Steven R. Oleson and Melissa L. McGuire

Glenn Research Center, Cleveland, Ohio

COMPASS Final Report: 

Lunar Relay Satellite (LRS)

NASA/TM—2012-217140

April 2012

CD–2007–12
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Steven R. Oleson and Melissa L. McGuire

Glenn Research Center, Cleveland, Ohio

COMPASS Final Report: 

Lunar Relay Satellite (LRS)

NASA/TM—2012-217140

April 2012

CD–2007–12

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135
Available from

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subject to revision as analysis proceeds.
NASA/TM—2012-217140 iii

Contents

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

2.0 Study Background and Assumptions....................................................................................................1

2.1 Introduction ..................................................................................................................................1

2.1.1 Human Phase Relay (LAT 1 Baseline Concept)...............................................................3

2.1.2 Concept of Operations.......................................................................................................3

2.2 Design Assumptions and Requirements.......................................................................................3

2.3 Assumptions.................................................................................................................................3

2.3.1 System Level Requirements..............................................................................................3

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

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

2.5.1 Mission Requirements.......................................................................................................4

2.5.2 Launch Vehicle Data and Trades......................................................................................4

2.5.3 Launch Vehicle Performance and Cost.............................................................................5

2.5.4 Mission Assumptions........................................................................................................5

2.5.5 Mission Timeline ..............................................................................................................6

2.5.6 Mission Trades: Electric Propulsion (EP) Versus Chemical.............................................7

2.5.7 Mission Analytical Methods .............................................................................................7

2.5.8 Mission Recommendation.................................................................................................7

2.5.9 Launch Vehicle Recommendations...................................................................................8

2.6 Design Trade Space ....................................................................................................................10

2.6.1 One Versus Two DTE Antennas.....................................................................................10

2.6.2 One Versus Two Relay Antennas ...................................................................................10

2.6.3 Atomic Clocks.................................................................................................................11

3.0 Baseline Design ..................................................................................................................................11

3.1 Top Level Design (MEL and PEL) ............................................................................................11

3.1.1 Master Equipment List (MEL)........................................................................................12

3.1.2 Power Equipment List (PEL)..........................................................................................12

3.2 Design Concept Drawing and Description .................................................................................13

4.0 Subsystem Breakdown........................................................................................................................13

4.1 Communications.........................................................................................................................13

4.1.1 Communications Requirements.........................................................................................13

4.1.2 Communications Assumptions ..........................................................................................14

4.1.3 Communications Design and MEL....................................................................................14

4.1.4 Communications Trades..................................................................................................16

4.2 Avionics......................................................................................................................................16

4.2.1 Avionics Requirements...................................................................................................16

4.2.2 Avionics Assumptions ....................................................................................................16

4.2.3 Avionics Design and MEL..............................................................................................16

4.2.4 GN&C Design.................................................................................................................16

4.2.5 Avionics Risk Inputs.......................................................................................................17

4.3 Structures and Mechanisms........................................................................................................17

4.3.1 Structures and Mechanisms Requirements......................................................................17

4.3.2 Structures and Mechanisms Assumptions.......................................................................17

4.3.3 Structures and Mechanisms Design and MEL ................................................................17

4.3.4 Structures and Mechanisms Analytical Methods............................................................18

4.3.5 Structures and Mechanisms Risk Inputs .........................................................................18

4.3.6 Structures and Mechanisms Recommendation................................................................18

4.4 Electrical Power System.............................................................................................................19

4.4.1 Power Requirements .......................................................................................................19
NASA/TM—2012-217140 iv

4.4.2 Power Assumptions.........................................................................................................19

4.4.3 Power Design and MEL..................................................................................................19

4.4.4 Power Trades...................................................................................................................19

4.4.5 Power Analytical Methods..............................................................................................20

4.4.6 Power Recommendation .................................................................................................20

4.5 Propulsion and Propellant Management.....................................................................................20

4.5.1 Propulsion and Propellant Management Requirements..................................................20

4.5.2 Propulsion and Propellant Management Assumptions....................................................20

4.5.3 Propulsion and Propellant Management Design and MEL.............................................20

4.5.4 Propulsion and Propellant Management Trades..............................................................24

4.5.5 Propulsion and Propellant Management Analytical Methods.........................................24

4.5.6 Propulsion and Propellant Management Risk Inputs......................................................24

4.5.7 Propulsion and Propellant Management Recommendation ............................................24

4.6 Thermal Control .........................................................................................................................25

4.6.1 Thermal Requirements....................................................................................................25

4.6.2 Thermal Assumptions .....................................................................................................25

4.6.3 Thermal Design and MEL...............................................................................................25

4.6.4 Thermal Analytical Methods...........................................................................................26

4.6.5 Thermal Risk Inputs........................................................................................................27

4.6.6 Thermal Recommendation ..............................................................................................27

5.0 Cost and Risk......................................................................................................................................27

5.1 Costing........................................................................................................................................27

5.2 Risk Analysis and Reduction......................................................................................................27

5.2.1 Risk Assumptions............................................................................................................27

5.2.2 Risk Matrix and Risk List ...............................................................................................27

5.2.3 Risk Summary.................................................................................................................30

Appendix A.—Acronyms and Abbreviations.............................................................................................31

Appendix B.—Rendered Design Drawings................................................................................................33

Appendix C.—Study Participants...............................................................................................................35

List of Tables 

Table 1.—COMPASS Mass Growth Allowance Standard...........................................................................4

Table 2.—Launch Service Costs from KSC.................................................................................................4

Table 3.—LRS Mission Timeline and Burn Magnitudes .............................................................................7

Table 4.—Comparison of Direct Chemical and SEP Missions ....................................................................7

Table 5.—Top Level LRS MEL Without System Level Growth ...............................................................12

Table 6.—LRS Master Equipment List with System Level Growth Calculations.....................................12

Table 7.—LRS Power Equipment List .......................................................................................................13

Table 8.—LRS Communications System MEL Breakdown ......................................................................15

Table 9.—LRS Avionics Detailed MEL.....................................................................................................17

Table 10.—LRS Structures and Mechanical Systems Detailed MEL ........................................................18

Table 11.—LRS Power System Detailed MEL ..........................................................................................19

Table 12.—LRS Propulsion Hardware Detailed MEL ...............................................................................22

Table 13.—LRS Propellant Management Detailed MEL...........................................................................23

Table 14—LRS Propellant Detailed MEL..................................................................................................23

Table 15.—LRS Thermal Control System Detailed MEL..........................................................................26

Table 16.—LRS Risk Matrix......................................................................................................................28

Table 17.—LRS Risk List and Detailed Data Table...................................................................................28

Table 17.—Concluded. ...............................................................................................................................30
NASA/TM—2012-217140 v

List of Figures

Figure 1.—An artists rending of the LCT and LRS systems in use at the lunar South Pole. .......................2

Figure 2.—LRS concept design....................................................................................................................2

Figure 3—Direct injection—lunar inertial coordinate frame........................................................................6

Figure 4.—Comparison of direct injection chemical to SEP mission performance. ....................................8

Figure 5.—Atlas V performance curve for C3 = –2 km2

/sec2. ......................................................................9

Figure 6.—Atlas V 401 fairing dimensions with notional two LRS launch packaging................................9

Figure 7.—Falcon launch vehicle family....................................................................................................10

Figure 8.—LRS design with major subsystem components identified.......................................................11

Figure 9.—Viewing angle of Q-band antenna on LRS...............................................................................14

Figure 10.—Bottom view of main engines on LRS....................................................................................21

Figure 11.—Propulsion system schematic diagram....................................................................................22

Figure 12.—LRS thermal control system. ..................................................................................................26

Figure B.1.—LRS stowed configuration K/Ka band antenna face.............................................................33

Figure B.2.—LRS stowed configuration, solar array panel side. ...............................................................33

Figure B.3.—LRS deployed configuration, K/Ka band antenna face.........................................................34

Figure B.4.—LRS deployed configuration.................................................................................................34

NASA/TM—2012-217140 1

COMPASS Final Report: Lunar Relay Satellite (LRS)

Steven R. Oleson and Melissa L. McGuire

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

1.0 Executive Summary

The Lunar Relay Satellite (LRS) COllaborative Modeling and Parametric Assessment of Space 

Systems (COMPASS) session was tasked to design a satellite to orbit in an elliptical lunar polar orbit to 

provide relay communications between lunar South Pole assets and the Earth. The design included a 

complete master equipment list, power requirement list, configuration design, and brief risk assessment 

and cost analysis. 

