 34

4.6.4 Propulsion and Propellant Management Design and MEL......................................... 35

4.6.5 Propulsion and Propellant Management Risk Inputs.................................................. 38

4.6.6 Propulsion and Propellant Management Recommendation ........................................ 39

4.7 Thermal Control.........................................................................................................................39

4.7.1 Thermal Requirements ................................................................................................. 39

4.7.2 Thermal Assumptions................................................................................................... 39

4.7.3 Thermal Design and MEL ............................................................................................ 40

4.7.4 Thermal Trades........................................................................................................... 41

4.7.5 Thermal Analytical Methods....................................................................................... 41

4.7.6 Thermal Risk Inputs.................................................................................................... 43

5.0 Cost, Risk and Reliability...................................................................................................................43

5.1 Costing: Six ASRG Configuration............................................................................................43

5.2 Reliability..................................................................................................................................45

5.2.1 Reliability Methodology ............................................................................................. 45

5.2.2 Reliability Assumptions.............................................................................................. 46

5.2.3 Reliability Results....................................................................................................... 46

5.2.4 Interpretation of the Reliability Results...................................................................... 48

6.0 Trades.................................................................................................................................................48

6.1 Case 1—REP Centaur Orbiter With Six ASRGs—Baseline ....................................................48

6.2 Case 2—REP Centaur Orbiter With Eight ASRGs ..................................................................48

6.2.1 Eight ASRG REP Centaur Orbiter Summary ............................................................. 49

6.2.2 Eight ASRG System Summary Masses (MEL and System Sheet)............................. 52

7.0 Challenges, Lessons Learned, and Areas For Future Study...............................................................53

7.1 Six ASRG REP Centaur Orbiter Summary...............................................................................53

7.1.1 Potential Mass Reductions.......................................................................................... 53

7.1.2 Areas for Future Work ................................................................................................ 53

Appendix A.—Acronyms and Abbreviations.............................................................................................55

Appendix B.—Six ASRG Design Rendered Drawings..............................................................................57

Appendix C.—Eight ASRG Design Rendered Drawings...........................................................................59

Appendix D.—Mission Analysis of Centaur Body Bienor.........................................................................63

D.1 Initial Potential Centaur Target Trajectory Analysis................................................................63
NASA/TM—2011-216971 v

Appendix E.—Study Participants...............................................................................................................67

Bibliography ...............................................................................................................................................68

NASA/TM—2011-216971 1

COMPASS Final Report: Radioisotope Electric Propulsion (REP)

Centaur Orbiter New Frontiers Mission

Steven R. Oleson and Melissa L. McGuire

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

1.0 Executive Summary

Radioisotope Electric Propulsion (REP) has been shown in past studies to enable missions to outer 

planetary bodies including the orbiting of Centaur asteroids (Figure 1.1). Key to the feasibility for REP 

missions are long life, low power electric propulsion (EP) devices, low mass Radioisotope Power System 

(RPS) and light spacecraft (S/C) components. In order to determine the key parameters for EP devices to 

perform these REP missions a design study was completed to design an REP S/C to orbit a Centaur in a 

New Frontiers (NF) cost cap. The design shows that an orbiter using several long lived (~200 kg xenon 

(Xe) throughput), low power (~700 W) Hall thrusters teamed with six (150 W each) Advanced Stirling 

Radioisotope Generators (ASRG) can deliver 60 kg of science instruments to a Centaur in 10 yr within 

the NF cost cap. Optimal specific impulses (Isp) for the Hall thrusters were found to be around 2000 s with 

thruster efficiencies over 40 percent. Not only can the REP S/C enable orbiting a Centaur (when 

compared to an all chemical mission only capable of flybys) but the additional power from the REP 

system can be used to enhance science and simplify communications.

Figure 1.1.—Mission overview.
NASA/TM—2011-216971 2

Figure 1.2.—Conceptual REP Science S/C with Star 48 Rocket.

The mission design detailed in this report is a Radioisotope Power System (RPS) powered EP science 

orbiter to the Centaur Thereus with arrival 10 yr after launch, ending in a 1 yr science mapping mission. 

Along the trajectory, approximately 1.5 yr into the mission, the REP S/C does a flyby of the Trojan 

asteroid Tlepolemus. The total ΔV of the trajectory is 8.9 km/s. The REP S/C is delivered to orbit on an 

Atlas 551 class launch vehicle with a Star 48 B solid rocket stage. Figure 1.2 shows the conceptual S/C

reported on in this document.

Table 1.1 summarizes the REP S/C and mission. It is evident that the largest systems are power and 

propulsion, which enable the timely mission to orbit a Centaur. While the power is more abundant than 

usual for S/C of this class, it is not large enough to allow for the high Isps (>3000 s) normally associated 

with solar electric propulsion (SEP) missions to Mars or nearer the Sun. Thus propellant loading is 

significant. The additional power used for transit is available for science and communications at the 

Centaur. Preliminary assessments point towards the use of higher power communications to reduce Deep 

Space Network (DSN) time and reduce operations costs.
NASA/TM—2011-216971 3

TABLE 1.1—MISSION AND S/C SUMMARY

Mission 10 yr mission, orbit Centaur (Thereus) in 10 yr, flyby Trojan (Tlepolemus) in 1.5 yr, 

8.9 km/s

Total mass 

with growth

Launch Atlas 551/Star 48, C3 97.28 km2/s2 1260 kg 

(wet)

Science Six instruments consuming 60 W, remote imaging (wide angle camera (WAC) and 

narrow angle camera (NAC) cameras, Laser Detection and Ranging (LIDAR), NearInfrared Mapping Spectrometer (NIMS) and in-situ sampling (Neutral Gas and Ion Mass 

Spectrometer (NGIMS), DFM, 130 Mb per day

57 kg

Power Six ASRG with multilayer insulation (MLI), attached (loaded) in pairs, 900 W end of life

(EOL) (10-yr)

199 kg

Propulsion Primary electric: 3+1 Long Life Hall Thrusters, operated serially, 500 kg Xe propellant 

Load, 650 W into thruster, 1920 s Isp, 170 kg Xe throughput each, single string power 

processing units (PPU) (95%), thruster feed, thruster 30,000 hr

Secondary chemical: Blow-down hydrazine, 80 kg propellant, eight 0.25 lbf monoprop 

thrusters

137 kg

Mechanical Hexagonal aluminum-lithium (Al-Li) bus with propulsion and science decks, capable of 

carrying 6 g launch loads

118 kg

Communications 220 W, 8 kb/s data rate, Ka Band, 2.1 m antenna 52 kg

Command and Data 

Handling (C&DH) 

RAD 750 computer with x MB storage capacity 45 kg

Attitude, 

Determination and

Control (AD&C)

Two Star cameras, inertial measurement unit (IMU), four reaction wheels, hydrazine 

propulsion

22 kg

Thermal MLI and heaters, ASRG isolated 49 kg

2.0 Study Background and Assumptions

2.1 Introduction

The COllaborative Modeling and Parametric Assessment of Space Systems (COMPASS) team was 

approached by the NASA Glenn Research Center (GRC) In-Space Project to perform a design session to 

develop REP S/C Conceptual Designs (with Cost/Risk/Reliability) for missions of three different classes: 

NF Class Centaur Orbiter (with Trojan Flyby), Flagship, and Discovery Class. The designs will allow 

trading of current and future propulsion systems. The results will directly support technology 

development decisions. The mission class documented in this final report is the NF class orbiter mission 

to a Centaur body, with a Trojan flyby along the trajectory.

Past studies have shown that REP can enable orbiters for outer planetary small bodies. The mission 

design detailed in this report is an REP powered EP science orbiter to the Centaur Thereus with arrival 

10 yr after launch, with a 1 yr science mapping mission, to make a total lifetime requirement for all 

systems of 11 yr. Along the trajectory, approximately 1.5 yr into the mission, the REP S/C does a flyby of 

the Trojan asteroid Tlepolemus. The total ΔV of the trajectory is 8.4 km/s. The REP S/C is delivered to 

orbit on an Atlas 551 class launch vehicle with a Star 48 B solid rocket stage. Figure 1.2 shows the 

conceptual S/C reported on in this document.

The Executive Summary (Section 1.0) provides a framework for where this study fits into NASA’s 

goals. The rest of the document describes the various aspects of the mission including: mission category, 

propulsion type and engines, mission duration, scientific payload and its requirements, power requirements, 

communications requirements, launch vehicle and a specific mission target (Thereus). As with any study, 

there are many options among these mission elements. Many of them are traded for comparison in order to 

provide optimal scientific results, minimize cost and provide the highest probability of mission success (as 

defined by established goals). There is a preliminary discussion on past missions that serves as a set of 

baseline designs. The science mission objective is to determine the origin and evolution of Centaurs and 

Trojan bodies. For example: What is the density, volume, rotation state, albedo of the “typical” Centaur? 
NASA/TM—2011-216971 4

Are Centaurs the direct descendants of Kuiper Belt Objects (KBOs)? How do Centaurs compare to the icy 

small satellites of Saturn (Phoebe, Hyperion) that are supposedly KBOs? Were Trojan asteroids formed in 

place, as part of the outer main belt of asteroids, or were they scattered into the Kuiper Belt as the planets 

migrated outward about 800 M years after the formation of the Earth?

2.2 Purpose

The goal of the REP science S/C is to send a science payload to a Centaur class body. The Centaurs 

are planetary bodies in orbit about the Sun and located between the orbits of Jupiter and Neptune. These 

bodies are typically asteroidal to comet-like in appearance and physical makeup. If possible, the mission 

flight plan will include a flyby of one of the Jupiter Trojan asteroids. The Trojan asteroids occupy one of 

the Lagrange points created by Jupiter and the Sun.

This study will utilize the COMPASS S/C Conceptual Design team to provide complete Science 

Class Reference Mission Designs.

Prework that was completed before the official study kickoff

• Determine Science Class missions that are compelling and relevant: Applied Physics Laboratory 

(APL)

• Setup Mission analyses capabilities for compelling missions: GRC

• Develop a suite of EP systems to trade on S/C designs: GRC 

• Develop Concept of Operations (CONOPS) and suite of science instruments and requirements for 

each science mission class: APL 

2.3 Assumptions and Approach

The following section contains the mission description of the class of mission being designed in this 

REP study. Design details are the most current available at the time of this design session. This study is 

focusing on a NF Class Mission design.

2.3.1 NASA NF Class Mission Definition

NASA has named New Horizons (NH) as the first NF class mission of the solar system. From the 

NASA NF Website, the definition of a NF mission at the time of this study was as follows

“The Office of Space Science (OSS) of the National Aeronautics and Space Administration (NASA) 

announces an opportunity to propose for two different types of scientific investigations through the New 

Frontiers Program: New Frontiers Mission investigations specifically for the planet Venus, the Earth’s 

moon, Jupiter, or a comet nucleus (see Section 2.1); and Mission of Opportunity investigations.”

(Extract: AO-03-OSS-03 Introduction)

New Frontiers Mission investigations are to be completed through space flight missions launched on 

Expendable Launch Vehicles (ELVs) no later than June 30, 2010, in order to accomplish science 

objectives in compliance with those stated in Section 2.1 of this Announcement of Opportunity (AO) for 

a total NASA OSS Cost through mission completion not to exceed $700M in Fiscal Year (FY) 2003 

dollars (see Sections 5.1 through 5.11 for a description of both general and specific restrictions).

New Frontiers Mission of Opportunity (MO) investigations are part of non-OSS space missions of 

any size that will be launched no later than December 31, 2008, that require a commitment from 

NASA before December 31, 2005

2.3.2 Past Design Starting Points

The REP Centaur Science S/C and mission used the APL NH S/C as its starting point for the baseline 

design. The APL NH S/C is the first NF class vehicle (NH website). The NH mission is a 10 yr mission to 
NASA/TM—2011-216971 5

map the Kuiper belt region beyond Neptune, and as such, has both a mission and trajectory similar to the 

objectives of the Centaur body orbiter to be designed in this COMPASS session.

The initial design of the REP S/C was based on the specifications of the NH S/C. For reference, the 

NH mission and S/C Design Highlights are as follows

• Launch date/time: January 19, 2006, at 19:00 UTC on an Atlas V 551 using a Star 48B solid third 

stage

• On-orbit dry mass: 385 kg 

• Nominal power output: 228 W

• Life: +15 yr

• Propulsion: Blow down hydrazine: 80 kg of fuel

• Stabilization: Spin and three-axis depending on mission phase

• Thermal: ‘Thermal Bottle’ using heaters/louvers powered by shunts from ASRG

• Science: 32.5 kg/69.7 W total (25.9 W max mode)

• Utilized hibernation mode

The NH S/C supports seven science instruments and is powered by an ASRG. Figure 2.1 is a graphic 

of the NH science S/C portion without the Star 48B attached. Figure 2.2 shows the NH S/C being 

assembled in a clean room.

Figure 2.1.—The NH S/C.
NASA/TM—2011-216971 6

Figure 2.2.—NH S/C clean room assembly.

2.4 Growth, Contingency and Margin Policy

Mass Growth: For dry mass elements in the system design, the COMPASS team uses the American 

Institute Aeronautics and Astronautics (AIAA) R-020A-1999, “Recommended Practice for Mass 

Properties Control for Satellites, Missiles, and Launch Vehicles.”

Table 2.1 shows the percent mass growth separated into a matrix specified by level of design maturity 

and specific subsystem. In this document mass growth and weight growth are used interchangeably.

Weight growth allowance (WGA) and mass growth allowance (MGA) mean the same thing when 

expressed as a percent value.

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

vehicle level is calculated. The COMPASS team desired total growth to be 30 percent, and an additional 

growth is carried at the system level in order to achieve a total system growth of a 30 percent limit on the 

dry mass of the system. Note that for designs requiring propellant, growth in propellant is either book 

kept in the propellant itself or in the ∆V used to calculate the propellant necessary to fly a mission.

The COMPASS team uses the Discovery Announcement of Opportunity (AO) definitions of 

contingency and mass margin.

From the Discovery AO: Definitions of Contingency and Mass

Contingency (or Reserve), when added to a resource, results in the maximum expected value for that 

resource. Percent contingency is the value of the contingency divided by the value of the resource less the 

contingency.

Margin is the difference between the maximum possible value of a resource (the physical limit or the 

agree-to limit) and the maximum expected value for a resource. Percent margin for a resource is the 

available margin divided by its maximum expected value.
NASA/TM—2011-216971 7

Power growth: The COMPASS team uses a 30 percent margin on the bottoms-up power requirements 

in modeling the power system. See Section 4.4 for the power system assumptions.

Science payload mass and power margins assumptions are: The science payload package was 

designed by the engineering staff at APL. They recommend the following margins be used on their 

science package. Therefore, the science package uses a different MGA and power margin than the rest of 

the S/C design.

• 30 percent margin should be added for mass estimates (mass estimates do not include any 

additional shield that may or may not be needed with the REP mission)

• 30 percent margin should be added for power estimates

TABLE 2.1.—PERCENT MGA 

Code

Design maturity

(basis for mass 

determination)

Percent MGA

Electrical/electronic 

components

StructureThermal controlPropulsionBatteriesWire HarnessesMechanismsInstrumentation

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

Prerelease 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

2.5 Mission Description

The primary requirement for the mission is orbit a Centaur class body in the minimum amount of time 

while delivering the maximum payload. Along the trajectory, if possible, a flyby of a Trojan asteroid will 

be incorporated into the mission profile. The representative Centaur body chosen for the mission was 

Thereus. Thereus, an average sized Centaur in the Saturn-Uranus region, was found to be representative 

of several Centaur targets in terms of its known behavior and required delivery ∆V. 

Figure 2.3 shows the grouping of Centaur objects in the solar system region around Saturn, Uranus, 

and Neptune. The orbits of the known Centaurs and Neptune Trojans are show in Green. The inclination 

of the orbit is the Y-axis. Using the mapping of Centaur bodies in Figure 2.3, and the low thrust mission 

analysis code VARITOP, a suite of trajectories to Centaur bodies was flown to find a mission that would 

fit the science requirements and deliver the inert mass from the bottoms-up S/C system design.
NASA/TM—2011-216971 8

Figure 2.3.—Centaur body distribution.

Figure 2.4.—Performance comparison.

While the initial Centaur body selection is done over a wide range of bodies, the iterative process of 

mission and system design done during the design session limits the selection of Centaur bodies to those 

with trajectory C3 is similar to the performance C3 of the ELV (Atlas 551/Star 48B) to accommodate the 

inert mass of the S/C design.

Next, Figure 2.4 shows the Centaur body target comparison in terms of mission performance 

parameters of launch mass, necessary propellant mass and delivered payload mass to the target, as well as 

total mission ΔV. The two red-boxed areas are the potential target Centaurs that fit within the initial 

assumed performance of the REP S/C and thrusters.
NASA/TM—2011-216971 9

Figure 2.5.—Distribution of Trojan Asteroids in Jupiter’s orbit.

Figure 2.5 shows the distribution of the Trojan asteroids in the orbit of Jupiter. The clusters of green 

objects located behind Jupiter in its orbit are the Trojans, and the clusters in front of Jupiter in its orbit are 

the Greeks. Both areas in the Jovian orbit have been added into the trade space of target objects for close 

flyby reconnaissance along the way to the Centaur object.

2.5.1 Trajectory Baseline—Thereus

Preliminary design of the REP Centaur orbiter used the Centaur body, Bienor, as a target. After the 

REP S/C bottoms-up analysis and sizing, the amount of payload delivered to the Centaur body Bienor 

(see Appendix D) was determined to be insufficient to accommodate the baseline S/C design and science 

instruments mass. Therefore, mission analysis was performed to locate another Centaur body trajectory 

that would place the REP S/C and science instruments at the Centaur before it reaches the Centaur body’s 

perihelion (closest approach to the Sun as required by the mission science drivers, while allowing enough 

delivered mass to accommodate the REP S/C. Figure 2.6 shows the net delivered mass as a function of 

trip time to the Centaur body Thereus. It can be observed that net mass delivered for trip times greater 

than 9 yr begins to level off.

The Centaur Thereus was then chosen to fit those parameters. There is no coast phase in the mission 

but a 10 percent thrust margin is included in the analysis. The baseline trajectory parameters for the 

Thereus trajectory were:

• Launch mass (Mo): 1260.2 kg (Atlas 551 performance to C3)

• C3: 97.28 km2/s2

• Launch date: November 15, 2024

• Fly-by date: August 7, 2026

• Arrival date: November 13, 2034

• Power: 650 W

• Isp: 2057 s

• ∆V: 8.93 km/s
NASA/TM—2011-216971 10

• Propellant (Mp): 450.8 kg (used to perform ΔV)

• Initial mass (Mo): 1260 kg (starting mass from ELV performance to C3)

• Mo – Mp: (1260 – 450) 809 kg

• Fly-by target: Trojan Asteroid Tlepolemus

This trajectory allowed for an REP inert mass of 809 kg. This inert mass includes the residuals and 

margin in the Xe propellant. The 450 kg is the ideal Xe propellant used to fly the ideal trajectory ΔV.

With the 30 percent margin and the extra Reaction Control System (RCS) propellant, which is not 

included in the ideal ∆V modeling via the low thrust trajectory, the REP S/C was able to fit inside this 

“inert” mass box. See Table 4.3 in Section 4.3.3 to see the system masses, with margin, and S/C adapter 

to come up with the launch margin for the REP S/C.

Figure 2.7 shows the launch window analysis on the baseline Thereus trajectory. A 10-day launch 

window will result in minimal impact to the trajectory. A 20-day launch window may require slightly 

longer trip time. Re-optimizing the trajectory to include coast periods could alleviate this issue.

Figure 2.6.—Net mass delivered versus transfer time.

Figure 2.7.—Launch window analysis.
NASA/TM—2011-216971 11

2.5.2 Mission Analysis Analytic Methods

The low thrust mission analysis performed for this mission was done using the low thrust 

optimization code VARITOP, developed at the NASA Jet Propulsion Laboratory (JPL). Typically a 

maximum duty cycle of 90 percent is imposed on the trajectory optimization using the updated program, 

SEPTOP. Because VARITOP does not have that feature, the thrust and therefore efficiency was reduced 

to 90 percent of the anticipated performance. This interjected 10 percent thrust margin artificially puts 

10 percent coasting times into the trajectory to alleviate the launch window analysis implications in 

Figure 2.7. 

The optimization parameters for running this trajectory were as follows in Table 2.2. Note that the 

mission duration quoted in the mission analysis assumptions is trip time to the Centaur Body. An 

additional year of analysis is performed at the Centaur making the total mission time 11 yr.

TABLE 2.2—MISSION ANALYSIS ASSUMPTIONS

Launch vehicle Atlas 551 with Star 48 Optimized

Launch vehicle contingency 10% Fixed

Epoch date April 22, 2024 Fixed

Launch date November 15, 2024 Optimized

Trajectory mission duration 10 yr Fixed

Thruster power 750 W Fixed

Thruster efficiency Hall Curve Fixed

Isp 1800 s Fixed

The thruster performance was modeled using the thruster curves shown in Figure 2.8. These are based 

on a single operating point thruster performance estimates provided by GRCs EP division. The Isp and 

sensitivity was initially optimized for the mission and then were fixed for parametric analyses on the 

chosen system. Trades are shown later in Section 6.0.

Comparison of provided and input performance

Figure 2.8.—Thruster performance modeling assumptions.
NASA/TM—2011-216971 12

Figure 2.9.—Potential Trojan flybys along trajectory.

Figure 2.10.—Mission timeline.

2.5.3 Trojan Flyby Analysis

Given the launch dates for the optimal Thereus trajectory, the possibilities of encountering a Trojan 

asteroid were greater for the Thereus mission than they were for the Bienor mission. Figure 2.9 shows the 

potential Trojan flybys as the REP science S/C passes through the asteroid belt. With thousands of Trojan 

asteroids to choose from, it can be seen that multiple Trojans were within flyby range without any 

targeting necessary by the S/C, and therefore, without extra propellant budgeting necessary.

Based on the available Trojan asteroids during the flyby, the Trojan asteroid, Tlepolemus, was chosen 

as the flyby target. Tlepolemos is the named Trojan nearest to the Earth in an orbit closer than any other 

of the Trojans to Earth’s orbit.

2.5.4 Mission Analysis Event Timeline

• Launch date: November 15, 2024

• Fly-by Trojan date: August 7, 2026

• Arrival at Centaur date: November 13, 2034

A benefit of this mission and trajectory is that First Science with a Trojan flyby is available less than 

2 yr after launch. This allows for testing of the science instruments for the final Centaur mission along 

with the capture of interesting science data of a Trojan asteroid. Figure 2.10 shows the relative mission 

timeline of the major events in the mission: Trojan flyby, Centaur body arrival and science mapping.
NASA/TM—2011-216971 13

Figure 2.11.—Trajectory from Earth to Thereus, with Trojan Asteroid flyby.

2.5.5 Mission Trajectory Details: Six ASRG Case

Figure 2.11 shows the trajectory to Thereus with a flyby of the Trojan asteroid Tlepolemus. The green 

orbit is the Trojan asteroid belt, and the aqua blue orbit is that of the Centaur body Thereus. The small

triangle on the orbit of Thereus indicates perihelion. The goal of the arrival at Thereus is to arrive at the 

Centaur body prior to its perihelion, in order to take science data at the Centaur as it goes through its 

closest approach to the sun and encounters its maximum temperatures and has the highest probability of 

activity. For reference, the Earth’s orbit about the Sun is shown in blue.

The CONOPS of the mission is based on that of the Deep Space One Mission that validated EP 

technology with the flight of the NSTAR engine. The Ground Trajectory Planning updates monthly. Orbit 

determination occurs weekly via an on-board camera every 4 to 6 hr. The “Autonav” system can 

overwrite, during retargeting, a real-time ephemeris, schedules and executes events and navigation update 

computations. The Reduced State Encounter Navigation (RSEN) operations mode begins several hours 

prior to encounters and maintains visual lock from the Earth to the REP S/C.

2.6 Launch Vehicle Details

The baseline launch vehicle is the Atlas 551 with a Star 48 solid propellant upper stage. The launch 

vehicle performance versus launch C3 is shown in Figure 2.12. The launch vehicle contingency was 

assumed to be 10 percent, and was generated using the low thrust trajectory code VARITOP. This data 

assumed that the Star 48 had already performed its burn. Consequently, the mass of the Star 48 will not be 

included in the REP S/C master equipment list (MEL).

Figure 2.13 shows the packaging of the REP S/C and Star 48 engine in the Atlas 551 short 5 m 

payload fairing. Note that the Atlas V payload shroud volume more than accommodates the REP Science 

orbiter S/C, leaving plenty of room for design changes such as a larger antenna diameter or an increase in 

the number of ASRGs. See Figure B.1 in Appendix B for the dynamic envelop of the REP S/C inside of 

the launch vehicle fairing.
NASA/TM—2011-216971 14

Atlas V551/Star 48 V

C3 (km2/s2)

Figure 2.12.—Launch mass versus C3.

Figure 2.13.—REP S/C in 

Atlas 551 payload shroud.

2.7 Science and Science Instruments Overview

Typically, the science payload delivered by the bus in a COMPASS design session is the figure of 

merit (FOM) of the analysis. In this design, the APL science team brought the science instruments to the 

COMPASS team for inclusion in the REP orbiter design. The goal, then, of the design was to fit the bus 

and science orbiter package inside of the performance (delivered mass) to the chosen C3 target by the 

launch vehicle. Anything left over was considered margin.
NASA/TM—2011-216971 15

The initial desired science payload from APL consisted of six core instruments base-lined for the 

Centaur REP mission study. The instruments in order of priority are: 

(1) Imager—Long Range Reconnaissance Imager (LORRI)

(2) Thermal mapping—Thermal Emission Imaging System (THEMIS)

(3) Ranging and topography—LIDAR

(4) Imager—WAC

(5) Spectrometer—NIMS

(6) Gas spectrometer—NGIMS

The instruments are not required to have a separate processor (i.e., dual processing unit (DPU)) and 

will interface directly with the Main Processor of the REP bus. Because Centaurs cross the orbits of the 

planets, their own orbits are unstable, evolving rapidly. This makes the Centaur mission a challenging 

one. However, in doing the design, the science instrument NGIMS was removed from the package to save 

on mass, cost and power.

The proposed mission includes a flyby of a Trojan asteroid with limited science to be performed during 

that flyby. Once at the Centaur body, up to a year of science mission operations will be performed. Therefore, 

there must be enough power and attitude control propellant to sustain the vehicle for that amount of time. 

Table 2.3 lists the major details (mass, power, cost, etc.) of the six science instruments initially 

evaluated by APL for inclusion on the REP Centaur conceptual S/C being designed in this session. Note 

that not all of the instruments were included in the final science package used in the design and reported 

in the science instrument MEL and therefore flown on the S/C in this design. In order to save on mass, the 

science package suite was reduced. New THEMIS was the first instrument dropped in order to save mass.

TABLE 2.3.—SCIENCE INSTRUMENTS FOR CENTAUR MISSION

Instrument name Mass

(kg)

Power

(W)

Dimensions

(cm)

Output 

data rate

Estimated 

data volume*

Estimated cost 

(FY07)

LORRI 8.32 2.11 62 by 48 by 35.5 12.6 Mbps TBD TBD

New THEMIS 12.2 14 29 by 38 by 55 614.4 Kbps TBD $12 million

WAC 0.6 4 14.5 by 9.2 by 7.6 TBD TBD $2 million

NIMS 18 12 W average

13 W peak

83 by 37 by 39 (optics)

20 by 25 by 13 (electronics)

11.52 kbps TBD ??

LIDAR 5 16.5 W average

20.7 peak

37.5 by 21.6 by 22.9 51 bps TBD $8 million

NGIMS 10.50 24 19 by 24.7 by 36.2 1.5 Kbps TBD $14 million

*A total of 130 Mbits/day will be baselined and allocated between the instruments

Table 2.4 lists the MEL for the science instruments as modeled in the COMPASS design session.

Note that the value in the quantity column of the MEL for NGIMS has been zeroed. As the design

progressed, NGIMS was also scaled back out of the science instrument suite in the interest of saving 

mass. New THEMIS was entered back into the Science instrument suite during the analysis.

TABLE 2.4.—REP SCIENCE S/C TOP LEVEL BOTTOMS-UP MEL
NASA/TM—2011-216971 16

WBS

no.

Description

REP Centaur Mission Six ASRG

(October 11, 2007)

Qty Unit 

mass

(kg)

CBE mass

(kg)

Growth

(%)

Growth

(kg)

Total 

mass

(kg)

01 REP S/C (Payload and Stage) - ----- 1071.00 11.68 125.14 1196.14 

01.1 Science Payload - ----- 44.12 30.00 13.24 57.36 

01.1.1 APL science instruments - ----- 25.52 30.00 7.66 33.18

01.1.1.a LORRI 1 8.32 8.32 30.00 2.5 10.82

01.1.1.b New THEMIS 1 12.2 12.20 30.00 3.66 15.86

01.1.1.c LIDAR 1 5.0 5.00 30.00 1.50 6.50

01.1.1.d NGIMS 0 10.5 0.00 30.00 0.00 0.00

01.1.2 Additional science instruments - ----- 18.60 30.00 5.58 24.18

01.1.2.a WACS 1 0.60 0.60 30.00 0.18 0.78

01.1.2.b NIMS 1 18.00 18.00 30.00 5.40 23.40

2.7.1 Science CONOPS

Orbital CONOPS

• The S/C will orbit the Centaur Body for at least 1 yr

• The S/C will orbit the body in the plane of Earth’s sky (i.e., the S/C will always be able to 

communicate with the Earth and will not eclipse behind the body)

• Before orbit insertion, upon approach to the body, the LORRI instrument will be required to 

capture at least two optical navigation images per day

• The minimum daily science collection will be 130 Mb/day while in Thereus orbit

– The instruments and recorder allow for a higher collection rate if desired

• Primary ground contact operations concept:

– Three 8 hr contacts per week using the DSN—70 m Dish, yielding 24 hr/wk of contact time 

and ~19.5 hr/wk of actual downlink time (at 6.5 hr/contact).