The LRS is a half-TDRSS sized box spacecraft, which provides communications and navigation relay 

between lunar outposts (via Lunar Communications Terminals (LCT)) or Sortie parties (with user radios) 

and large ground antennas on Earth. The LRS consists of a spacecraft containing all the communications 

and avionics equipment designed by NASA Jet Propulsion Laboratory’s (JPL) Team X to perform the 

relay between lunar-based assets and the Earth. The satellite design is a standard box truss spacecraft 

design with a thermal control system, 1.7 m solar arrays for 1 kWe power, a 1 m diameter Ka/S band dish 

which provides relay communications with the LCT, and a Q-band dish for communications to/from the 

Earth based assets. While JPL’s Team X and Goddard Space Flight Center’s (GSFC) I M Design Center 

(IMDC) have completed two other LRS designs, this NASA Glenn Research Center (GRC) COMPASS

LRS design sits between them in terms of physical size and capabilities.

This LRS was used in conjunction with the LCT designed by COMPASS and documented in COMPASS 

final report CD-2007-11.

Figure 1 is a notional artist rending of the LCT and LRS systems in use at the lunar South Pole.

Figure 2 shows the LRS concept design.

2.0 Study Background and Assumptions

2.1 Introduction 

During crewed missions to the Moon, there is a need for redundant communications paths to Earth.

For early, shorter duration missions, direct to Earth (DTE) and relay provides dual paths. For missions 

longer than the 2-wk two LRS satellites are required for continuous coverage.

During the crewed phase, communication and navigation services are primarily provided via relay, in 

conjunction with a surface LCT, in support of the outpost. This LCT also provides outpost 

communications interconnectivity using 802-network technology.

In all cases, the nominal communications operations include all outpost vicinity communications 

routed through the LCT to other local users or to the relay. The LRS and LCT will provide these basic 

communication and navigation services:

• Forward command

• Return mission data and telemetry

• One- and two-way ranging and Doppler tracking 

• Beacon signals

• In situ routing

This design session concentrated on designing the LRS.
NASA/TM—2012-217140 2

Figure 1.—An artists rending of the LCT and LRS systems in use at the lunar South Pole.

Figure 2.—LRS concept design.
NASA/TM—2012-217140 3

2.1.1 Human Phase Relay (LAT 1 Baseline Concept)

For the Crewed Phase, Communication (Comm) and Navigation (Nav) services in the lunar vicinity 

will be provided primarily by LRS, backed up by Earth Based DTE antennas, which also provide service 

during Cis-lunar operations. The Surface Systems will provide wide area surface networking and surface 

links to the relay satellites.

2.1.2 Concept of Operations

Nominally, all communications during crew operations at the Moon will go through a relay. A single 

relay will provide about 60+ percent availability, or about 17 hr a day, from the baseline frozen orbit. 

When not in view of LRS or LCT, a user has the capability to communicate with Earth. (But for missions 

longer than 8 d, LRS is needed for daily Earth contact. LRS is also needed for redundant paths.) The LRS 

will broadcast a navigation signal and provide timing services. The LRS will have the capability to 

“locally” route communications in the lunar vicinity.

2.2 Design Assumptions and Requirements

2.3 Assumptions

Two LRS will be designed to provide Comm/Nav for piloted lunar missions, to be launched in 2018 

and 2020, operating for a 7-yr mission with a design ground rule of a 10-yr lifetime. 

2.3.1 System Level Requirements

The requirement for this design is to produce a satellite that meets the communications requirements.

The LRS provides communications to all lunar surface elements (south pole outpost site focused), but 

also global coverage as necessary, robotic, rover, hab, ISRU in-situ resource utilization (ISRU), Space 

Power Unit (SPU)/Makeup Power Unit (MPU), 500-km diameter circle (S band), 20-km Ka band.

Figures of Merit (FOM) for this study design were: Cost, Mass, Flexibility

2.4 Growth, Contingency and Margin Policy

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

for Mass Properties Control for Satellites, Missiles, and Launch Vehicles. Table 1 shows the growth 

percentage based on design maturity and design area. Percent Mass Growth divided by design maturity 

and subsystem.

Once the percent growth factors are applied on a subsystem per subsystem basis, the total system 

growth of the design is calculated. The COMPASS system desired total growth to be 30 percent, and an 

additional growth is carried at the system level in order to add up to a total system growth of 30 percent

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

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

From the Discovery Announcement of Opportunity: 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 

agreed-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.

Power Growth: The COMPASS team uses a 30 percent growth percentage on power requirements.
NASA/TM—2012-217140 4

TABLE 1.—COMPASS MASS GROWTH ALLOWANCE STANDARD

Code

Design maturity

(basis for mass 

determination)

Percent mass growth allowance

Electrical/electronic components

StructureThermal controlPropulsionBatteriesWire harnessesMechanismsInstrumentation

0 to 5 kg 5 to 15 kg >15 kg

E Estimated

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

L

Layout

(or major modification of 

existing hardware)

25 20 15 12 12 12 15 30 12 30

P

Pre-release drawings

(or minor modification of 

existing hardware)

20 15 10 8 8 8 10 25 8 25

C Released drawings

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

X

Existing hardware

(actual mass from another 

program)

3 3 3 2 2 2 3 3 2 3

A Actual mass

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

CFE Customer furnished 

equipment 0 0 0 0 0 0 0 0 0 0

2.5 Mission Description

The LRS mission launches on an Atlas V 501 launch vehicle to a C3 of approximately –2 km2/s2. This 

places the LRS on a direct trajectory to the Moon with a trip time of approximately 4 d. The LRS places 

itself into a lunar frozen orbit for long duration operation with minimal station keeping required. From 

this orbit, with a long dwell time over the lunar south pole, the LRS completes its primary mission 

providing communications coverage to assets on the lunar surface and relaying that information back to 

ground stations on the Earth. Two LRS satellites are delivered to the 12 hr frozen orbit, out of phase with 

each other by 180°. 

2.5.1 Mission Requirements

• Deliver sufficient mass to the lunar frozen orbit.

• Provide as much communications coverage as possible to the lunar South Pole.

• Remain in orbit for a sufficiently long period of time.

• Dispose of spacecraft safely.

2.5.2 Launch Vehicle Data and Trades

Table 2 lists the cost of launch vehicles available at the time of the LRS launch to lunar orbit. All 

available launch vehicles provide an excess of performance for the LRS mission.

TABLE 2.—LAUNCH SERVICE COSTS FROM KSC

Performance range Cost Launch vehicles

0 to 3,580 kg ~$175M Atlas V 401, Delta IV Medium

0 to 2,795 kg ~$185M Atlas V 501

4,711 to 5,400 kg ~$224M Delta IV Medium +
NASA/TM—2012-217140 5

2.5.3 Launch Vehicle Performance and Cost

Assumptions:

• Performance range based on a C3 = –2 km2

/s2 (lunar). 

• Used a launch date of January 1, 2018, and an eastern range.

• Per our PPBE07 assumptions, the prices reflect revised Launch Service Program (LSP) Pricing 

Strategy for the Evolved Expendable Launch Vehicles (EELV) Launch Service, which result in a 

slightly reduced amount of conservatism based on some recent acquisitions.

• All costs are estimated in real-year dollars based on current NASA Launch Services (NLS)

contract information. Prices provided are for a complete launch service, which includes nominal 

allocation for mission unique launch vehicle modifications/services, mission integration, launch 

site payload processing, range safety, and telemetry/communications.

• There is no launch delay penalty cost assumed in the budget.

• The launch service prices are estimates and are not to be considered commitments from the 

Launch Service Program.

• Launch Service Task Order (LSTO) is awarded to the Contractor that provides the best value in 

launch services to meet the Government's requirements based on technical capability/risk, 

reasonableness of proposed price, and past performance.

• Due to uncertainty with U.S. Air Force (AF) infrastructure cost allocations, NASA Headquarters 

(HQ) Programs should carry threat to cover resulting unexpected price fluctuations.

Cost/performance range difference associated with the 0 to 3580 kg and 0 to 2795 kg scenarios is 

associated with the use of a larger fairing.

2.5.4 Mission Assumptions

The ground rules for the mission design were based on work done and assumptions levied on the JPL 

and GSFC design sessions on alternate LRS systems.

• Launch vehicle: Atlas V 401

• Lunar orbit parameters:

○ Periapse altitude: 718 km

○ Apoapse altitude: 8090 km

○ Inclination: 57.7°

○ Argument of periapse: 90°

• Frozen orbit allows the relay satellite to maintain its apoapse in the lunar southern hemisphere 

without the need for station keeping

○ Semi-major axis, inclination and argument of perigee will vary but they remain bounded

Figure 3 shows the direct injection trajectory as displayed in the lunar inertial coordinate frame.
NASA/TM—2012-217140 6

Figure 3—Direct injection—lunar inertial coordinate frame.

2.5.5 Mission Timeline

Table 3 outlines the mission events, magnitude and duration of ∆V burns for the LRS mission. The 

ΔV table was adapted from GSFC IMDC presentations on their LRS design, with the addition of:

• 50 m/s for attitude control in lunar orbit

• 50 m/s for orbit maintenance in lunar orbit

• 80 m/s for disposal on lunar surface near perilune

Phasing maneuver ΔV provides for:

• Critical event coverage ~4 missions, 45° mean anomaly change

○ ~4 m/s per 45° mean anomaly change with 2 wk to execute

• Did not include 180° mean anomaly changes to avoid long Earth eclipses

○ ~18 m/s with 2 wk to execute

• More analysis needed to assess

○ Stability of the phasing orbits

○ Extra station keeping ΔV (if any)

○ Effect to operations during the phasing time period
NASA/TM—2012-217140 7

TABLE 3.—LRS MISSION TIMELINE AND BURN MAGNITUDES

Phase

no.

Phase name Main

∆V 

(m/s)

RCS

∆V 

(m/s)

Duration 

(hr)

Preburn

mass 

(kg)

Main prop

load

(kg)

Main burn

time

(min)

RCS load

(kg)

RCS burn

time

(min)

1

2

3

4

5

6

7

8

9

10

11

12

13

Launch from Earth

Checkout

Loiter to TLI window opening

TLI opening to ignition

TLI burn

Correct for ELV dispersions

Trans-Earth mid-course corrections

Trans-lunar coast

Lunar orbit capture burn 

lunar orbit adjustments 

phasing maneuvers

Lunar telecom orbit

Disposal

---

---

---

---

85

25

---

410

250

50

50

80

---

---

---

---

---

---

---

---

---

50

---

0.5

18.0

70.5

3.0

0.1

1.0

71.7

0.5

4.0

1.0

1.0

61320.0

1.0

1165

1165

1165

1165

1165

1165

1122

1110

1110

928

831

813

778

0.0

0.0

0.0

0.0

0.0

42.5

12.2

0.0

182.3

96.1

18.0

17.6

26.8

0.0

0.0

0.0

0.0

0.0

3.6

1.0

0.0

15.6

8.2

1.5

1.5

2.3

0.0

0.0

0.0

0.0

0.0

0.0

0.0

0.0

0.0

0.0

0.0

18.0

0.0

0.0

0.0

0.0

0.0

0.0

0.0

0.0

0.0

0.0

0.0

0.0

15.1

0.0

Total ∆V 950.0 50.0 Total burn

time 33.9 15.1

2.5.6 Mission Trades: Electric Propulsion (EP) Versus Chemical

The bottom line is that because of the excess of 60 percent launch margin using a chemical direct 

injection technique off of the available larger class launch vehicles for LRS, EP is not necessary for 

performance improvements.

Table 4 shows the comparison of the direct chemical injection mission versus a similar spacecraft 

configuration for the Solar Electric Propulsion (SEP) mission. Each delivers a similar final mass to the 

same lunar orbit. 

TABLE 4.—COMPARISON OF DIRECT CHEMICAL AND SEP MISSIONS

Direct Chemical Mission SEP Mission

Launch vehicle and C3 Atlas V 501 Launch to C3 of ~–2 km2/s2 Atlas V 501 Launch to C3 of ~–2 km2/s2

∆V (m/s) 900 1200

Trip time 4 to 5 d ~85 d including a lunar flyby

Final orbit Spacecraft placed in lunar frozen orbit Spacecraft placed in lunar frozen orbit

Delivered spacecraft dry mass ~1300 kg ~1300 kg

Propellant 735 kg 80 kg

Power required ~1.5 kWe ~6 kWe

Figure 4 further points out the cost of the SEP performance over the direct injection chemical mission 

due to increased trip time.

2.5.7 Mission Analytical Methods

• Analysis performed using an internally developed mission architecture tool

○ This tool recursively calculates propellant masses based on propulsion system parameters, 

inert masses, and ∆v schedule.

○ Can accommodate multiple vehicles and non-propulsive mass depletion

2.5.8 Mission Recommendation

• Because of the large launch vehicle, a significant amount of mass margin can be carried.

• A direct chemically propelled mission is the preferred architecture for this mission.

• This mission can meet all of the performance goals set forth.
NASA/TM—2012-217140 8

Figure 4.—Comparison of direct injection chemical to SEP mission 

performance.

2.5.9 Launch Vehicle Recommendations

LRS is ideally a Delta-II class payload, but Delta II’s future is uncertain. Cost of keeping the Delta II 

pad operational may offset the savings in the cost of a Delta II launch. Moving up in EELV class, the 

Atlas V provides ~66 percent mass margin for LRS (i.e., it could launch almost three LRS) with sufficient 

volume for a dual launch or secondaries. For lunar orbit [C3 = –2.0 km2

/s2], Atlas V 501 can launch 

2,795 kg, Atlas V 401 3,580 kg, and Atlas V 511 3,915 kg (Figure 5 was generated using KSC’s online 

ELV performance estimation tool on the Atlas class of Launch vehicles), compared to LRS mass of 

1124 kg with growth. LRS fits within a 4-m diameter payload envelope, so the 4-m fairing (Atlas V 401 

series) is sufficient. 

• Atlas V 401 and Delta IV Medium (EELV) provide excess capability for LRS 

○ Performance excess mitigates risk of low performing launch vehicle

○ Atlas V is a NASA certified, flight proven launch vehicle

○ Excess launch vehicle performance can be provided for secondary payload opportunities 

using the ESPA adapter

○ Spacecraft can provide accommodations (mass, volume, power, etc.) for science mission 

payloads

• Savings could also be realized by launching LRS 1 and 2 on a single launch vehicle. See Figure 6

for an illustration of how two LRS vehicles could fit in the Atlas V fairing.