– Possible ~1 Gbits/day of science/housekeeping collection

 Assumed minimum 45 kbits/s downlink at 70 m dish with 100 W of radio frequency (RF)

radiated power 

– DSN total 1 yr orbital cost: FY07 $7.4M ($23M for 1 yr of 7 tracks per week)

Trojan Flyby CONOPS

• Science collection at the Trojan flyby consists of two data collection periods

– For 10 days leading up to closest approach to the asteroid, the LORRI instrument will capture 

at least six images per day

 Assume 50 percent duty cycle of the EP thrusters during this time

– The 3 days of closest approach (1 day before, 1 day during, 1 day after) will consist of all 

instruments collecting data

 Assume 0 percent duty cycle of the EP thrusters during this time

– It is estimated that ~ 8 Gbits of science data will be collected during the entire flyby

• Total non-thrusting period during asteroid flyby would be approximately 12 days broken down as 

follows

– 5 days of non-thrusting leading up to closest approach

– 3 days of non-thrusting during closest approach

– 52 hr of data downlink for 8 Gbits during non-thrusting periods

 Assume 70 m dish and S/C can downlink at minimum 45 kbits/s

– 2 days of contingency
NASA/TM—2011-216971 17

2.8 System Design Trade Space

The COMPASS design trade space will cover the following mission classes and subdesigns within 

those classes of mission as noted herein. This first document will only deal with the initial baseline NF

class mission, similar to the NH mission. A second design session will be run to study a Flagship mission 

(see COMPASS document CD–2007–18).

2.9 Baseline System Design

Figure 2.14 shows the basic design of the Centaur REP driven S/C. From the top down in this 

diagram, the REP S/C consists of a 2.1 m antenna dish located on the top for relay to the DSN. Below the 

dish sits the payload deck where the science instruments (shown in purple and grey) and avionics 

instruments are mounted. There are six ASRG units, providing at least 900 W of total S/C power, 

mounted by twos (dark blue grey) to the S/C bus superstructure, pointing out perpendicularly from the 

main body of the S/C. Also on the s/c are two Xe tanks (orange), a helium (He) pressurization tank, four 

Xe Hall EP Thrusters, and a Star 48 solid rocket engine sitting directly below the propellant management 

equipment.

The REP S/C, shown in Figure 2.15, will be launched on an Atlas 551 with Star 48 Upper Stage 

(similar to NH launch). The payload will be located in the middle of the REP/Star 48 motor stack just 

below the antenna dish.

Figure 2.14.—Baseline conceptual REP S/C design on top of Star 48 rocket motor.
NASA/TM—2011-216971 18

Figure 2.15.—Conceptual REP science S/C.

To first order, the S/C configuration is built around the following major components

• Bottom propulsion/power deck

– ~500 kg of Xe in two Composite Overwrapped Pressure Vessel (COPV) tanks

– Six Advanced RPS

• Stacked by twos

• Supported by ends with plate struts

– Four advanced Hall or Ion propulsion systems

• Thrusting mostly tangential

– Hydrazine secondary propulsion system

• Top payload/avionics deck

– 2.1 m dish antenna

– Four Science payloads—side pointing

– Avionics/Communications/GN&C

• Bottom launch vehicle interface

The main design challenge for an REP S/C is minimizing S/C masses while integrating a significant 

Xe propellant and multiple RPS. Fortunately, Xe propellant is very dense and can be stored in carbon 

fiber overwrapped tanks. The selected RPS was ASRG due to their high efficiency and subsequently 

small plutonium load for the power delivered. The best balance of power and thrust was found to be 

around 900 W EOL, with 700 W being fed to the electric thruster power processors. The remaining 

200 W was for housekeeping. This power requirement was provided by six ASRGs, each containing 

about 0.9 kg of plutonium for a total of 5.4 kg. By comparison the NH S/C, using an 80 percent loaded 

RTG had 11 kg of plutonium.

Besides the obvious advantage of enabling orbiting a Centaur, the REP approach also provides much 

more power to the science and communications system once the asteroid is reached when compared to a 

chemical flyby system. During this design it was found that the power was actually more valuable to the 

communication system. While additional power for science will require additional mass allowance, the 
NASA/TM—2011-216971 19

communication and data storage systems and operations can be reduced by allowing for higher power 

communications and either a smaller antenna or reduced DSN time.

2.10 Internal COMPASS Details

COMPASS is a multidisciplinary collaborative engineering team whose primary purpose is to 

perform integrated vehicle systems analysis and provide trades and designs for both Exploration and 

Space Science Missions.

2.10.1 GLIDE Study Share

GLIDE (GLobal Integrated Design Environment) is a data collaboration tool that enables secure 

transfer of data between a virtually unlimited number of sites from anywhere in the world. GLIDE is the 

primary tool used by the COMPASS design team to pass data real-time between subsystem leads. The 

study team members can find documents saved at the following permissions-limited, webdav-accessible, 

share.

Study Share: https://glide.grc.nasa.gov/REP_Sept2007

2.10.2 GLIDE Study Container (Architecture)

GLIDE Architecture: REP_Sept2007

2.10.3 GLIDE Study Container(s)

Six ASRG Case: Thereus_6ASRG

Eight ASRG Case: Bienor1

Study Description

NF Class Mission to the Centaur Bienor 1. Although the target has been shifted to Thereus, per the 

mission section, the study container was not renamed with that new target name. This information is 

specific to the inner workings of the COMPASS design team and session and is noted here for posterity in 

case future work must build upon these studies.

3.0 Baseline Design (Six ASRG)

3.1 Top Level Design (MEL and PEL) for Six ASRG Case

The six ASRG configuration of the REP Centaur orbiting S/C is built on a hexagonal, two platform 

design with ASRGs attached by twos. The main propulsion system consists of four, long life, Hall 

thrusters, serially operated, and mounted to the side of the S/C. The bottom deck, to which the Hall 

thrusters are attached, carries all the electric and chemical propulsion as well as various power conversion 

systems. The top deck of the S/C, below the antenna, contains all of the avionics and science instruments.

The surface area of the top face of the S/C is dominated by the 2.1 m Ka-band antenna. The S/C is 

designed to optimize ASRG output power and provide science pointing while providing communications.

3.1.1 MEL for Six ASRG Case

The bottoms-up MEL for the final six ASRG mission designed is shown in Table 3.1.
NASA/TM—2011-216971 20

TABLE 3.1.—BASELINE SIX ASRC CASE—TOP LEVEL MEL

Description

REP Centaur Mission Six ASRG

(October 11, 2007)

Qty Unit mass

(kg)

CBE mass

(kg)

Growth

(%)

Growth

(kg)

Total mass

(kg)

REP S/C (Payload and Stage) - - 1071.00 11.68 125.14 1196.14 

Science Payload - - 44.12 30.00 13.24 57.36 

REP Bus - - 1026.88 10.90 111.91 1138.79 

Attitude Determination and Control - - 18.40 20.00 3.68 22.08 

Command and Data Handling - - 33.30 34.26 11.41 44.71 

Communications and Tracking - - 39.00 34.10 13.30 52.30 

Electrical Power Subsystem - - 169.82 16.97 28.82 198.64 

Thermal Control (Non-Propellant) - - 42.81 15.00 6.42 49.23 

Propulsion - - 107.99 26.56 28.68 136.67 

Propellant - - 517.57 0.00 0.00 517.57 

Structures and Mechanisms - - 97.99 20.00 19.60 117.59 

3.1.2 Power Equipment List (PEL) for Six ASRG Case

The power listing for nominal loads is modeled using a 900 W max power (with 30 percent margin on 

all but the EP system) budget. A 30 percent margin is added on all but the REP systems (thruster able to 

handle lower powers down to 500 W). Table 3.2 lists the concepts of operations and what items are turned 

on at which point on the mission trajectory.

TABLE 3.2.—PEL PER SUBSYSTEM OVER MISSION PHASES Propulsion (W) Avionics (W) Comm. (W) Thermal (W) GN&C (W) Power (W) Science (W)

CBE Total(W)30% MarginTotal(W)

Launch 0 24 0 33 27 63 0 146 48.63 195

Star 48 Operation 0 24 420 33 27 63 0 566 174.63 741

S/C separation 16 24 420 33 27 63 0 582 174.63 757

S/C checkout 16 24 420 33 36 63 60 652 195.48 847

REP Thrusting 700 24 0 33 29 63 2 850 49.86 900

REP Coast 16 24 0 33 29 63 2 166 49.86 216

Communications 16 53 420 33 29 63 2 615 184.44 799

Flyby 16 53 420 33 29 63 60 673 201.87 875

Centaur Targeting 700 24 0 33 29 63 2 850 49.86 900

Centaur Science 16 53 0 33 29 63 60 253 75.87 329

Centaur Communications 16 53 420 33 29 63 60 673 201.87 875

3.2 System Level Summary

Table 3.3 breaks out the system level summary of the REP S/C designed in this COMPASS session.

The bottoms-up masses of the subsystem with the growth estimates applied per line item in the subsystem 

yielded a total growth on the dry mass of 22.6 percent. Since the desired total growth is 30 percent per 

COMPASS operating procedure, an additional 7.4 percent of dry mass is carried as system level growth. 

This mass is “flown” through the mission, and adds to the inert mass that the propulsion system has to 

navigate with the trajectory ΔV. 

The performance of the Atlas 551 launch vehicle to the mission C3 of 97 km2

/s2 is 1260 kg. A S/C

adapter mass of 10 kg is taken out of that number to give the available launch performance of 1250 kg to 

the C3 as reported in Table 3.3. The total wet mass of the S/C of 1237 kg is subtracted from the available 

launch performance. The remaining 13 kg is the launch margin available. This mass is not flown along 
NASA/TM—2011-216971 21

the trajectory. In future iterations, the science payload may be increased to get as close to a zero launch 

margin as possible and still fit in the launch vehicle performance.

TABLE 3.3.—SYSTEM INTEGRATION SUMMARY SHEET: SYSTEM LEVEL GROWTH TRACKING

COMPASS study: Radioisotope Electric Propulsion (REP) Study date: November 26, 2007 

GLIDE container: REP_Sept2007: Thereus_6ASRG 

REP Spacecraft MEL rack-up (mass) COMPASS 

REP design 

WBS 

no.

Main subsystems CBE Mass 

(kg) 

Growth 

(kg) 

Total mass 

(kg) 

Aggregate 

growth 

(%) 

01 REP Spacecraft (payload and bus) 1071.0 125.1 1196.1 ------

01.1 Science Payload 44.1 13.2 57.4 30.0

01.2 REP Bus 1026.9 111.9 1138.8 ------

01.2.1 AD&C 18.4 3.7 22.1 20.0

01.2.2 C&DH 33.3 11.4 44.7 34.3

01.2.3 Communications and tracking 39.0 13.3 52.3 34.1

01.2.4 Electric power 169.8 28.8 198.6 17.0

01.2.5 Thermal control 42.8 6.4 49.2 15.0

01.2.6 Propulsion 108.0 28.7 136.7 26.6

01.2.7 Propellant 517.6 ----- -------- ------

01.2.8 Structures and mechanisms 98.0 19.6 117.6 20.0

Estimated REP S/C dry mass 553 125 678.6 22.6

Estimated REP S/C wet mass 1071 125 1196.1 ------

System level growth calculations Total growth 

(%)

Desired system level growth 553 166 719.5 30.0 

Additional growth (carried at system level) ------- 41 -------- 7.4 

Total wet mass with growth 1071 166 1237.0 

Available launch performance to C3 (kg) 1250.2 

Launch margin available (kg) 13.2 

Figure 3.1.—REP Science S/C dimensions.
NASA/TM—2011-216971 22

3.3 Design Concept Drawing and Description

Figure 3.1 shows a side view of the REP Centaur S/C, without the Star 48 engine attached, with 

dimensions. All dimensions are in metric units.

4.0 Subsystem Breakdown 

4.1 Attitude Control System (ACS)

The starting design is borrowed from NH

• ACS hydrazine

• Two Star Trackers (Adcole Corporation) Figure 4.1. These star trackers were the ones used on the 

NH S/C (Adcole)

• Eight Sun Sensors (EDO Corp Barnes Engineering Division)

• Four Reaction Wheels (Valley Forge Bearcat 5 Nms)

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

Table 4.1 lists the items in the ACS MEL for the COMPASS REP S/C design. All growth allowances 

follow the AIAA MGA schedule in Section 2.4. Note that an IMU from Northrop Grumman was added to 

the MEL (Figure 4.2).

Figure 4.1.—Adcole Star Tracker.

Figure 4.2.—Northrop Grumman 

Scalable Inertial Reference Unit 

(NG SIRU) for space.
NASA/TM—2011-216971 23

TABLE 4.1.—ACS BOTTOMS-UP MEL

Description

REP Centaur Mission Six ASRG

(October 11, 2007)

Qty Unit mass

(kg)

CBE mass

(kg)

Growth

(%)

Growth

(kg)

Total mass

(kg)

REP Spacecraft (Payload and Stage) - ---- 1071.00 11.68 125.14 1196.14 

REP Bus - ---- 1026.88 10.90 111.91 1138.79 

AD&C - ---- 18.40 20.00 3.68 22.08 

GN&C - ---- 18.40 20.00 3.68 22.08 

Sun sensors 8 0.01 0.04 20.00 0.01 0.05 

Reaction wheels 4 1.27 5.08 20.00 1.02 6.10 

Star Trackers 2 3.19 6.38 20.00 1.28 7.66 

IMU 1 6.90 6.90 20.00 1.38 8.28 

Figure 4.3 shows the avionics deck of the REP S/C. This deck is where all the electronics are located. 

The graphic has all non-electronics items in the S/C invisible, and highlights the components of Avionics, 

ACS, Power and Communications, aside from the main antenna, and the science instruments (all labeled 

at the bottom of the graphic).

4.1.1 ACS Trades

No trades were done on the ACS, except a conceptual trade of whether or not to include reaction wheels.

4.1.2 ACS Analytical Methods

The design was based on current hardware. 

Figure 4.3.—Science and avionics deck of REP S/C.
NASA/TM—2011-216971 24

4.1.3 ACS Recommendation

Analysis into the amount of ΔV necessary for station keeping and attitude control throughout the 

mission life needs to be performed to determine whether the 50 m/s assumption is sufficient.