At this time, the use of the excess capacity of the Atlas V class launch vehicle used in delivering the 

LRS as a cost saving solution has not been determined. It is offered up here as a suggestion.

0

500

1000

1500

2000

2500

3000

Direct,

Chemical,

Atlas V 501

Lunar Flyby,

Atlas V 501,

SEP

Mass, kg

0

30

60

90

Trip Time, days

Dry Mass Prop. Mass Trip Time
NASA/TM—2012-217140 9

C3 (km2/s2)

Figure 5.—Atlas V performance curve for C3 = –2 km2/sec2.

Figure 6.—Atlas V 401 fairing dimensions with notional 

two LRS launch packaging.
NASA/TM—2012-217140 10

Figure 7.—Falcon launch vehicle family.

Commercial alternative launch vehicles may be available in the LRS launch timeframe (2018). The 

Space-X Falcon 1 is too small, and the Falcon 5 will not be available, but Falcon 9 is appropriate, and at 

an advertised price of $50M, provides a 75 percent savings to LRS over the $175M Atlas V or Delta IV 

cost. Falcon 9 has the potential to provide a low cost launch option for LRS, should that service 

materialize. Figure 7 details the Falcon family of launch vehicles.

2.6 Design Trade Space

For the design of the LRS, the following major functional trades were identified.

2.6.1 One Versus Two DTE Antennas

• Previous (initial COMPASS payload design) had two 0.5-m Ka band DTE antennas to enable 

simultaneous Earth, Moon, Sun pointing

• Current design is one 0.5-m Ka band dish for DTE communications

• Impact: Reduces mass of one antenna and eliminates 3 dB loss in power due to Traveling Wave 

Tube Amplifier (TWTA)

2.6.2 One Versus Two Relay Antennas

• Previous: Relay link had two antennas: 1- by 1-m S band patch array + 0.5-m Ka band dish

• Current: Relay link has one 1-m dish with dual feed (S and Ka)
NASA/TM—2012-217140 11

• Impact 1: Reduces required surface area on s/c bottom, but we lose flexibility of pointing Ka and 

S band antennas independently

• Impact 2: Decreases Ka band footprint from 60- to 30-km radius

2.6.3 Atomic Clocks 

• Previous: None

• Current: 2 atomic clocks (1 + backup)

• Impact 1: Increase mass by 6 kg and DC power by 20 W

• Impact 2: Signal-to-noise (SN) signal pseudo noise (PN) code transmissions from LRS 2,3 and 

LCT are synchronous

3.0 Baseline Design 

The simplicity of this spacecraft design rests in a 110° joint, which enables the Earth antenna to view 

more than a hemisphere, so that when Earth is edge-on to the LRS orbit, oscillating from slightly above to 

slightly below 90° relative to the Earth antenna, the Earth antenna can look down over the solar arrays to 

see Earth without requiring the spacecraft to flip over. In this design, a spacecraft flip is required twice a 

lunar month, every 14 d. The solar arrays are mounted to the rear, Russian style, to avoid blocking the 

Earth antenna view. The arrays have two degrees of freedom rotation to track the Sun. The lunar antenna 

is gimbaled. See Figure 8 for a design drawing with main components labeled.

3.1 Top Level Design (MEL and PEL)

The bottoms-up current best estimate (CBE) wet mass of the LRS satellite was 1030.5 kg. See Table

5 for detail breakdown per subsystem. With growth factors added at appropriate levels for technology 

development in the subsystems, the total bottoms up wet mass for the LRS was 1120.6 kg. Total 

propellant was 467.3 kg. This leaves the total CBE dry mass at 563.1 kg and the total dry mass with 

growth at 653.3 kg without the additional system level growth from Section 3.1.1.

Figure 8.—LRS design with major subsystem components identified.
NASA/TM—2012-217140 12

TABLE 5.—TOP LEVEL LRS MEL WITHOUT SYSTEM LEVEL GROWTH

Description Quantity Unit

mass

(kg)

CBE 

mass

(kg)

Growth

()

Growth

(kg)

Total

mass

(kg)

Nominal

power

(W)

Peak

power

(W)

Lunar COMM Relay Satellite (LRS) ---- ---- 1030.5 8.7 90.1 1120.6 683.6 259.6

LRS-2 ---- ---- 1030.5 9 90.1 1120.6 683.6 259.6

Communications ---- ---- 78.4 3 2.4 80.8 494.0 0.0

Avionics ---- ---- 89.6 24 21.1 110.7 189.6 259.6

Structures & Mechanical Systems ---- ---- 180.5 14 25.1 205.6 0.0 0.0

Power System ---- ---- 72.6 31 22.2 94.8 0.0 0.0

Propulsion (Electric) ---- ---- 0.0 0 0.0 0.0 0.0 0.0

Propellant management (EP) ---- ---- 0.0 0 0.0 0.0 0.0 0.0

Propellant (EP) ---- ---- 0.0 0 0.0 0.0 0.0 0.0

Propulsion (Chemical) ---- ---- 21.8 8 1.7 23.5 0.0 0.0

Propellant management (Chemical) ---- ---- 72.9 15 10.9 83.9 0.0 0.0

Propellant (Chemical) ---- ---- 467.3 0 0.0 467.3 0.0 0.0

Thermal Control (Non-Propellant) ---- ---- 47.3 14 6.7 54.0 0.0 0.0

3.1.1 Master Equipment List (MEL)

Where the MEL (Table 6) captures the bottoms up estimation of CBE and growth percentage line 

item by item from the subsystem designer, Table 6 wraps up those total masses, CBE and total mass after 

applied growth percentage. The total growth from the bottoms up is 90.1 kg (16 percent of the dry mass).

In order to reach the system level growth mass of 30 percent (168.9 kg) an additional 78.8 kg is carried at 

the system level.

TABLE 6.—LRS MASTER EQUIPMENT LIST WITH SYSTEM LEVEL GROWTH CALCULATIONS

LRS MEL CBE 

(kg)

Growth 

()

Growth 

(kg)

Total Mass

(kg)

Percent of 

dry mass

Communications 78.4 3 2.4 80.8 12.4

Avionics & GN&C 89.6 24 21.1 110.7 16.9

Structures & Mechanisms 180.5 14 25.1 205.6 31.5

Power System 72.6 31 22.2 94.8 14.5

Propulsion (Chemical) 21.8 8 1.7 23.5 3.6

Propellant Management (Chemical) 72.9 15 10.9 83.9 12.8

Propellant (Chemical) 467.3 0 0.0 467.3 -----

Thermal Control 47.3 14 6.7 54.0 8.3

LRS dry mass without system growth 563.1 16.0 90.1 653.3

Desired system growth (contingency) 563.1 30 168.9 732.1

Added system level growth (contingency) 563.1 14.0 78.8 ------

LRS dry mass (with system growth) 563.1 30 168.9 732.1

Total useable propellant 453.0 ---- ---- ----

Total residuals + pressurant 14.3 ---- ---- ----

LRS inert mass (with system growth) 577.5 N/A 168.9 746.4

LRS total wet mass (with system growth) 1016.1 N/A 168.9 1185.1

3.1.2 Power Equipment List (PEL)

The power listing for nominal loads is 684 W (see Table 7). This includes both the communications 

packages and the avionics. At the time of the study, propulsion and thermal had not yet gotten their power 

requirements into the system design. Therefore, the power system included a 30 percent margin to 

account for these unknowns.
NASA/TM—2012-217140 13

TABLE 7.—LRS POWER EQUIPMENT LIST

LRS PEL Nominal power 

(W)

Peak power

(W)

Standby power 

(W)

LRS Subsystems

Communications 494 0 0

Avionics 189.6 259.6 24.5

Power System 0 0 0

Structures & Mechanisms 0 0 0

Propulsion 0 0 0

Thermal Control 0 0 3.4

Total LRS 684 260 28

3.2 Design Concept Drawing and Description

The LRS in this design is a half-TDRSS sized box spacecraft, which provides communications and 

navigation relay between lunar outposts (via LCT) or Sortie parties (with user radios) and large ground 

antennas on Earth. 