Additionally, further research is necessary to determine whether the star trackers and sun sensors are 

capable of operating at the distances of the Centaur bodies at the EOL of the trajectory.

4.2 Communications

4.2.1 Communications Requirements

Provide uplink and downlink capability throughout the primary and/or extended mission. Meet 

science mission requirements during Thereus orbital operations of 8 hr/day of downlink pointed to Earth 

with a minimum 6.3 kbps downlink at 70-m dish (or about 147 Mbits/day of downlink including a 

minimum of 10 percent for housekeeping).

4.2.2 Communications Assumptions

Assume the DSN will be capable of supporting Ka-Band downlink via a 70-m (or equivalent antenna 

array) antenna by through 2024. The design is based on the NH concept of two onboard, integrated 

electronics modules (IEMs). The overall harness requirements are reduced if the NH IEM design is 

implemented.

4.2.3 Communications Design and MEL

• REP orbiter communications subsystem consists of 

– A fixed 2.1-m diameter X/Ka-Band high gain antenna (HGA)

– Two IEMs, based on the NH, housing many S/C functions, including C&DH, instrument 

interface circuitry, telemetry interface, solid state recorder, and receiver and exciter sections 

of the communications subsystem

– Two 200-W Traveling Wave Tube Amplifier (TWTA) to provide high power radio frequency 

communications power (RF downlink output)

– RF switch assembly to interconnect antenna with two TWTAs and the rest of the 

communications subsystem

– Cabling

• Ka-Band downlink to 70-m ground stations

– Ka-Band downlink frequency: 32 GHz

• X-Band support between orbiter and Earth’s 70-m ground stations

– Forward frequency: 8.4 GHz

– Return frequency: 7.75 GHz

– Use of a fixed 2.1-m HGA

– 200 W RF power

– TT&C will share the uplink and downlink bandwidth

Table 4.2 lists the items in the Communications system MEL for the COMPASS REP S/C design. All 

growth allowances follow the AIAA MGA schedule in Section 2.4 and do not contain the additional 

8.8 percent carried at the system level.
NASA/TM—2011-216971 25

TABLE 4.2.—REP COMMUNICATIONS SYSTEM MEL

Description 

REP Centaur Mission Six ASRG

(October 11, 2007)

Qty Unit mass

(kg)

CBE mass

(kg)

Growth

(%)

Growth

(kg)

Total mass

(kg)

REP Spacecraft (Payload and Stage) - ---- 1071.00 11.68 125.14 1196.14 

REP Bus - ---- 1026.88 10.90 111.91 1138.79 

Communications and Tracking - ---- 39.00 34.10 13.30 52.30 

X/Ka HGA - ---- 27.00 31.48 8.50 35.50 

Transmitter/receiver 2 4.00 8.00 30.00 2.40 10.40 

Power amp 2 3.00 6.00 30.00 1.80 7.80 

Switch unit 0 0.00 0.00 0.00 0.00 0.00 

Antenna 1 9.00 9.00 30.00 2.70 11.70 

Band pass filter 0 0.00 0.00 0.00 0.00 0.00 

Band reject filter 0 0.00 0.00 0.00 0.00 0.00 

Sensor 0 0.00 0.00 0.00 0.00 0.00 

Cabling 2 2.00 4.00 40.00 1.60 5.60 

Diplexer 0 0.00 0.00 0.00 0.00 0.00 

Coupler 0 0.00 0.00 0.00 0.00 0.00 

Miscellaneous no. 1 0 0.00 0.00 0.00 0.00 0.00 

Miscellaneous no. 2 0 0.00 0.00 0.00 0.00 0.00 

Ka-band antenna - ---- 0.00 0.00 0.00 0.00 

Transponder 0 0.00 0.00 0.00 0.00 0.00 

RF assembly 0 0.00 0.00 0.00 0.00 0.00 

Processing module 0 0.00 0.00 0.00 0.00 0.00 

Antenna 0 0.00 0.00 0.00 0.00 0.00 

Communications instrumentation - ---- 12.00 40.00 4.80 16.80 

Coaxial cable 2 6.00 12.00 40.00 4.80 16.80 

Installation—mounting and circuitry 0 0.00 0.00 0.00 0.00 0.00 

4.2.4 Communications Trades

Further analysis on optimization of the communications components was not performed.

4.2.5 Communications Analytical Methods

The link budgets provides for values of RF transmit powers at 40 and 200 W and antenna gains for 

Ka-Band and X-Band downlinks. Link margins of 3 dB or better exist for all links. 

4.2.6 Communications Risk Inputs

Heritage of success of design and components assumed.

4.2.7 Communications Recommendation

In the future, further analysis should be done to consider the communications system used in the NH

mission, as its performance behavior is characterized during flight mission ops. Figure 4.4 shows a 

detailed block diagram of the NH full S/C design. This includes C&DH as well as communications.

4.3 Command and Data Handling (C&DH)—Avionics

4.3.1 Avionics Requirements

The design requirements, from the science payload and the REP Bus, for the C&DH system were as 

follows

• Storage for 7 days of data or fly-by (TBD, est. 8 to 16 GB)

• Avionics for systems command, control, and health management

• Payload control will be done by the C&DH system
NASA/TM—2011-216971 26

• Single fault tolerant avionics

4.3.2 Avionics Assumptions

• All electronics are ≥65 Krad avionics

• Cabling mass is estimated as 50 percent of the avionics hardware

• Avionics spares are cold spares to minimize power consumption

• NH S/C was used as the starting point for the avionics hardware design

4.3.3 Avionics Design and MEL

All avionics components used in the design are based on commercially available components from 

British Aerospace (BAE) and SEAKR Engineering, Inc. (SEAKR). There are two independent avionics 

boxes to provide for single fault tolerance. Each avionics box contains a GN&C/C&DH RAD6000 

processor, 256 MB GN&C solid-state memory card, solid state recorder (SSR) card, a communications

interface card, and a payload interface card. The 1553 processor is used for communications between the 

GN&C processor and hardware, i.e., Star trackers, IMUs, etc. The GN&C and C&DH computers 

communicate via the 1553 bus. This new REP avionics design, as shown below, is based on the NH S/C.

With the exception of GN&C and C&DH, the processors for each IEM are combined into one, using 

either RAD6000 or RAD750.

Table 4.3 lists the components used in the COMPASS C&DH MEL design. These are the inputs from 

the subsystem lead. All growth allowances follow the AIAA MGA schedule in Section 2.4 and do not 

contain the additional 8.8 percent carried at the system level.

Figure 4.4.—NH full S/C design.
NASA/TM—2011-216971 27

TABLE 4.3.—AVIONICS MEL MASS DETAILS

Description 

REP Centaur Mission Six ASRG

(October 11, 2007)

Qty Unit mass

(kg)

CBE mass

(kg)

Growth

(%)

Growth

(kg)

Total mass

(kg)

REP Spacecraft (Payload and Stage) - ------ 1071.00 11.68 125.14 1196.14 

REP Bus - ------ 1026.88 10.90 111.91 1138.79 

C&DH - ------ 33.30 34.26 11.41 44.71 

Flight computer 2 8.00 16.00 25.00 4.00 20.00 

Command and telemetry computer 0 0.00 0.00 0.00 0.00 0.00 

Data interface unit 2 2.00 4.00 30.00 1.20 5.20 

Data bus operations amplifier 0 0.00 0.00 0.00 0.00 0.00 

Operations recorder 2 1.10 2.20 30.00 0.66 2.86 

Command and control harness (data) 1 11.10 11.10 50.00 5.55 16.65 

Instrumentation and wiring - ------ 0.00 0.00 0.00 0.00 

Operational Instrumentation, sensors 0 0.00 0.00 0.00 0.00 0.00 

Data cabling 0 0.00 0.00 0.00 0.00 0.00 

4.3.4 Avionics Trades

Off-the-shelf (OTS) components used in the design. Further analysis of the choice of components to 

needs be done.

4.3.5 Avionics Analytical Methods

Much of the design is based on the NH S/C.

4.3.6 Avionics Risk Inputs

Assuming heritage on components.

4.3.7 Avionics Concerns, Comments, Recommendations

• No ultra-stable oscillator (USO)/Atomic Clock included in avionics hardware. Should it be 

included in communications system?

• Processing power of the RAD6000 is assumed to be adequate for GN&C, C&DH, and science 

payload.

• Storage requirements are driven by fly-by storage needs (which are still being estimated).

• Only one SSR would be active at a time and thus susceptible to single event upsets (SEU).

• Total radiation dose is a concern with all deep space missions. This preliminary design has 

attempted to use only hardware which has already been proven in a deep space mission to assure 

the life of the electronics over the 12-yr mission.

4.4 Electrical Power System

4.4.1 Power Requirements

Six ASRG’s (12 general purpose heat source (GPHS)) are designed to provide 960 W to power the 

REP S/C at beginning of life (BOL). The system is designed to provide 900 W to the REP S/C at EOL 

(10-yr). There are negligible thermal interactions between the ASRG’s. Figure 4.5 shows a typical ASRG 

with the main components called out in the graphic. The six ASRGs are connected together with via a 

Shunt Regulator/Bus Protection (RBI) assembly. This RBI isolates the ASRG’s from S/C bus and each 

other and follows load demands from S/C bus. There is an approximately 6 percent loss through the RBI 

and monitoring circuitry (94 percent of power flows through to loads) with 53 W used for fault detection/

monitoring. Included in this system is a bus capacitance of 3000 μf which provides some bus rigidity.

Power cabling and harness systems design assumes a 1 percent line loss.
NASA/TM—2011-216971 28

Figure 4.5.—ASRG computer aided design (CAD) model.

4.4.2 Power Assumptions

This baseline power system design consisted of six ASRG for the generation of power.

4.4.3 Power Design and MEL

Minimize power for non-propulsion during EP operation (minimize plutonium needed).

Specific performance details on each ASRG unit are as follows

• Power: 160 W at 28 ± 0.2 V BOL

150 W at 28 ± 0.2 V EOL (10 yr)

• End of mission (EOM) (Deep Space (14 yr)) -126 W

• Mass: 19.5 kg with mounting isolator plate

• Envelope: 30.5-cm W, 46-cm H, 76-cm L (12-in. W, 18-in. H, 30-in. L)

• Specific Power: 8 We/kg

ASRG Design Attributes

• Two Stirling converters

– Co-axially aligned for dynamic balance

– One GPHS module per converter

• Integrated, single-fault tolerant controller 

• Autonomous operation and fault isolation from S/C

• S/C disturbance torque requirement < 35 N-m

– Based on 1000 kg, 1-m cube S/C with 5-μrad pointing accuracy and a safety factor of 5

Table 4.4 lists the items in the power system MEL for the COMPASS REP S/C design. All growth 

allowances follow the AIAA MGA schedule in Section 2.4. Figure 4.6 shows the power and propulsion 

deck of the REP S/C. The ASRGs are mounted to the bus structure via trusses, at a 120° angle between 

sets of pairs around the perimeter of the main bus.
NASA/TM—2011-216971 29

TABLE 4.4.—POWER SYSTEM MEL FOR REP BUS

Description 

REP Centaur Mission Six ASRG

(October 11, 2007)

Qty Unit mass

(kg)

CBE mass

(kg)

Growth

(%)

Growth

(kg)

Total mass

(kg)

REP Spacecraft (Payload and Stage) - ------- 1071.00 11.68 125.14 1196.14 

REP Bus - ------- 1026.88 10.90 111.91 1138.79 

Electrical Power Subsystem - ------- 169.82 16.97 28.82 198.64 

Radioisotope Power System - ------- 116.82 10.00 11.68 128.50 

RPS main system 6 19.47 116.82 10.00 11.68 128.50 

Miscellaneous no. 2 0 0.00 0.00 0.00 0.00 0.00 

Power management and distribution - ------- 26.75 15.00 4.01 30.76 

Power management/control electronics 0 0.00 0.00 0.00 0.00 0.00 

Power distribution/monitoring wiring harness 0 0.00 0.00 0.00 0.00 0.00 

DC Switchgear/Shunt Regulator 1 26.75 26.75 15.00 4.01 30.76 

Miscellaneous no. 2 0 0.00 0.00 0.00 0.00 0.00 

Power cable and harness subsystem - ------- 26.25 50.00 13.13 39.38 

Spacecraft bus harness 1 5.25 5.25 50.00 2.63 7.88 

PMAD Harness 1 5.25 5.25 50.00 2.63 7.88 

Electric propulsion harness 1 5.25 5.25 50.00 2.63 7.88 

RPS to S/C harness 1 5.25 5.25 50.00 2.63 7.88 

Power cabling 1 5.25 5.25 50.00 2.63 7.88 

Figure 4.6.—Power and propulsion system deck.
NASA/TM—2011-216971 30

4.4.4 Power Trades

For the power system, two other power system design options were considered but rejected

Option 1 (Not used since more power was needed than the 10/1 power ratio allowable on the dualwindings allowed. Significant advantages to the mission exist if the full ASRG power is available during 

non-thrusting periods)

• Direct drive the Hall thrusters 

• Use dual wound alternator (providing 600 V and 28 V, 100 Hz AC)

• 10/1 power ratio on dual alternators

• The 600 V AC converts to 400 V DC

• Power to Thrusters EOM 646 W

• Power to Payload EOM 76 W

Option 2 (not used due to the heavy converter needed.)

• Each ASRG provides 28 V DC as designed

• DC/DC conversion to 400 V DC for hall thruster

• The current mass estimate of single 600 W DC/DC converter at 30 kg

• Eight ASRG provides (1120 W BOL, 1040 W EOL) 750 W power into the thruster with excess 

14 W EOL

• Loss of a single SRG would then provide ~650 W into Thruster

• If the spacing of the ASRGs is changed to a 90°: Does this provide better viewing for radiative 

coupling between the ASRGs? How does this affect structure and loads balancing?

Table 4.5 lists the impact of trade in the number of ASRGs and total power available, as well as 

excess power to be radiated as mentioned in the Option 2 analysis up above.

TABLE 4.5.—TRADE ANALYSIS OF VARYING NUMBER OF SRGS

Number SRGs 4 5 6 7 8

Power (EOL, 10 yr) 130 130 130 130 130

Total power EOL (W) 520 650 780 910 1040

Into thruster (W) 250 400 500 650 750

PPU, line loss 25 40 50 65 75

Housekeeping (cruise only) 155 155 155 155 155

Housekeep margin (30%) 46.5 46.5 46.5 46.5 46.5

Excess 44 9 29 –7 14

4.4.5 Power Analytical Methods

ASRGs based on current hardware designs.

4.4.6 Power Recommendations for Future Analysis

• Would a change to a 90° separation between the ASRGs provide a better or worse radiative 

coupling between the ASRGs?

• Plate/strut support in ‘middle’

– Vibration/thermal leak to science payload instruments

• Science Payload view of SRGs, is a 90° or 120° (current design) separation better? Should the 

ASRG's be offset above or below the flight instrument deck?

Is there necessary plutonium availability?
NASA/TM—2011-216971 31

4.5 Structures and Mechanisms

The following section describes the design structures and mechanisms methodology and details.

4.5.1 Structures and Mechanisms Requirements

The REP S/C structure must contain necessary hardware for research instrumentation, avionics, 

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

and provide minimum deflections, sufficient stiffness, and vibration damping. The goal of the design is to 

minimize weight of the components that make up the structure of the S/C bus, and must fit within 

confines of launch vehicle.

4.5.2 Structures and Mechanisms Assumptions

Because of the use of the Star 48 engine, the structure must be able to withstand an axial acceleration 

load of up to 6g maximum. The launch vehicle also imparts a maximum of 3.5g lateral acceleration. The 

structure must also accommodate a Conical Star 48 adapter. The Hall thrusters must thrust through the 

center of mass (COM) and need to be canted. The structural design and S/C layout must accommodate a 

thruster gimbal angle.

The basic assumptions made in the design process of the S/C bus structure were

• Material: Al alloy 2090-T3

• Space frame with tubular members

• Composite sandwich structure shelf assumed to be 100 percent Al composition using Al 2090-T3 

face sheets and an Al honeycomb core with the trade name, Alcore Higrid.

• Welded and threaded fastener assembly

Figure 4.7.—REP Main Bus.
NASA/TM—2011-216971 32

4.5.3 Structures and Mechanisms Design and MEL

The basic structural design of the REP S/C shown in Figure 4.7 consists of

• Tubular space frame in a hexagonal configuration

• Deck consists of a composite sandwich architecture with an Al 25 mm thick low density 

honeycomb core and 1.5 mm thick Al face sheets to mount hardware

• Thin sheets utilized as sheer panels and to enclose the structure

• Struts are used to support the paired ASRGs externally. The ASRGs are mounted to plates with 

vibration isolators.

All growth allowances follow the AIAA MGA schedule in Section 2.4.

The initial assumptions used in the design were: 2600 kg, 6g axial loading, 3.5g lateral loading (not 

concurrent with max axial loading), Al 2090-T3 and a 1.4 safety factor. The maximum stress in axial 

members was set at 58 MPa. Initial assumptions for lateral load capacity were based on S/C heritage 

design. The Secondary structural mass (called installations in the MEL above in Table 4.6) assumes a 

4 percent factor of the mounted hardware current best estimate (CBE) mass.

TABLE 4.6.—REP SIX ASRG S/C STRUCTURES AND MECHANISMS BOTTOMS-UP MEL

Description 

REP Centaur Mission Six ASRG

(October 11, 2007)

Qty Unit mass

(kg)

CBE mass

(kg)

Growth

(%)

Growth

(kg)

Total mass

(kg)

REP Spacecraft (Payload and Stage) - ------ 1071.00 11.68 125.14 1196.14 

REP Bus - ------ 1026.88 10.90 111.91 1138.79 

Structures and Mechanisms - ------ 97.99 20.00 19.60 117.59 

Structures - ------ 78.41 20.00 15.68 94.09 

Primary structures - ------ 40.95 20.00 8.19 49.15 

Main bus structure 1 40.95 40.95 20.00 8.19 49.15 

Secondary structures - ------ 37.45 20.00 7.49 44.94 

Balance mass 0 0.00 0.00 0.00 0.00 0.00 

Tank supports and bracketry 1 18.18 18.18 20.00 3.64 21.82 

SRG support structure 3 5.51 16.54 20.00 3.31 19.85 

SRG vibration isolation hardware 3 0.91 2.73 20.00 0.55 3.27 

Mechanisms - ------ 19.58 20.00 3.92 23.50 

Adaptors and separation - ------ 4.72 20.00 0.94 5.67 

Spacecraft adapter 1 4.72 4.72 20.00 0.94 5.67 

Separation mechanism (pyros) 0 0.00 0.00 0.00 0.00 0.00 

Miscellaneous no. 1 0 0.00 0.00 0.00 0.00 0.00 

Installations - ------ 14.86 20.00 2.97 17.83 

Science payload installation 1 1.76 1.76 20.00 0.35 2.12 

C&DH installation 1 1.33 1.33 20.00 0.27 1.60 

Communications and tracking installation 1 1.56 1.56 20.00 0.31 1.87 

GN&C installation 1 0.74 0.74 20.00 0.15 0.88 

Electrical power installation 0 0.00 0.00 0.00 0.00 0.00 

Thermal control installation 1 3.03 3.03 20.00 0.61 3.63 

Electric Propulsion installation 1 6.44 6.44 20.00 1.29 7.73 

4.5.4 Structures and Mechanisms Trades

Trades were conducted on the use of composite for the main bus compartment structure. The frame of 

the main bus is sized to accommodate volume and space requirements for antenna, ASRGs, and 

instrumentation while fitting within confines of the launch vehicle.
NASA/TM—2011-216971 33

4.5.5 Structures and Mechanisms Analytical Methods

Preliminary structural analysis and modeling was performed using the given launch loads and 

dimensions of the desired S/C bus. An additional installations mass was held for each subsystem in the 

mechanisms section of the structures system. These installations were modeled using 4 percent of the 

CBE dry mass of each of the subsystems. No growth margin was applied to that installation mass.

4.5.6 Structures and Mechanisms Risk Inputs

Risk analysis is still to be performed.

4.5.7 Structures and Mechanisms Recommendation

A more detailed structural analysis for loads and vibrations using a modeling tool, i.e., finite element 

analysis (FEA) would be beneficial in further modeling the structure with assurance of sustaining launch 

loads. Analysis is needed to look into using the current shelf face sheets, outer sheets, and/or support struts 

substituted with graphite/polymer composites for further weight savings but possibly at increased cost.

4.6 Propulsion and Propellant Management

4.6.1 Propulsion and Propellant Management Requirements

The REP Centaur mission is a relatively long life mission. The goal of each subsystem and the 

propulsion system is to minimize the overall mass of the REP bus in order to deliver the maximum 

amount of payload.

The mission requirements that impact the choice of EP system components are as follows

• Approximately 10 yr mission (~ 600 kg throughput)

• 500 to 1000 W total power

• Essentially constant power from ASRG power supply

• Preliminary optimum Isp ~ 2050 s

4.6.2 Propulsion and Propellant Management Assumptions

The baseline system redundancy assumption was based on single string units. In other words, a 

propulsion system unit consists of a string of thruster, PPU, gimbal, and Propellant Management System 

(PMS). Spares or redundant units are assumed to consist of all of the above subsystems. It is important to 

note that the mission is already modeled with a 90 percent duty cycle. So, 10 percent of the time, the S/C

is coasting along its trajectory.

The baseline propulsion system design consists of the following items

• One active 750 W long life Hall engine with two extra engines for lifetime issues and one cold 

spare

• Four PPU: no cross-strap

• Two-axis range of motion: TBD

• 5.0 percent Xe navigation allowance, 3.6 percent Xe residuals

• PSI cylindrical COPV Xe tanks

• OTS hydrazine system with NH heritage

• Instruments to be fully operational with 900 W ASRG power supply and thrusters full on.

Hall thruster and PPU performance and masses were based on published or in-house calculations by 

the GRC RPP branch. Thruster performance over a range of Isp was examined as a series of custom 

designs, rather than a single thruster design throttled over a range of Isp. Thruster mass was assumed to be 
NASA/TM—2011-216971 34

50 percent greater than a commercial thruster (SPT-70) which operates at a similar power level. PPU 

performance and mass were based on a single module of a PPU unit under development and test at GRC.

4.6.3 Propulsion and Propellant Management Design Trades

The trades considered in designing the two major propulsion systems on the S/C (main and RCS) are 

as follows

• ACS hydrazine

– OTS blow-down similar to NH

– Single tank with ~20 kg hydrazine

The possible main EP system options to be considered for this design are

• New Advanced Technology Small Hall Thruster (Figure 4.8)

– Based on ongoing High Voltage Hall Accelerator (HiVHAC) program at GRC

– Optimized design to allow up to 2000 s Isp at powers below 1 kW

– Allows long life needed for mission (Figure 4.9)

• Derated HiVHAC (Figure 4.10)

– Maximum Isp at 1 kW ~1570 s

– Performance inadequate for range of REP missions

• Commercial-off-the-shelf (COTS)

– SPT-70/BPT-600

– 600 W, ~ 1500 s 

– Limited life/throughput (35 to 50 kg)

• Low power (20 cm) Ion

The Advanced Hall thruster option was initially chosen both for its potential for direct drive operation 

(see power system discussion), and for its superior performance in terms of efficiency (or equivalently, 

thrust-to-power) at the low power levels characteristic of REP. The 20 cm ion thruster projected 

performance was inferior to that of the Hall below 1 kW and at 2000 s or less Isp. The commercial Hall 

thrusters increased system mass and complexity through the increased number of propulsion strings (13 

or more) needed to meet lifetime and redundancy requirements.

Figure 4.8.—Advanced Technology Small 

Hall Thruster.
NASA/TM—2011-216971 35

Figure 4.9.—Assumed performance of Small Hall Thruster.

Figure 4.10.—HiVHAC Hall Electric Thruster.

The possible EP thruster system options, once an EP thruster type has been chosen, to be considered 

are

• Hall

– Standard PPU 

– “Direct drive” from Stirling alternator

Because of limitations in the ASRG alternator design, the “direct drive” option was discarded and a 

standard PPU option was selected.

4.6.4 Propulsion and Propellant Management Design and MEL

Figure 4.10 is a graphic of the Advanced Technology Small Hall Thruster being considered in this 

conceptual design.

4.6.4.1 Main EP System (Xe)

The main EP system is comprised of

 Four extended life, high Isp Hall Thrusters (three operating, one active spare)

− Thruster performance 

• 30,000 hr life, 300 to 700 V

 Standard PPU

 Two cylindrical, COPV high pressure (2800 psi) Xe tanks
NASA/TM—2011-216971 36

 Propellant distribution system: Single string PMS to each thruster from balanced tank feed

 Thermal details of prop system

− Number of heaters on tanks, etc.

 Total propellant

− 540 kg used

− 8.6 percent residual + margin

Figure 4.11 is a schematic of the EP system and propellant management tankage, etc. The main EP 

subsystem is comprised of: four HiVHAC Hall Thrusters—three operating, one spare, gimbals on each 

thruster was used for thrust vector control, and two COPV titanium (Ti)-lined high-pressure cylindrical 

storage tanks for the Xe propellant (nominal) storage. The Xe distribution system was based on newly 

developed pressure and flow control units and four PPU for delivering power to each ion thruster.

Figure 4.11.—EP system schematic.
NASA/TM—2011-216971 37

4.6.4.2 Secondary RCS System (Hydrazine)

The attitude-reaction control propulsion subsystem was comprised of: eight 0.25 lbf monoprop 

thrusters placed around the S/C body The Rocket Research MR-103H monomethyl hydrazine and 

nitrogen tetroxide bipropellant system (MMH/NTO) thrusters were used. Fuel was stored in an Al-Li

Metallic Tank Single spherical tank using a blow down pressurization with discrete He pressurization 

system (Cassini heritage). The propellant distribution system used a design similar to systems developed 

for the Constellation program, including fault tolerance configuration. Multiple tank and line heaters were 

included in the mass model to prevent propellants from freezing. Additionally, insulation was included for 

the same elements. The Instrumentation included was a nominal suite of temperature and pressure 

sensors. Figure 4.12 shows the notional setup from hydrazine tank to the hydrazine thrusters and 

illustrated the feed system linkages between the two.

Table 4.7 lists the propulsion system hardware MEL. All growth allowances follow the AIAA MGA 

schedule in Section 2.4.

Table 4.8 lists the propellant used in this mission. Note, the margins and residuals are called out as 

separate line items in this mass listing, and no additional weight growth schedule (WGS) is necessary on 

the propellants.

Figure 4.12.—RCS system schematic.
NASA/TM—2011-216971 38

TABLE 4.7.—ELECTRIC AND CHEMICAL PROPULSION SYSTEM MEL

Description 

REP Centaur Mission Six ASRG

(October 11, 2007)

Qty Unit mass

(kg)

CBE mass

(kg)

Growth

(%)

Growth

(kg)

Total mass

(kg)

REP Spacecraft (Payload and Stage) - ------- 1071.00 11.68 125.14 1196.14 

REP Bus - ------- 1026.88 10.90 111.91 1138.79 

Propulsion - ------- 107.99 26.56 28.68 136.67 

Primary EP system - ------- 9.00 12.00 1.08 10.08 

Primary EP thrusters 4 2.25 9.00 12.00 1.08 10.08 

EPS power processing and control 0 0.00 0.00 0.00 0.00 0.00 

EPS structure - ------- 0.00 0.00 0.00 0.00 

EP thruster pod 0 0.00 0.00 0.00 0.00 0.00 

EP thruster boom 0 0.00 0.00 0.00 0.00 0.00 

Miscellaneous no. 1 0 0.00 0.00 0.00 0.00 0.00 

EPS thermal control subsystem - ------- 0.00 0.00 0.00 0.00 

EPS MLI 0 0.00 0.00 0.00 0.00 0.00 

EPS heaters and sensors 0 0.00 0.00 0.00 0.00 0.00 

Miscellaneous no. 1 0 0.00 0.00 0.00 0.00 0.00 

Propellant management - ------- 66.29 31.18 20.67 86.96 

Xe propellant tank(s) 2 27.39 54.77 30.00 16.43 71.21 

High pressure feed system 1 7.62 7.62 30.00 2.29 9.90 

Low pressure feed system 0 0.00 0.00 0.00 0.00 0.00 

Residual Xe propellant (nondeterministic) 0 0.00 0.00 0.00 0.00 0.00 

Temperature sensors 1 3.90 3.90 50.00 1.95 5.85 

PPU - ------- 16.00 12.00 1.92 17.92 

PPU mass 4 4.00 16.00 12.00 1.92 17.92 

Cabling 0 0.00 0.00 0.00 0.00 0.00 

RCS - ------- 16.70 30.00 5.01 21.71 

RCS tank subassembly 1 2.79 2.79 30.00 0.84 3.62 

RCS propellant management subassembly 1 9.45 9.45 30.00 2.83 12.28 

RCS thruster subassembly 2 2.23 4.46 30.00 1.34 5.80 

TABLE 4.8.—PROPELLANT MEL

Description 

REP Centaur Mission Six ASRG

(October 11, 2007)

Qty Unit mass

(kg)

CBE mass

(kg)

Growth

(%)

Growth

(kg)

Total mass

(kg)

REP Spacecraft (Payload and Stage) - -------- 1071.00 11.68 125.14 1196.14 

REP Bus - -------- 1026.88 10.90 111.91 1138.79 

Propellant - -------- 517.57 0.00 0.00 517.57 

Primary EP propellant - -------- 489.53 0.00 0.00 489.53 

Primary EP propellant used 1 450.76 450.76 0.00 0.00 450.76 

Primary EP propellant residuals (unused) 1 16.23 16.23 0.00 0.00 16.23 

Primary EP propellant performance margin (unused) 1 22.54 22.54 0.00 0.00 22.54 

RCS propellant - -------- 27.94 0.00 0.00 27.94 

RCS used 1 27.26 27.26 0.00 0.00 27.26 

RCS residuals 1 0.68 0.68 0.00 0.00 0.68 

Pressurant 1 0.10 0.10 0.00 0.00 0.10 

4.6.5 Propulsion and Propellant Management Risk Inputs

The EP system is expected to have a performance not previously demonstrated at the power levels of 

interest. Because the chosen trajectory does not have large coast periods, the mission performance would 
NASA/TM—2011-216971 39

not be achievable without reaching the thruster performance goals. Also, the mission is not very “robust” 

in that missed thrust periods could place the mission at risk. Mitigation would include imposing coast 

period requirements or mandating a propulsion system with a less than optimal Isp, therefore higher thrust; 

to increase the capability to makeup mission thrust periods. Effect of thrust on coast periods is shown in 

the initial analysis done for the Bienor mission in Figure D.4. 