Figure 8 shows a front view of the LRS spacecraft. Called out are the two antenna, the solar arrays, 

and the two types of thrusters.

4.0 Subsystem Breakdown 

4.1 Communications

The multi-center COMPASS design session worked to achieve a simple spacecraft design which 

allows the Earth antenna, lunar antenna, solar arrays and Roll Control System (RCS) thrusters to all 

remain pointed at their targets without blockage, as LRS traverses its elliptical orbit around the Moon 

over the lunar South Pole, while the Moon orbits the Earth, and Earth orbits the Sun. Prior LRS designs, 

developed by JPL’s Team-X and GSFC’s IMDC design teams, over three to four months, were unable to 

resolve this problem without requiring two spacecraft, spacecraft flips as often as four times a day, two 

Earth Ka-band antennas, or heavy mechanically complex gimbals to support antennas beyond the solar 

arrays.

4.1.1 Communications Requirements

The LRS provides the following communications for the manned lunar base.

• Provide connectivity between Habitat and 

○ Surface users not on LCT network 

○ Earth (Deep Space Network (DSN))

• Provide connectivity between LCT network and 

○ Earth (DSN)

○ Surface users not on LCT network

• Provide connectivity between users not on LCT network and 

○ Earth (DSN)

○ Habitat

○ Other users not on LCT network

• Support priority based demand access service for following types of IP-based connections

○ Two way full-duplex voice and data between 

– Surface user and surface user

– Surface user and habitat

– Surface user and Earth
NASA/TM—2012-217140 14

○ One way video from 

– Surface user to habitat and to Earth

– Habitat to Earth 

– Earth to habitat

4.1.2 Communications Assumptions

• Physical layer and MAC layer protocol for satellite and LCT can be designed independently

• Users can carry separate radios for LCT and sat communications

• Satellite-based tracking services use SN signaling 

○ Communications and tracking services are not required to be integrated for data rates above 

192 kbps

○ Users requiring sat communications services above 192 kbps carry a separate Ka band 

transceiver in the 26/23 GHz band

• Users have the ability to point directional antennas to the satellite with sufficiently high accuracy

• 2 kbps emergency voice service is acceptable for communications with stranded astronauts 

• No satellite services to lunar surface when satellite is in safe mode

○ Safe mode/relay links operating frequencies are in reverse S-band configuration

4.1.3 Communications Design and MEL

On the top face, a Q-band Earth antenna (40/37 GHz) with a 0.5 m dish provides communication with 

Mission Control. A unique 110° joint in the gimbals allows the Earth antenna to look over the solar arrays 

to see Earth thus avoiding the need to flip the spacecraft over as often as four times a day (as previous 

designs by JPL and GSFC had required) during times in the orbit when Earth is sometimes above and 

sometimes below 90°. A low gain S band (2 GHz) antenna transmits/receives navigation from Network 

Control and telemetry from Mission Control. Figure 9 illustrates the viewing angle of the 110° joint 

gimbals on the Q-band antenna, and their ability to look over and around the two solar arrays mounted on 

the sides of the spacecraft. The brown cone is the cutout zone below which the Q-band antenna cannot see 

the Earth. This illustrates that the antenna can look more than over a 180° range, but slightly over each 

side. When the Earth enters into that brown cutout zone, the spacecraft will have to perform a maneuver 

to flip about that axis.

On the front face (velocity vector), a Ka/K band antenna (23/26 GHz) with a 1-m dish provides 

communications with the Moon. This antenna, from the Lunar Reconnaissance Orbiter, has an integral 

Cassegrainian S band (2 GHz) antenna for navigation. This antenna will be flight qualified in 2009 at the 

new Ka/K frequencies (23/26 GHz).

Table 8 captures the line items (quantity, mass and power) from the MEL of the communication 

system for LRS as used in the COMPASS design session. This system was provided by the 

communications consultants from JPL and adapted as the study progressed to the final version seen here.

Figure 9.—Viewing angle of Q-band antenna on LRS.
NASA/TM—2012-217140 15

TABLE 8.—LRS COMMUNICATIONS SYSTEM MEL BREAKDOWN

Description Quantity Unit mass

(kg)

CBE mass

(kg)

Growth

(%)

Growth

(kg)

Total mass

(kg)

Nominal

power

(W)

Peak

power

(W)