A development schedule risk lies in the choice of an undeveloped thruster technology. Because the 

projected thruster performance and life have not been demonstrated, additional schedule and cost 

uncertainty are introduced. The assigned levels of risk for the various consequence categories are

• Cost: 4

• Schedule: 4

• Performance: 4

• Safety: 1

• Risk mitigation can be achieved by appropriate scheduling assumptions and by incorporating 

adequate preliminary testing into the development program.

4.6.6 Propulsion and Propellant Management Recommendation

Future trades to reduce mass on the main propulsion system are as follows

• Lower power ion thruster (8 cm) for possible mass benefit or mission benefit

• Re-optimize mission for lower Isp or higher power to capture HiVHAC or COTS regime

Further trades on the secondary propulsion system to reduce mass are

• Utilize primary propulsion for some maneuvering, and modeling the trade between attitude 

control using only wheels (control moment gyros) versus propulsion or a combination of the two.

4.7 Thermal Control

4.7.1 Thermal Requirements

The thermal requirements for the REP Centaur mission were to provide a means of cooling the S/C

during operation as well as provide heat to vital components and systems to maintain a minimum 

temperature throughout the mission. The goal of the thermal control system is to provide for the rejection 

of heat and maintain a safe operating environment for the electronics and other systems on the S/C.

The maximum heat load to be rejected by the thermal system was 125 W from the electronics, and the 

desired operating temperature for the electronics and propellant was 300 K. The ASRGs have dedicated 

built in thermal control systems and therefore were not part of the S/C thermal system.

4.7.2 Thermal Assumptions

The thermal subsystem used in the COMPASS SEP Enceladus design (CD–2007–09) was used as a 

starting point for these analyses. The system was modeled for Deep Space Operation. The radiator always 

sees deep space with a small (0.05) view factor to the Sun.

The assumptions utilized in the analysis and sizing of the thermal system were based on the 

operational environment. It was assumed that the worst case operational conditions would be in near 

Earth space. The following assumptions were utilized to size the thermal system. 

• The view factors for the radiator to the Earth, lunar surface and ASRG radiators were assumed to 

be 0.1, 0.25 and 0.1, respectively

• The maximum angle of the radiator to the Sun was 15°

• The radiator temperature was 320 K
NASA/TM—2011-216971 40

4.7.3 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 during periods of inactivity. 

Excess heat is collected from a series of Al 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 

S/C body and are integrated to the radiator are protected with a micrometeor shield. The radiator has 

exterior louvers on it to provide some control over its heat transfer capability. 

The radiator was sized with approximately 50 percent margin in its heat rejection area. This added 

margin insures against unforeseen heat loads, degradation of the radiator and increased view factor 

toward the sun or other thermally hot body not accounted for in the analysis. 

To provide internal heating for the electronics and propulsion systems a series of electric heaters are 

utilized. These heaters are controlled by an electronics controller, which reads a series of thermocouples 

through a data acquisition system. 

MLI is also utilized on the S/C, and propellant system to regulate and maintain the desired temperatures. 

Table 4.9 lists the items in the thermal system MEL for the COMPASS REP S/C design. All growth 

allowances follow the AIAA MGA schedule in Section 2.4 and do not contain the additional 8.8 percent

carried at the system level. Systems modeled: Micrometeor Shielding on Radiator, Radiator Panels, 

Thermal Control of Propellant Lines and Tanks, S/C Insulation, Avionics and Power Management and 

Distribution (PMAD) Cooling.

TABLE 4.9.—THERMAL SYSTEM MEL

Description 

REP Centaur Mission Six ASRG

(October 11, 2007)

Qty Unit mass

(kg)

CBE mass

(kg)

Growth

(%)

Growth

(kg)

Total mass

(kg)

REP Spacecraft (Payload and Stage) --- ----- 1071.00 11.68 125.14 1196.14 

REP Bus --- ----- 1026.88 10.90 111.91 1138.79 

Thermal control (nonpropellant) --- ----- 42.81 15.00 6.42 49.23 

Active thermal control --- ----- 16.90 15.00 2.54 19.44 

Heaters 15 1.00 15.00 15.00 2.25 17.25 

Thermal control/heaters circuit 2 0.20 0.40 15.00 0.06 0.46 

Data acquisition 1 1.00 1.00 15.00 0.15 1.15 

Thermocouples 50 0.01 0.50 15.00 0.08 0.58 

Miscellaneous no. 1 0 0.00 0.00 15.00 0.00 0.00 

Miscellaneous no. 2 0 0.00 0.00 15.00 0.00 0.00 

Passive thermal control --- ----- 23.76 15.00 3.56 27.33 

Heat Sinks 4 3.46 13.85 15.00 2.08 15.93 

Heat Pipes 1 1.02 1.02 15.00 0.15 1.17 

Radiators 1 2.34 2.34 15.00 0.35 2.69 

MLI 1 3.77 3.77 15.00 0.57 4.33 

Temperature sensors 25 0.01 0.25 15.00 0.04 0.29 

Phase change devices 0 0.00 0.00 15.00 0.00 0.00 

Thermal coatings/paint 1 0.93 0.93 15.00 0.14 1.07 

Micrometeor shielding 0 0.00 0.00 15.00 0.00 0.00 

S/C RTG MLI 1 0.00 0.00 15.00 0.00 0.00 

S/C engine MLI 1 1.60 1.60 15.00 0.24 1.84 

Semi-passive thermal control --- ----- 2.15 15.00 0.32 2.47 

Louvers 1 1.35 1.35 15.00 0.20 1.55 

Thermal Switches 4 0.20 0.80 15.00 0.12 0.92 
NASA/TM—2011-216971 41

4.7.4 Thermal Trades

No additional trades were run in this design session. The thermal system was designed to match the 

requirements.

4.7.5 Thermal Analytical Methods

The radiator panel area has been modeled along with a rough estimate of its mass. The model was 

based on a 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. Worst-case thermal 

environment assumptions were used to size the radiator. See Table 4.10 for thermal environment constant 

assumption.

TABLE 4.10.—THERMAL ENVIRONMENT

CONSTANT ASSUMPTIONS

Variable Value

Radiator solar absorptivity.....................................0.14

Radiator emissivity ...............................................0.84

Radiator sun angle ................................................. 90°

Radiator operating temperature ..........................320 K

S/C radiator dissipation power ..........................250 W

Power requirements and mass have been modeled. This modeling included propellant tank MLI and

heaters and propellant line insulation and heaters. Worst case thermal environment assumptions were used 

to calculate the heat loss (Figure 4.13). 

The model was based on a first principles analysis of the radiative heat transfer from the tanks and 

propellant lines to space. The heat loss through the insulation set the power requirement for the tank and 

line heaters. See Table 4.11 for thermal system details.

Figure 4.13.—Insulation mass versus number of layers of insulation.
NASA/TM—2011-216971 42

TABLE 4.11—THERMAL SYSTEM DETAILS

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 and power level.................................1.02 kg at 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 and power..........................0.143 kg/m at up to 39 W/m

Line foam insulation density....................................................................................56 kg/m3

The mass of the S/C MLI on the engine bulkhead was modeled to determine the mass of the 

insulation and heat loss (Table 4.12). The model was based on a first principles analysis of the heat 

transfer from the S/C through the insulation to space. The near Earth thermal environment was used to 

size the insulation for the maximum expected operating heat load during flight.

TABLE 4.12.—S/C MLI 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) ......................................................... 25

MLI layer spacing (di) ................................................................ 1.0 mm

S/C radius (rsc)........................................................................... 1.145 m

Mass estimates for the ATCS system have been completed. The components of the system included: 

Cold plates and heat pipes (for cooling plate assumptions see Table 4.13). The model was based on a first 

principle 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. 

TABLE 4.13.—COOLING PLATE ASSUMPTIONS

Variable Value

Cooling plate and lines material ..........................................................Al

Cooling plate and lines material density .............................. 2,770 kg/m3

Number of cooling plates........................................................................2

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

For more detailed information on the thermal analysis a summary white paper titled “Preliminary 

Thermal System Sizing,” was produced. This paper is presently under publication as a NASA Contractor 

Report (CR). 
NASA/TM—2011-216971 43

4.7.6 Thermal Risk Inputs

The risks associated with the thermal system are based mainly on the failure of a component or 

multiple components of the system. The majority of the system operation is passive and therefore has a 

fairly high reliability. Some of the major failure mechanisms are listed below.

• Heat pipe failure. This can be due to cracking due to thermal stresses, micro-meteor impact or 

design defect. This likelihood of this type of failure is low. The impact of this failure would be a 

loss of all or a portion of the S/C’s capability. 

• Heater system failure. This would most likely be due to wire breakage or a controller failure. The 

likelihood of this type of failure is low. The impact of this failure would be a loss of certain 

components or propulsion capability once the vehicle is exposed to an extended period of cold 

• Radiator louver failure. The thermal controller on the system can fail due to an electronics failure 

or power failure. Subsequently this will cause a failure of the radiator louvers reducing the 

effectiveness of the radiator and limiting control of the thermal system. The louvers can also 

experience mechanical failure causing them to be held in a fixed position or limiting their range 

of motion. 

To improve the reliability of the system and compensate for the identified failure risks, the following 

system design changes can be made.

• Redundant heat pips can be utilized for each cold plate. The heat pipes can be individually run to 

the radiator to provide independent cooling paths. The radiator can be separated into two 

independent units providing additional redundancy.

• Redundant heating system controllers can be utilized. The heaters can be wired individually so 

that a single heater failure does not bring down any additional heaters. Additional insulation can 

be added to the S/C to insure that the interior components do not drop below their desired 

minimum temperature based on a know shadow period of operation. 

• The radiator louvers can be designed to fail opened to a specified angle. This will enable the 

radiator to continue to operate, although not optimally, for the remainder of the mission. 

5.0 Cost, Risk and Reliability

5.1 Costing: Six ASRG Configuration

The following items represent the assumptions in the costing analysis of the six ASRG REP S/C

design. S/C costs reflect 50 percent confidence level. The ASRG is assumed to be flight ready by its own 

development project. The S/C fee is assumed at 10 percent and is not applied to science instruments 

(assumed to be furnished equipment). The NASA project office and technical oversight is based on 

5 percent of all other costs. The costing for Phase A is based on 5 percent of S/C costs. The Launch 

services cost is based on guidance from 2003 NF AO. The 25 percent reserves are not applied to Launch 

Services or RPS costs per 2003 NF AO. Table 5.1 shows the estimations for the REP S/C and science 

instruments life cycle costs (LCC).

TABLE 5.1.—REP S/C AND SCIENCE INSTRUMENTS LCC

REP Thereus—NF Mission FY08 $M

NASA project office/technical oversight 31

Phase A 17

S/C with science instruments* 315

S/C prime contractor fee (10%) 28

Launch services 172

Mission operations 94

Reserves (25%) 98

Life cycle cost 756
NASA/TM—2011-216971 44

Table 5.2 shows the costing per work breakdown structure (WBS) line items in the REP S/C MEL in 

FY08 $M.

TABLE 5.2.—REP S/C COST PER MEL LINE ITEM IN FY08 $M

WBS 

element

Element name DDT&E 

total

Flight 

hardware

S/C

total

0.1.1 Science Payload 22.7 12.3 35.0

01.1.1.a LORRI 6.0 2.6 8.6

01.1.1.c LIDAR 4.8 2.1 6.9

01.1.1.d NGIMS

01.1.2.a WAC 1.2 0.5 1.7

01.1.2.b NIMS 10.7 7.1 17.8

01.2.1 AD&C 9.0 7.6 16.6

01.2.1.a.a Sun Sensors 1.4 3.0 4.5

01.2.1.a.b Reaction Wheels 1.0 0.8 1.8

01.2.1.a.c Star Trackers 0.9 1.8 2.7

01.2.1.a.d IMU 5.7 1.9 7.6

01.2.2 C&DH 11.8 3.7 15.5

01.2.2.a.a Flight Computer 2.6 3.0 5.5

01.2.2.a.c Data Interface Unit 0.3 0.3 0.6

01.2.2.a.e Operations Recorder 0.1 0.1 0.2

01.2.2.a.f Command and Control Harness (data) 3.1 0.3 3.4

Flight Software/Firmware 5.8 5.8

01.2.3 Communications and Tracking 8.9 4.6 13.5

01.2.3.a X/Ka HGA

01.2.3.a.a Transmitter/Receiver 3.0 1.2 4.2

01.2.3.a.b Power Amp 1.4 1.1 2.5

01.2.3.a.d Antenna 2.3 1.6 3.9

01.2.3.a.h Cabling 0.6 0.2 0.8

01.2.3.c.a Coaxial Cable 1.7 0.4 2.1

01.2.4 Electrical Power Subsystem 3.8 125.4 129.2

01.2.4.a.a RPS Main System 123.4 123.4

01.2.4.b PMAD 2.7 1.1 3.8

01.2.4.c Power Cable and Harness Subsystem (C and HS) 1.1 0.9 2.0

01.2.5 Thermal Control (Non-Propellant) 3.6 1.0 4.7

01.2.5.a Active thermal control 0.5 0.8 1.2

01.2.5.b Passive and semi-passive thermal control 3.2 0.3 3.4

01.2.6 Propulsion 10.0 6.8 16.8

01.2.6.a.a Primary EP thrusters 0.8 1.2 2.0

01.2.6.b.a Xe propellant tank(s) 1.4 0.7 2.1

01.2.6.b Balance of propellant management system 2.0 0.7 2.7

01.2.6.c.a PPU mass 2.1 2.8 4.9

01.2.6.d RCS

01.2.6.d.a RCS tank subassembly 0.1 0.0 0.2

01.2.6.d.b RCS propellant management subassembly 1.7 0.8 2.5

01.2.6.d.c RCS thruster subassembly 1.9 0.5 2.4

01.2.8 Structures and mechanisms 4.7 4.0 8.7

01.2.8.a Structures 4.5 3.8 8.3

01.2.8.b.e S/C adapter 0.2 0.2 0.4

Subtotal 74.6 165.4 240.1
NASA/TM—2011-216971 45

TABLE 5.2.—REP S/C COST PER MEL LINE ITEM IN FY08 $M

WBS 

element

Element name DDT&E 

total

Flight 

hardware

S/C

total

Systems Integration 46.3 29.1 75.3

Integration, Assembly and Check Out 3.8 5.2 9.0

System Test Operations 4.4 4.4

Ground Support Equipment 7.9 7.9

System Engineering and Integration 14.8 16.4 31.1

Project Management 8.4 7.5 15.9

Launch Operations and Orbital Support 7.0 7.0

Total Prime Cost 120.9 194.5 315.4

5.2 Reliability

5.2.1 Reliability Methodology 

A first-order reliability analysis was performed based on limited information available during the 

conceptual design phase. Hardware elements, which have major impact on mission success (such as 

propulsion and power systems), were given the most attention. Many of the component failure rates 

utilized were assumed using engineering judgment due to a lack of knowledge—partly because new 

technologies are used with unknown reliability parameters and partly because relevant reliability 

databases are not generally accessible. Hence, the reliability numbers cited below are considered to be 

ballpark-accurate only and subject to considerable revision pending detailed analyses.