Lunar COMM Relay Satellite (LRS) -- ---- 1030.5 8.7 90.1 1120.6 683.6 259.6

LRS-2 -- ---- 1030.5 9 90.1 1120.6 683.6 259.6

Communications -- ---- 78.4 3 2.4 80.8 494.0 0.0

Ka band (40/37) DTE Link -- ---- 17.1 3.00 0.5 17.6 217.0 0.0

Ka DTE transponder 2 2.9 5.8 3.00 0.2 6.0 17.0 0.0

Ka TWTA (100W) 2 2.6 5.2 3.00 0.2 5.4 200.0 0.0

Ka-band diplexer 2 0.4 0.8 3.00 0.0 0.8 0.0 0.0

WG transfer switch 1 0.4 0.4 3.00 0.0 0.4 0.0 0.0

Coax switch 1 0.2 0.2 3.00 0.0 0.2 0.0 0.0

Attenuator 1 0.1 0.1 3.00 0.0 0.1 0.0 0.0

Ka band hybrid coupler 1 0.1 0.1 3.00 0.0 0.1 0.0 0.0

Ka band Wave Guide 9 0.5 4.5 3.00 0.1 4.6 0.0 0.0

Miscellaneous no. 1 0 0.0 0.0 3.00 0.0 0.0 0.0 0.0

Miscellaneous no. 2 0 0.0 0.0 3.00 0.0 0.0 0.0 0.0

Ka band (23/26) Relay Link -- ---- 21.1 3.00 0.6 21.7 115.0 0.0

Ka band modulator 2 1.0 2.0 3.00 0.1 2.1 10.0 0.0

Ka band hybrid coupler 1 0.1 0.1 3.00 0.0 0.1 0.0 0.0

Ka TWTA (20W) 2 2.6 5.2 3.00 0.2 5.4 40.0 0.0

WG transfer switch 1 0.4 0.4 3.00 0.0 0.4 0.0 0.0

Ka-band diplexer 1 0.4 0.4 3.00 0.0 0.4 0.0 0.0

LNA-splitter 1 1.0 1.0 3.00 0.0 1.0 5.0 0.0

Ka receiver (up to four on, two spare) 6 2.0 12.0 3.00 0.4 12.4 60.0 0.0

Miscellaneous no. 1 0 0.0 0.0 3.00 0.0 0.0 0.0 0.0

Miscellaneous no. 2 0 0.0 0.0 3.00 0.0 0.0 0.0 0.0

Ka band (40/37) TT&C and Ranging -- ---- 17.1 3.00 0.5 17.6 42.0 0.0

Ka DTE transponder 2 2.9 5.8 3.00 0.2 6.0 17.0 0.0

Ka TWTA (10 W) 2 2.6 5.2 3.00 0.2 5.4 25.0 0.0

Ka-band diplexer 2 0.4 0.8 3.00 0.0 0.8 0.0 0.0

WG transfer switch 1 0.4 0.4 3.00 0.0 0.4 0.0 0.0

Coax switch 1 0.2 0.2 3.00 0.0 0.2 0.0 0.0

Attenuator 1 0.1 0.1 3.00 0.0 0.1 0.0 0.0

Ka band hybrid coupler 1 0.1 0.1 3.00 0.0 0.1 0.0 0.0

Ka band wave guide 9 0.5 4.5 3.00 0.1 4.6 0.0 0.0

Miscellaneous no. 1 0 0 0.0 3.00 0.0 0.0 0.0 0.0

Miscellaneous no. 2 0 0 0.0 3.00 0.0 0.0 0.0 0.0

S band Relay Link -- ---- 9.3 3.00 0.3 9.6 80.0 0.0

S band transceiver (10 W) 2 3.5 7.0 3.00 0.2 7.2 80.0 0.0

S band coax switch 1 0.3 0.3 3.00 0.0 0.3 0.0 0.0

S band coax 10 0.2 2.0 3.00 0.1 2.1 0.0 0.0

Data Router and SSR -- ---- 4.0 3.00 0.1 4.1 20.0 0.0

Data router and SSR 1 4 4.0 3.00 0.1 4.1 20.0 0.0

Antennas (Booms excluded) -- ---- 9.8 3.00 0.3 10.1 20.0 0.0

S band relay, 1- by 1-m patch array 0 2 0.0 3.00 0.0 0.0 0.0 0.0

Gimbal 0 1.4 0.0 3.00 0.0 0.0 5.0 0.0

S/Ka band relay, 1 m dia. reflector 1 4 4.0 3.00 0.1 4.1 0.0 0.0

Feed 1 0.5 0.5 3.00 0.0 0.5 0.0 0.0

Gimbal 1 1.4 1.4 3.00 0.0 1.4 5.0 0.0

Ka band DTE , 0.5 m dia. reflector 1 1 1.0 3.00 0.0 1.0 0.0 0.0

Feed 1 0.5 0.5 3.00 0.0 0.5 0.0 0.0

Gimbal 1 1.4 1.4 3.00 0.0 1.4 10.0 0.0

S band LGA (T/R) for safing 2 0.5 1.0 3.00 0.0 1.0 0.0 0.0
NASA/TM—2012-217140 16

4.1.4 Communications Trades

In order to overcome the pointing issues with the LRS communications system, several design 

scenarios for the communications antenna were considered.

• Options: Single ‘110°’ pointing or two Earth pointing antennas to eliminate ‘flipping’ the 

spacecraft (S/C) as often as previous studies had indicated was necessary

○ Chosen: Found design to ‘point around’ S/C, use single 0.5-m Q band antenna (see Figure 9

for viewing angle of the top antenna on LRS)

• Options: Two (1 S and 1 Ka band) or one (combined S and Ka band) antenna 

○ Chosen: Single combined Ka and S 1-m antenna

4.2 Avionics

For the purposes of modeling the LRS system, the Avionics section included both Command and 

Data Handing (C&DH) and Guidance, Navigation and Control (GN&C) hardware.

4.2.1 Avionics Requirements

• Storage array for 24 hr of storage (0.3 TB)

• Satellite capacity requirements

○ Support the following peak offered load from the lunar surface to relay sat 

– 100 Mbps from Habitat

– 100 Mbps from LCT

– 50 Mbps from surfaces users not on the LCT network or in the Habitat

○ Support a forward link rate of up to 100 Mbps from relay sat to lunar surface 

○ Support two-way ranging to up to five users simultaneously

4.2.2 Avionics Assumptions

The avionics shall be single fault tolerant.

4.2.3 Avionics Design and MEL

Avionics components are based on commercially available components come from British Aerospace 

(BAE). All avionics assume 3U-160 form factor cards.

4.2.4 GN&C Design

The GN&C subsystem line items were kept under the Avionics subsystem in the LRS MEL. The LRS 

GN&C subsystem provides navigation and attitude determination using

• Twelve coarse Sun sensor suites

• Two star tracking cameras

• Two inertial measurement units 

LRS GN&C provides momentum control using

• Four reaction wheels with momentum storage capability of 20 Nms

• RW thrust assembly

LRS GN&C controls and logic are included in the Avionics system elements. Table 9 details the line 

items contained in the LRS Avionics MEL rolled up into the LRS design.
NASA/TM—2012-217140 17

TABLE 9.—LRS AVIONICS DETAILED MEL

Description Quantity Unit

mass

(kg)

CBE 

Mass

(kg)

Growth

()

Growth

(kg)

Total

mass

(kg)

Nominal 

power

(W)

Peak 

power

(W)

Lunar COMM Relay Satellite (LRS) --- ------- 1030.5 8.7 90.1 1120.6 683.6 259.6

LRS-2 --- ------- 1030.5 9 90.1 1120.6 683.6 259.6

Avionics --- ------- 89.6 24 21.1 110.7 189.6 259.6

Command & Data Handling --- ------- 53.7 23.22 12.5 66.1 157.6 157.6

General Avionics Processor 2 11 22.0 20.00 4.4 26.4 65.0 65.0

Time generation unit 2 0.4 0.8 3.00 0.0 0.8 0.6 0.6

Command and control harness (data) 1 4.88 4.9 50.00 2.4 7.3 0.0 0.0

Instrumentation & wiring 0 0 0.0 50.00 0.0 0.0 0.0 0.0

Data recorder 2 11 22.0 20.00 4.4 26.4 72.0 72.0

Atomic clock 2 2 4.0 30.00 1.2 5.2 20.0 20.0

Miscellaneous no. 2 0 0 0.0 0.00 0.0 0.0 0.0 0.0

Guidance, Navigation and ACS --- ------- 35.9 23.96 8.6 44.5 32.0 102.0

Reaction wheel assembly (RWA) 4 7 28.0 30.00 8.4 36.4 15.0 80.0

RWA mount 1 0.2 0.2 18.00 0.0 0.2 0.0 0.0

Star camera 2 3 6.0 2.00 0.1 6.1 5.0 10.0

Star camera mount 1 0.1 0.1 18.00 0.0 0.1 0.0 0.0

Inertial measurement units 2 0.75 1.5 2.00 0.0 1.5 12.0 12.0

Course Sun sensor suite 12 0.01 0.1 2.00 0.0 0.1 0.0 0.0

4.2.5 Avionics Risk Inputs

The 0.3 TB storage array is design around future technology. If the technology for the storage array is 

not yet available when the LRS is built, this component could be larger than expected.

4.3 Structures and Mechanisms

4.3.1 Structures and Mechanisms Requirements

Requirements include the following

• Provide sufficient space to house the necessary hardware for avionics, communications, 

propulsion, and power.

• Sustain maximum loads with minimum deflections without damage.

• Minimize weight.

• Minimize vibrations.

• Fit within the confines of the launch vehicle.

4.3.2 Structures and Mechanisms Assumptions

The primary structures material is aluminum (Al) 2090. It consisted of a tubular space frame 

construction with honeycomb core composite sandwich structure panels. A maximum axial load is 5.5 g 

and a maximum lateral load is 3 g.

4.3.3 Structures and Mechanisms Design and MEL

The material of the structure is Al 2090. The architecture of the structure consists of a tubular space 

frame with honeycomb core composite sandwich structure panels. Welds and threaded fasteners are used 

for the assembly.

Table 10 details the line items used in the Structures and Mechanical Systems subsystem used to 

build the LRS. CBE Mass, Growth and Total Mass are displayed.
NASA/TM—2012-217140 18

TABLE 10.—LRS STRUCTURES AND MECHANICAL SYSTEMS DETAILED MEL

Description Quantity Unit mass

(kg)

CBE mass

(kg)

Growth

()

Growth

(kg)

Total mass

(kg)

Lunar COMM Relay Satellite (LRS) -- ----- 1030.5 8.7 90.1 1120.6

LRS-2 -- ----- 1030.5 9 90.1 1120.6

Structures & Mechanical Systems -- ----- 180.5 14 25.1 205.6

Primary Structures -- ----- 164.1 14.43 23.7 187.7

Separation ring/flange 1 18.72 18.7 10.00 1.9 20.6

Top/bottom deck 2 5.735 11.5 15.00 1.7 13.2

Vertical inner panels 4 3.66 14.6 15.00 2.2 16.8

Side panels 4 13.09 52.4 15.00 7.9 60.2

Corner post 4 1.899 7.6 15.00 1.1 8.7

Top cover 0 0 0.0 15.00 0.0 0.0

Thrust tube 1 59.28 59.3 15.00 8.9 68.2

Miscellaneous clips/fasteners 0 0 0.0 15.00 0.0 0.0

Miscellaneous no. 1 0 0 0.0 15.00 0.0 0.0

Miscellaneous no. 2 0 0 0.0 15.00 0.0 0.0

Secondary Structures -- ----- 9.4 15.00 1.4 10.8

Antenna boom 1 5.126 5.1 15.00 0.8 5.9

Solar array boom(s) 2 2.148 4.3 15.00 0.6 4.9

EP thruster booms (N/A) 0 0.537 0.0 15.00 0.0 0.0

Chemical thruster booms 0 0.537 0.0 15.00 0.0 0.0

Miscellaneous no. 1 0 0 0.0 0.00 0.0 0.0

Miscellaneous no. 2 0 0 0.0 0.00 0.0 0.0

Installation -- ----- 7.0 0.00 0.0 7.0

C&DH installation 1 0.916 0.9 0.00 0.0 0.9

Communications installation 1 5.212 5.2 0.00 0.0 5.2

GN&C installation 1 0.916 0.9 0.00 0.0 0.9

Power installation 0 0 0.0 0.00 0.0 0.0

Propulsion installation 0 0 0.0 0.00 0.0 0.0

Propellant storage installation 0 0 0.0 0.00 0.0 0.0

Thermal installation 0 0 0.0 0.00 0.0 0.0

Miscellaneous no. 1 0 0 0.0 0.00 0.0 0.0

Miscellaneous no. 2 0 0 0.0 0.00 0.0 0.0

Miscellaneous no. 3 0 0 0.0 0.00 0.0 0.0

Mechanisms -- ----- 0.0 0.00 0.0 0.0

Solar array deployment mech (in power) 0 0 0.0 0.00 0.0 0.0

Radiator deployment mechanism (if applicable) 0 0 0.0 0.00 0.0 0.0

Separation mechanism (pyros) 0 0 0.0 0.00 0.0 0.0

Bi-axial antenna gimbal 0 0 0.0 0.00 0.0 0.0

4.3.4 Structures and Mechanisms Analytical Methods

A preliminary structural analysis using basic calculations with given launch loads was conducted.

4.3.5 Structures and Mechanisms Risk Inputs

Primary risks to maintaining structural integrity include the following

• Potential impact with foreign object.

• Excessive vibrations.

• Excess inertial loading from the launch vehicle.

4.3.6 Structures and Mechanisms Recommendation

A complete finite element analysis (FEA) to determine stresses and displacements along with a modal 

analysis for vibrations is necessary for further study.
NASA/TM—2012-217140 19

4.4 Electrical Power System

4.4.1 Power Requirements

During eclipse periods longer than 2.5 hr, the power available must meet nominal power 

requirements. 

4.4.2 Power Assumptions

The solar arrays are assumed to use GaAs triple-junction cells that are 28 percent efficient at the 

beginning-of-life. An off-point angle is assumed that accounts for 15° cosine losses between the Sun and 

solar array. The battery modules are assumed to be lithium-ion chemistry with a cell specific energy of 

170 W-hr/kg and a module specific energy of 72 W-hr/kg. The spacecraft bus is assumed to be nominally 

28 Vdc. A 30 percent power margin is assumed and 30 percent mass margin on everything except the 

wire harness, which is at 50 percent mass margin.

4.4.3 Power Design and MEL

Located on the sides of the LRS, the 1.7- by 1.7-m (area of 2.9 m2

) solar array panels rotate to face 

the Sun to generate a total of 1 kW (1040 We) useful power. The arrays are mounted to the base of the 

spacecraft to avoid blockage to the Earth antenna. Radiator panels on the sides provide heat rejection. 

Two equal area solar array wings are used to have some fault tolerance and help with spacecraft 

balance. The arrays are sized to meet the peak power requirement plus 30 percent power margin. A single 

axis gimbal is assumed. The battery system is sized to meet the nominal power requirement for a 2.5 hr 

eclipse duration. For longer eclipses, a lower amount of power will be available.

Table 11 details the line items used in the Power Systems subsystem used to build the LRS. CBE 

Mass, Growth and Total Mass are displayed.

TABLE 11.—LRS POWER SYSTEM DETAILED MEL

Description Quantity Unit mass

(kg)

CBE mass

(kg)

Growth

(%)

Growth

(kg)

Total mass

(kg)

Lunar COMM Relay Satellite (LRS) -- ------ 1030.5 8.7 90.1 1120.6

LRS-2 -- ------ 1030.5 9 90.1 1120.6

Power System -- ------ 72.6 31 22.2 94.8

Battery System -- ------ 40.3 30.00 12.1 52.4

Battery assembly-secondary 1 40.3 40.3 30.00 12.1 52.4

Miscellaneous no. 1 0 0 0.0 0.00 0.0 0.0

Miscellaneous no. 2 0 0 0.0 0.00 0.0 0.0

Miscellaneous no. 3 0 0 0.0 0.00 0.0 0.0

Solar Array -- ------ 26.1 30.00 7.8 33.9

Solar array panel 2 6.4 12.8 30.00 3.8 16.6

Solar array structure 2 0.64 1.3 30.00 0.4 1.7

Solar array drive assembly 2 3 6.0 30.00 1.8 7.8

Solar array interface (carried in structures) 0 0 0.0 0.00 0.0 0.0

Solar array deployment mech 2 3 6.0 30.00 1.8 7.8

Miscellaneous no. 2 0 0 0.0 0.00 0.0 0.0

Miscellaneous no. 3 0 0 0.0 0.00 0.0 0.0

Miscellaneous no. 4 0 0 0.0 0.00 0.0 0.0

Miscellaneous no. 5 0 0 0.0 0.00 0.0 0.0

Miscellaneous no. 6 0 0 0.0 0.00 0.0 0.0

Power Management & Distribution -- ------ 6.2 37.10 2.3 8.5

Battery charge controller 1 4 4.0 30.00 1.2 5.2

Wire harness 2 1.1 2.2 50.00 1.1 3.3

4.4.4 Power Trades

One solar array wing would reduce structure and mechanism mass while potentially reducing fault 

tolerance and spacecraft balance. The battery sizing is sensitive to power level and eclipse duration. If a 

lower power load can be used during eclipse cycles, the battery mass will be reduced proportionally.
NASA/TM—2012-217140 20

4.4.5 Power Analytical Methods

The battery and solar array system sizing is based on bottoms-up algorithms used in the System 

Power Analysis for Capability Evaluation (SPACE) code. The Solar Array Drive Assembly (SADA)

system is based on Moog designs.

4.4.6 Power Recommendation

A detailed timeline of power loads versus time with the eclipse time period included would improve 

the sizing of both the solar array and battery systems. A more accurate peak and eclipse requirement 

would be identified that would directly affect the design.

4.5 Propulsion and Propellant Management

4.5.1 Propulsion and Propellant Management Requirements

There were two operational requirements for the propulsion subsystem:

• Two large burns for lunar orbit insertion and modification

• Several smaller burns required for orbit modifications

• The 100-lbf thruster was selected to perform these burns efficiently

• Attitude Control burns for momentum wheel desaturation.16 thrusters provide sufficient couples 

for rotating spacecraft on three axes and provide single fault (~1.5) tolerance

4.5.2 Propulsion and Propellant Management Assumptions

Assumption 1—Thruster Operation

• Thruster operation constant over duration of burn

Assumption 2—Propellant Storage

• Propellant delivery assumed to be isothermal

• Same with helium (He) pressurant into main tank

Assumption 3—Design for Minimum Risk (DMR) applied to propellant storage tank and propellant 

delivery lines, following Exploration Program lead. Additionally, DMR applied to passive elements only.