Figure 5.1 shows the general methodology used to estimate the reliability of the REP S/C. To cope 

with the limited study duration, a simple spreadsheet reliability model was developed that contained 

numerous and/or gates that mimicked the behavior of a full-blown fault tree analysis. This tool affords a 

convenient technique to generate approximate reliability values relatively quickly for alternative design 

options. It is a deterministic methodology, however, and therefore yields a single reliability value that is 

subject to considerable uncertainty. The systems uncertainty analysis (SUA) tool was therefore invoked to 

assess the impact of the uncertainties on the mission reliability as well as to identify the most important 

uncertainties.

Figure 5.1.—Reliability prediction methodology.
NASA/TM—2011-216971 46

5.2.2 Reliability Assumptions

The baseline S/C propulsion systems were assumed to consist of

• Four extended-life Hall-effect thrusters (three of the thrusters are activated sequentially one-at-atime—3 yr = 26298 hr duty for each—for 9-yr of S/C thrusting and the fourth thruster is a 

standby spare). 

• An RCS system consisting of eight MMH/NTO thrusters placed around the S/C body (the RCS 

provides some additional ∆V capability for minor orbit adjustments).

An important uncertainty is the failure rate of the advanced Hall-effect thrusters. Only two data points 

were available to estimate this failure rate—successful life tests of 16,265 and 30,352 hr on key elements 

of current-technology thrusters applicable to life-limiting elements of advanced Hall Thrusters. Two 

procedures are available for estimating mean-time-to-failure (MTTF) of hardware that has demonstrated 

lifetimes without failure

• Using “Calculating MTTF When You Have Zero Failures” by Relex Software, MTTF was 

estimated as 1.4427×30352 hr = 43789 hr, which results in an estimated thruster failure rate

λ = 1/MTTF = 2.284×10–5

/hr; 

• A second “rule-of-thumb” is the 1/3 rule, in which estimated MTTF = demonstrated life divided 

by 1/3 = 30352/0.33333 = 91056 hr, resulting in an estimated failure rate of 1.098×10–5

/hr. 

Feedback from thruster designers indicated the first estimate is more likely at this point.

The ASRG power system is the second critical element of the S/C. The baseline configuration 

assumed that six ASRG units would operate for the entire 10-yr (= 87,660 hr) mission to provide 

electrical power and that one of these units could fail without causing complete mission failure (although 

operational compromises would certainly be required). Estimates of an ASRG MTTF was reverseengineered from reliability results performed for a February 2007 ASRG Engineering Unit Final Design 

Review (FDR). For the FDR study option which assumed the use of high-strength components, 3 yr in 

fueled storage, and 14 yr operating life, an ASRG mission reliability of 0.9906 had been estimated.

Assuming the MTTF for the storage portion of the mission was 10 times that of the operational phase, an 

ASRG operational MTTF of 13.900,000 hr was calculated (resulting in an estimated failure rate of 

7.1942×10–8

/hr), which is significantly longer than values used in earlier analyses.

Additional hardware assumption changes made from phase 1 of this study included

(1) Propulsion—Decreased the common cause factor (CCF) value from 0.10 to 0.05 by assuming that 

extensive ground testing would precede flight.

(2) RCS—Reduced duty life (once every 2 wk for 20 min rather than the original 1 min every 6 hr);

(3) Avionics—Two redundant pairs of star trackers rather than a single pair.

(4) Engine gimbals—Replaced the initial guess of 1.0×10–6

/demand with a Cassini gimbal failure 

rate 2 orders of magnitude less.

5.2.3 Reliability Results

Reliability analyses were performed for the 10-yr mission starting with the baseline configuration 

with three of the thruster strings (all the elements downstream of the Xe propellant tank and regulator—

thruster, flow controller, and PPU) activated serially for 9 yr of EP thrusting. As the sensitivity results in 

Figure 5.3 indicate, the Hall thruster system configuration has the biggest impact on overall S/C

reliability. Thruster system and overall S/C reliabilities determined for different numbers of thruster 

spares and values of thruster MTTF are summarized in Table 5.3.
NASA/TM—2011-216971 47

TABLE 5.3.—RELIABILITY CASES FOR VARYING NUMBER OF THRUSTERS AND MTTFS 

Total no. of Hall 

thrusters*

No. of Hall 

thruster spares

Assumed Hall 

thruster MTTF 

(hr)

Thruster strings 

system reliability

Overall s/c 

reliability

4 1 43789 0.807 0.754

5 2 43789 0.930 0.869

6 3 43789 0.953 0.891

4 1 91056 0.932 0.871

5 2 91056 0.966 0.903

6 3 91056 0.970 0.906

* Total minus spares = 3 operating thrusters (serially one at a time)

The highlighted result shows the value of adding an additional thruster string to the configuration 

(adding two additional spares has little impact) and technology efforts to improve thruster MTTF to the 

higher of the estimated values.

Table 5.4 summarizes the reliability contributions of the various S/C systems to the overall S/C

reliability (exclusive of the science payload) for the highlighted configuration. 

TABLE 5.4.—RELIABILITY SUMMARY FOR SIX REP ASRG’S 

AND FIVE HIGH-MTTF THRUSTERS CONFIGURATION

Contribution

Propulsion system (including RCS) ................................................................. 41.3%

Electrical power ................................................................................................. 0.9%

AD&C................................................................................................................ 8.1%

Communication................................................................................................ 18.3%

Thermal control................................................................................................ 31.0%

Structure and mechanical systems ..................................................................... 0.4%

Science (not modeled).......................................................................................-------

Mission reliability ............................................................................................. 0.903

The impact of four major failure rate uncertainties (thruster, Xe tank, RCS, ASRG) on mission 

reliability is displayed in Figure 5.2 in terms of the probability and cumulative density functions (pdf and 

cdf). Each uncertainty was assumed to be normally distributed with a 0.50 coefficient of variation. These 

uncertainties have a large impact on predicted mission reliability: the 5 to 95 percentile range of mission 

success is 0.88 to 0.93 out of 1.0.

Figure 5.2.—Probabilistic S/C reliability (capturing major uncertainties).
NASA/TM—2011-216971 48

Figure 5.3.—Probabilistic sensitivity factors.

The probabilistic sensitivity factors are shown in Figure 5.3. The longer bars identify the more 

important uncertainties. In this case, the thruster failure rate (Pf = 1.0982×10–5

/hr) uncertainty is dominant. 

5.2.4 Interpretation of the Reliability Results 

The low mission reliability estimate is somewhat discouraging at first glance. Furthermore, a large 

improvement is not possible with just a single added redundancy since the unreliability is spread 

relatively evenly over many components. This suggests that reliability improvement would come from 

many design improvements rather than just one. 

It should be noted, however, that the modeling fidelity for this design exercise was quite low. Many 

failure rates were merely educated guesses due to the rapid pace of the analysis and especially efforts to 

come up with reasonable estimates for the hardware elements driving the overall S/C reliability. Hence, it 

is not known whether the low values reflect the reality of substantial risks in long duration deep space 

missions or are artificially low because of the modeling fidelity. A more in-depth reliability analysis is 

required to resolve this uncertainty.

6.0 Trades

The original baseline at the start of the COMPASS session was an eight ASRG configuration for the 

bus. Upon costing, the number of ASRGs was reduced to six, and that design became the new baseline.

Both designs are reported here in the trade section.

6.1 Case 1—REP Centaur Orbiter With Six ASRGs—Baseline

This design became the baseline of the mission since it accomplished the science goals and fit inside 

the NF cost box. The details of this design trade are documented in this report in detail.

6.2 Case 2—REP Centaur Orbiter With Eight ASRGs

Initially, the REP Centaur orbiter design contained eight ASRGs. This design was able to deliver the 

payload required in a reasonable amount of trip time. It did not, however, fit inside the NF cost cap.

Therefore, the first trade performed was to reduce the number of ASRGs down to six, and adjust the 

mission in order to fly to the centaur body with reduced power available to the thrusters. In order to 

reduce the S/C mass, one of the science instruments (NGMIS) was dropped off of the science payload 

(see Section 2.7). Note that New THEMIS was put back into the science instrument listing at the final 

design.
NASA/TM—2011-216971 49

6.2.1 Eight ASRG REP Centaur Orbiter Summary

The eight ASRG case is still a NF Class Mission, to be launched on an Atlas 551 using a Star 48 solid 

motor.

• 8.5 yr trip time to Centaur Thereus

– Flyby a Trojan in 2 yr

• 1 yr Centaur Science Period

– Science during Centaur perihelion for maximum activity

 Non-Perihelion targeting possible if active instruments used

– 55 kg/60 W of passive science instruments

– 6 kbps with DSN

• Eight ASRG

– 1040 W EOL (10 yr) power output

– Less plutonium than NH

• 3+1 advanced, long life Hall thrusters

– Thruster throughput >4 times current flight systems

• Relatively high power communications system

– Using EP power when coasting

– Could allow ‘high’ data rates, less DSN time, less cost

• Avionics, GN&C, hydrazine propulsion, thermal systems similar to other Deep Space Missions

Figure 6.1 shows the eight ASRG configuration of the Centaur orbiter from the beginning of this 

study. This was originally the baseline but was supplanted by the six ASRG case.

Figure 6.1.—Concept design using eight ASRGs.
NASA/TM—2011-216971 50

6.2.1.1 Eight ASRG Mission Trajectory—Thereus

After the REP S/C bottoms-up analysis and sizing, the amount of payload delivered to Bienor was 

found to be insufficient to accommodate the baseline S/C design and science instruments mass. Therefore, 

analysis was performed to locate another centaur body and trajectory that would place the REP S/C and 

science instruments at the Centaur before perihelion for science requirements, and allow enough delivered 

mass to accommodate the REP S/C.

The Centaur Thereus was chosen to fit those parameters. The baseline trajectory parameters for the

Thereus trajectory were

• Launch Mass (Mo): 1296.6 kg

• C3: 95.33 km2/s2

• Launch Date: November 6, 2024

• Fly-by Date: July 29, 2026

• Arrival Date: May 8, 2033

• Power: 750 W

• Isp: 2000 s

• ∆V: 8.59 km/s

• Propellant (Mp): 469.73 kg

• Mo – Mp: 836.87 kg

• Fly-by Target: Prylis

6.2.1.2 Mission Trajectory Details: Eight ASRG Case

Figure 6.2 shows the baseline trajectory to Thereus for the REP Science Orbiter using eight ASRG 

power levels (750 W). Earth’s orbit is in blue in the center, Thereus’s is in the aqua color on the outside. 

The REP trajectory is in red, taking it through the Trojan asteroid belt shown in green. This baseline 

trajectory is for the thruster operating at 2000 Isp and a power level of 750 W.

Figure 6.2.—Trajectory to Thereus.
NASA/TM—2011-216971 51

Figure 6.3.—Launch, propellant and delivered mass as a function of Isp.

Figure 6.4.—Delivered Mass as a function of Isp.

The Centaur Thereus is a more difficult target to hit than Bienor. Thereus is a difficult target to reach 

in case of a failure in an RPS for baseline missions beyond 8 yr. Should there be an RPS failure inside the 

8.5-yr mission lifetime, the possibility of reaching Thereus becomes more difficult. There might not be 

the propellant budget in the current design to accommodate this sort of failure. Figure 6.3 shows the 

relationship of mass and Isp for the Thereus trajectory for differing power levels. Bienor showed 

approximately linear relationship of mass versus trip time during period of interest.

Analysis was performed on the sensitivity of the trajectory to Isp. Figure 6.3 shows that the launch 

mass and propellant mass are greatly affected by changed in Isp. 

However, Figure 6.4 shows that net delivered mass was very insensitive to Isp changes, varying only 

about 6 kg through a variation of 400 s in Isp.
NASA/TM—2011-216971 52

6.2.2 Eight ASRG System Summary Masses (MEL and System Sheet)

The following table is the system summary of the eight ASRG design configuration.

The MEL (Table 6.1) captures the bottoms-up estimation of CBE and growth percentage line item by 

item from the subsystem engineer.

The system integration in Table 6.2 summarizes those total masses, CBE and total mass after applied 

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

for system level growth. The total additional growth amounted to 8.8 percent (51 kg). The bottoms-up

average growth on the dry mass of the REP S/C totaled 21.2 percent (122 kg). This was for the dry mass 

of 698.8 kg. With the additional system level growth, the total dry mass is 750 kg. Given the available 

launch performance of 1296.6 kg to the C3 targets, and the 10 kg S/C adaptor taken out of that, the 

remaining mass available to the REP S/C is 1286.6 kg. The total wet mass, with 30 percent total growth 

applied on top of it yields a total wet mass of the REP S/C of 1278 kg, which fits in with an 8.5 kg 

margin.

TABLE 6.1.—EIGHT ASRG CASE TOP LEVEL MEL

Description 

REP Centaur Mission 

(September 4, 2007))

CBE mass

(kg)

Growth

(%)

Growth

(kg)

Total mass

(kg)

REP Spacecraft (Payload and Stage) 1105.19 11.07 122.36 1227.55 

Science Payload 42.42 30.00 12.73 55.15 

REP Bus 1062.77 10.32 109.64 1172.41 

AD&C 13.52 20.00 2.70 16.22 

C&DH 33.30 34.26 11.41 44.71 

Communications and Tracking 39.00 34.10 13.30 52.30 

Electrical power subsystem 234.76 16.56 38.88 273.64 

Thermal control (non-propellant) 36.49 18.00 6.57 43.06 

Propulsion 85.98 25.67 22.07 108.05 

Propellant 528.78 0.00 0.00 528.78 

Structures and mechanisms 90.94 16.17 14.70 105.65 

TABLE 6.2.—EIGHT ASRG REP CONFIGURATION SYSTEM SUMMARY

REP Spacecraft Master Equipment List Rack-up (Mass) COMPASS REP Design 

WBS Main subsystems CBE Mass 

(kg) 

Growth 

(kg) 

Total mass 

(kg) 

Aggregate 

growth 

(%) 

01 REP Spacecraft (payload and bus) 1105.2 122.4 1227.6 -----

01.1 Science Payload 42.4 12.7 55.1 30.0 

01.2 REP Bus 1062.8 109.6 1172.4 

01.2.1 AD&C 13.5 5.4 16.2 40.0 

01.2.2 C&DH 33.3 11.4 44.7 34.3 

01.2.3 Communications and Tracking 39.0 13.3 52.3 34.1 

01.2.4 Electric Power 234.8 38.9 273.6 16.6 

01.2.5 Thermal Control 36.5 6.6 43.1 18.0 

01.2.6 Propulsion 86.0 22.1 108.1 25.7 

01.2.7 Propellant 528.8 ------ -------- -----

01.2.8 Structures and Mechanisms 90.9 14.7 105.6 16.2 

Estimated REP S/C dry mass 576 122 698.8 21.2 

Estimated REP S/C wet mass 1105 122 1227.6 

System Level Growth Calculations Total growth 

Desired System Level Growth 576 173 749.3 30.0 

Additional Growth (carried at system level) ----- 51 -------- 8.8 

Total Wet Mass with Growth 1105 173 1278.1 -----

Available launch performance to C3 (kg) 1286.6 

Launch margin available (kg) 8.5 
NASA/TM—2011-216971 53

7.0 Challenges, Lessons Learned, and Areas For Future Study

In order to determine what are the key parameters for EP devices to perform these REP missions a 

design study was completed to design an REP S/C to orbit a Centaur in a NF cost cap. The design shows 

that an orbiter using several long lived (~200 kg Xe throughput), low power (~700 W) Hall thrusters 

teamed with six (150 W each) ASRG can deliver 60 kg of science instruments to a Centaur in 10 yr 

within the NF cost cap. Optimal Isps for the Hall thrusters were found to be around 2000 s with thruster 

efficiencies over 40 percent. Not only can the REP S/C enable orbiting a Centaur (when compared to an 

all chemical mission only capable of flybys) but the additional power from the REP system can be reused 

to enhance science and simplify communications.

Key to the feasibility for REP missions are long life, low power EP devices, low mass RPS and light 

S/C components. Performance of the REP mission could be improved by increasing ASRG power density 

and increasing thruster lifetime and efficiency.

7.1 Six ASRG REP Centaur Orbiter Summary

7.1.1 Potential Mass Reductions 

• Improved ASRG (greater power density)

– Allows for elimination of one to two ASRGs from the baseline design

• ASRG mounting methods

– Trade of structure and thermal radiative coupling (effective SRG power output)

• Optimize structures, more composites

• Propulsion

– Higher EP Isp is shown to not be an important driver

– Efficiency and longer life thrusters are a major driver

• Communications

– High data rates cost and mission selling rate: reduced data rate not recommended

– Lighter antennas designs are recommended

7.1.2 Areas for Future Work

Further work is needed to trade ASRG mounting methods, optimizing structures (utilize more 

composites), and optimizing communications performance/mass based on the additional power available 

from the ASRG system.

Fix the issues open in the current design.

• Exploring Centaur Design Space

– Propulsion: Other Hall, ion options

– Power: Other power levels, power supplies

– Other targets

• Flagship and Discovery designs

• Inputs to EP Technology Development Program

NASA/TM—2011-216971 55

Appendix A.—Acronyms and Abbreviations

ACS Attitude Control System

AD&C Attitude, Determination and

Control

AIAA American Institute for Aeronautics 

and Astronautics

Al aluminum

ANSI American National Standards 

Institute

AO Announcement of Opportunity

APL Applied Physics Laboratory

ASRG Advanced Stirling Radioisotope 

Generators

BAE British Aerospace

BOL beginning of life

C&DH Command and Data Handling

CAD computer aided design

CBE current best estimate

CCF common cause factor 

COM center of mass

Comm Communications

COMPASS COllaborative Modeling and 

Parametric Assessment of Space 

Systems

CONOPS Concept of Operations

COPV Composite Overwrapped Pressure 

Vessel

COTS commercial-off-the-shelf

DPU dual processing unit

DSN Deep Space Network

ELV Expendable Launch Vehicle

EOL end of life

EOM end of mission

EP Electric Propulsion

FEA finite element analysis 

FOM figure of merit

FY fiscal year

GLIDE GLobal Integrated Design 

Environment

GN&C Guidance, Navigation and Control

GPHS general purpose heat source

GRC NASA Glenn Research Center

He helium

HGA high gain antenna

HiVHAC High Voltage Hall Accelerator

IEM integrated electronics module

IMU Inertial Measurement Unit

Isp specific impulse

JPL NASA Jet Propulsion Laboratory

KBO Kuiper Belt Objects

LCC life cycle costs

Li lithium

LIDAR Laser Detection and Ranging

LORRI Long Range Reconnaissance 

Imager

MEL Master Equipment List

MGA Mass Growth Allowance

MLI multilayer insulation

MMH/NTO monomethyl hydrazine and 

nitrogen tetroxide bipropellant 

system

MO Mission of Opportunity

MTTF mean-time-to-failure

NAC narrow angle camera

NASA National Aeronautics and Space 

Administration

NG SIRU Northrop Grumman Scalable

Inertial Reference Unit

NGIMS Neutral Gas and Ion Mass 

Spectrometer

NH New Horizons

NIMS Near Infrared Mapping 

Spectrometer

NIST National Institute of Standards and 

Technology

OSS Office of Space Science

OTS off-the-shelf

PEL Power Equipment List

PMAD Power Management and 

Distribution

PMS Propellant Management System

PPU Power Processing Units

RBI Regulator/Bus Protection 
NASA/TM—2011-216971 56

RCS Reaction Control System

REP Radioisotope Electric Propulsion 

RF radio frequency

RPS Radioisotope Power Systems

RSEN Reduced State Encounter 

Navigation 

S/C spacecraft

SEAKR SEAKR Engineering, Inc.

SEP Solar Electric Propulsion

SEU single event upset

SRG Stirling Radioisotope Generator

SSR solid state recorder

SUA Systems Uncertainty Analysis

TBD to be determined

THEMIS Thermal Emission Imaging 

System

Ti titanium

TT&C Telemetry, Tracking and 

Command

TWTA Traveling Wave Tube Amplifier

USO ultra-stable oscillator

WAC wide angle camera

WBS work breakdown structure

WGA weight growth allowance

WGS weight growth schedule

Xe xenon
NASA/TM—2011-216971 57

Appendix B.—Six ASRG Design Rendered Drawings

Figure B.1 shows the dynamic envelope of the REP S/C and Star 48 motor inside the Delta IV fairing.

Figure B.2 shows the rendered image of the REP Stage with the science payload side shown. Figure B.3

shows the rendered image of the REP Stage with the Hall Thruster side shown.

Figure B.1.—REP S/C inside Delta IV Fairing.

Figure B.2.—REP Stage after Star Motor separation.
NASA/TM—2011-216971 58

Figure B.3.—Thruster face of the REP S/C.
NASA/TM—2011-216971 59

Appendix C.—Eight ASRG Design Rendered Drawings

Figure C.1 shows the dynamic envelope of the REP S/C and Star 48 motor inside the Atlas 551 

fairing. Figure C.2 shows the dimensions of the Eight ASRG Case. Figure C.3 shows the REP S/C face 

with the four hall thrusters pointing out of the power and propulsion deck. Figure C.4 displays the science 

package face of the REP S/C. Figure C.5 shows the Avionics and flight deck of the REP S/C with all 

major components labeled. Figure C.6 presents the Power and Propulsion deck of the REP S/C with all 

major components labeled.

Figure C.1.—Dynamic envelope of the REP S/C inside the Atlas 551 Fairing.

Figure C.2.—Dimensions of the eight ASRG Case.
NASA/TM—2011-216971 60

Figure C.3.—EP Thruster face of the eight ASRG REP Design.

Figure C.4.—Science package face of the REP S/C.
NASA/TM—2011-216971 61

.

Figure C.5.—Avionics deck of the eight ASRG design.

Figure C.6.—Power and propulsion deck of the eight ASRG design.

NASA/TM—2011-216971 63

Appendix D.—Mission Analysis of Centaur Body Bienor

D.1 Initial Potential Centaur Target Trajectory Analysis

A Centaur class body, Bienor, was chosen as the first representative body for this Centaur mission.

Bienor (Asteroid 54598, Bienor, is a Centaur-type asteroid discovered on August 27, 2000, by the Deep 

Ecliptic Survey project at Cerro Tololo near Pedernales, Coquimbo, Chile. It has a period of 66 yr, 305 

days) is representative of this class of body. If possible, the mission will add on a Trojan asteroid body 

flyby. 

Several trades were run on the affect of thruster power for trip times to see the impact on net 

delivered mass. Performance is very sensitive to both trip time and available power. A trade was 

performed at Hall Effect Thruster performance fit (Isp may be too high) to analyze the affect of power on 

trip time and delivered mass. It was noted that optimal Isp rises quickly with available power and trip 

times greater than 9 yr.

Figure D.1 shows the net delivered mass to the Centaur, Bienor, as a function of trip time for three 

varying power levels of the REP S/C thruster system (500 to 750 W).

Figure D.2 shows net delivered mass to the Centaur Bienor as a function of thruster input power for a 

series of different trip time trajectories. The mission was baselined on a 750 W power system. The trades 

in the chart were made varying mission and launch date used a total power of 750 W.

Figure D.3 shows the launch mass (initial mass), propellant mass, and net delivered mass to Bienor1 

as a function of trip time at the 750 W ASRG configuration. The requirement on trip time is to arrive at 

Bienor1 in about a 10-yr total trip time. Therefore, the initial mass of the REP S/C can be no more than 

roughly 1100 kg at launch.

Figure D.1.—Bienor delivered mass versus trip time and thruster system.
NASA/TM—2011-216971 64

Figure D.2.—Bienor delivered mass versus thruster input.

Figure D.3.—Launch, propellant, and net masses to Bienor versus trip time for 750 W ASRG.

Figure D.4 shows the mission sensitivities (mass, thruster operation time, and coast period) to Isp of 

the Electric thrusters. This trade will help to decide which Isp to operate the electric thrusters at.

Figure D.5 is the baseline trajectory from Earth to Bienor1 for this mission, at an Isp of 2306 s and 

power of 750 W. The thruster Isp choice will be 1950 s, which will change all of the performance analysis 

in the previous four figures. The erosion of the thrusters at higher Isp are high, and the stability issues of 

the current at higher power caused the thruster Isp selection change for the mission.
NASA/TM—2011-216971 65

Figure D.4.—Mission sensitivities versus Isp.

Figure D.5.—Baseline Earth to Bienor.

NASA/TM—2011-216971 67

Appendix E.—Study Participants

Radioisotope Electric Propulsion Design Session

Subsystem Name Center Email

APL Lead Paul Ostdiek APL Paul.Ostdiek@jhuapl.edu

James Leary APL James.Leary@jhuapl.edu

Science Rob Gold APL Robert.Gold@jhuapl.edu

Science Carey Lisse APL Carey.Lisse@jhuapl.edu

Science Karl Hibbits APL Karl.Hibbitts@jhuapl.edu

Science Carey Lisse APL Carey.Lisse@jhuapl.edu

In-Space Program Len Dudzinski GRC Len.A.Dudzinski@nasa.gov

In-Space Program Scott Benson GRC Scott.W.Benson@nasa.gov

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

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

Launch Vehicle Integration TBD GRC TBD

Ground Systems GRC with APL support TBD

Mission John Dankanich GRC John.W.Dankanich@nasa.gov

Operations, GN&C Doug Fiehler GRC Douglas.I.Fiehler@nasa.gov

Structures and Mechanisms John Gyekenyesi GRC John.Z.Gyekenyesi@nasa.gov

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

Propulsion Jim Gilland GRC James.H.Gilland@nasa.gov

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

Power Paul Schmitz GRC Paul.C.Schmitz@nasa.gov

Command and Data Handling Jeff Juergens GRC Jeffrey.R.Juergens@nasa.gov

Communications O. Scott Sands GRC Obed.S.Scott@nasa.gov

Communications Bin Nguyen GRC Binh.V.Nguyen@nasa.gov

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

Avionics, Communications and 

Software T.C. Nguyen GRC Thanh.C.Nguyen@nasa.gov

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

Risk Anita Tenteris GRC Anita.D.Tenteris@nasa.gov

Reliability Bill Strack GRC bstrack@wowway.com

Reliability Joseph Hemminger GRC Joseph.A.Hemminger@grc.nasa.gov
NASA/TM—2011-216971 68

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14. ABSTRACT

Radioisotope Electric Propulsion (REP) has been shown in past studies to enable missions to outer planetary bodies including the orbiting of Centaur asteroids. Key to the 

feasibility for REP missions are long life, low power electric propulsion (EP) devices, low mass Radioisotope Power System (RPS) and light spacecraft (S/C) components. In 

order to determine the key parameters for EP devices to perform these REP missions a design study was completed to design an REP S/C to orbit a Centaur in a New Frontiers 

(NF) cost cap. The design shows that an orbiter using several long lived (~200 kg xenon (Xe) throughput), low power (~700 W) Hall thrusters teamed with six (150 W each) 

Advanced Stirling Radioisotope Generators (ASRG) can deliver 60 kg of science instruments to a Centaur in 10 yr within the NF cost cap. Optimal specific impulses (Isp) for 

the Hall thrusters were found to be around 2000 s with thruster efficiencies over 40 percent. Not only can the REP S/C enable orbiting a Centaur (when compared to an all 

chemical mission only capable of flybys) but the additional power from the REP system can be used to enhance science and simplify communications. The mission design 

detailed in this report is a Radioisotope Power System (RPS) powered EP science orbiter to the Centaur Thereus with arrival 10 yr after launch, ending in a 1 yr science 

mapping mission. Along the trajectory, approximately 1.5 yr into the mission, the REP S/C does a flyby of the Trojan asteroid Tlepolemus. The total ΔV of the trajectory is 8.9 

km/s. The REP S/C is delivered to orbit on an Atlas 551 class launch vehicle with a Star 48 B solid rocket stage.

15. SUBJECT TERMS

Centaur launch vehicle; Electric propulsion; Costs; Hall thrusters; Trajectories; Deep Space Network; Planets; Asteroids 

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