4.5.3 Propulsion and Propellant Management Design and MEL

The propulsion subsystem is comprised of the following main elements

• Two 100-lbf-thrust MR-104 hydrazine monopropellant engines—one operating, one spare

• Sixteen 0.5-lbf-thrust MR-111E hydrazine monopropellant engines—all operating

• One metallic titanium (Ti) hydrazine storage tank

• Two metallic Ti high pressure He storage tanks for pressurant gas

• Propellant distribution system to control propellant delivery from storage tank to thrusters

• Pressurant distribution system to control pressurant delivery to hydrazine storage tank

Figure 10 shows the bottom view of the LRS highlighting the propulsion system.
NASA/TM—2012-217140 21

Figure 10.—Bottom view of main engines on LRS.

On the bottom face, the OMS main engine provides hydrazine-fueled propulsion. Four RCS thruster 

pods at the corners provide yaw, pitch, and roll. 

The main propulsion thruster selected for the LRS was the Aerojet MR-104 Hydrazine thruster. A 

monoprop engine was selected to simplify the propulsion subsystem. The primary characteristics of the 

main thruster are:

 Thrust: 441 N (100 lbf); 204.6 to 572.5 N (46 to 128.7 lbf)

 Specific Impulse: 239; 239-223 s

 Mass: 1.86 kg

 Power: 43.1 W

 Lifetime: 2,654 s (cumulative)

The thruster selected for the reaction control system on LRS was the Aerojet MR-111E Hydrazine 

engine. This is also a monoprop device with the following characteristics:

 Thrust: 2.2 N (0.5 lbf); 0.5 to 2.2 N (0.11 to 0.5 lbf)

 Specific Impulse: 224; 224-213 s

 Mass: 0.33 kg

 Power: 13.64 W

 Lifetime: 26.7 hr (cumulative)
NASA/TM—2012-217140 22

TABLE 12.—LRS PROPULSION HARDWARE DETAILED MEL

Description Quantity Unit mass

(kg)

CBE mass

(kg)

Growth

(%)

Growth

(kg)

Total mass

(kg)

Lunar COMM Relay Satellite (LRS) -- ------- 1030.5 8.7 90.1 1120.6

LRS-2 -- ------- 1030.5 9 90.1 1120.6

Propulsion (Chemical) -- ------- 21.8 8 1.7 23.5

Main Engine -- ------- 5.5 8.00 0.4 6.0

Main engine 1 5.543 5.5 8.00 0.4 6.0

Main engine gimbal 0 0 0.0 0.00 0.0 0.0

Reaction Control System -- ------- 16.3 8.00 1.3 17.6

RCS engine 4 4.064 16.3 8.00 1.3 17.6

Figure 11.—Propulsion system schematic diagram.

Table 12 lists the MEL for the main hardware of the propulsion system (main engine, RCS system, 

etc.). The propellant and pressurant distribution systems are comprised of gas/fluid delivery lines, the

control components (isolation/latching valves, control valves, filters, and service valves). The 

instrumentation for pressure and temperature are included in the distribution subassemblies. Finally, 

hardware for attachment/positioning are included here. The propellant subsystem schematic shown in 

Figure 11 provides a notional drawing of the physical components included in it.

The hydrazine propellant for main and reaction control propulsion was stored in a Ti pressure vessel 

based on a COTS unit from ATK-PSC, Inc., Model No. 80356-1. The baseline tank dimensions were:

 Size: 0.91- by 1.1-m L (3.0 ft by 3.6 ft L)
NASA/TM—2012-217140 23

 Internal volume: 0.59 m3 (20.8 ft3)

Table 13 details the propellant management system line item breakdown including tanks and fuel 

lines. Table 14 lists the MEL for the propellant used to perform the LRS Mission.

TABLE 13.—LRS PROPELLANT MANAGEMENT DETAILED MEL

Description Quantity Unit mass

(kg)

CBE mass

(kg)

Growth

(%)

Growth

(kg)

Total mass

(kg)

Lunar COMM Relay Satellite (LRS) -- -------- 1030.5 8.7 90.1 1120.6

LRS-2 -- -------- 1030.5 9 90.1 1120.6

Propulsion (Chemical) -- -------- 21.8 8 1.7 23.5

Propellant Management (Chemical) -- -------- 72.9 15 10.9 83.9

Main Propellant Management -- -------- 69.2 15.00 10.4 79.6

Fuel tanks 1 41.35 41.4 15.00 6.2 47.6

Fuel lines 0 0 0.0 0.00 0.0 0.0

Oxidizer tanks 0 0 0.0 0.00 0.0 0.0

Oxidizer lines 0 0 0.0 0.00 0.0 0.0

Pressurization system: tanks, panels, lines 1 23.67 23.7 15.00 3.6 27.2

Feed system: regulators, valves, etc 1 4.201 4.2 15.00 0.6 4.8

RCS Propellant Management -- -------- 3.7 15.00 0.6 4.3

Fuel tanks 0 0 0.0 0.00 0.0 0.0

Fuel lines 0 0 0.0 0.00 0.0 0.0

Pressurization system: tanks, panels, lines 0 0 0.0 0.00 0.0 0.0

Feed system: regulators, valves, etc 1 3.705 3.7 15.00 0.6 4.3

Propellant (Chemical) -- -------- 467.3 0 0.0 467.3

TABLE 14—LRS PROPELLANT DETAILED MEL

Description Quantity Unit mass

(kg)

CBE mass

(kg)

Growth

(%)

Growth

(kg)

Total mass

(kg)

Lunar COMM Relay Satellite (LRS) -- --------- 1030.5 8.7 90.1 1120.6

LRS-2 -- --------- 1030.5 9 90.1 1120.6

Propellant (Chemical) -- --------- 467.3 0 0.0 467.3

Main Engine Propellant -- --------- 464.4 0.00 0.0 464.4

Fuel -- --------- 464.4 0.00 0.0 464.4

Fuel usable 1 453.1 453.1 0.00 0.0 453.1

Fuel boiloff 0 0 0.0 0.00 0.0 0.0

Fuel residuals (unused) 1 11.33 11.3 0.00 0.0 11.3

Oxidizer -- --------- 0.0 0.00 0.0 0.0

Oxidizer usable 0 0 0.0 0.00 0.0 0.0

Oxidizer boiloff 0 0 0.0 0.00 0.0 0.0

Oxidizer residuals (unused) 0 0 0.0 0.00 0.0 0.0

Pressurant 0 0 0.0 0.00 0.0 0.0

RCS Propellant -- --------- 2.9 0.00 0.0 2.9

Fuel -- --------- 0.0 0.00 0.0 0.0

Fuel usable 0 0 0.0 0.00 0.0 0.0

Fuel boiloff 0 0 0.0 0.00 0.0 0.0

Fuel residuals (unused) 0 0 0.0 0.00 0.0 0.0

Oxidizer -- --------- 0.0 0.00 0.0 0.0

Oxidizer usable 0 0 0.0 0.00 0.0 0.0

Oxidizer boiloff 0 0 0.0 0.00 0.0 0.0

Oxidizer residuals (unused) 0 0 0.0 0.00 0.0 0.0

Pressurant 1 2.929 2.9 0.00 0.0 2.9

Minor size changes were to made to match propellant load determined by the in-house developed 

Mission analysis tool mentioned in the mission assumptions section. These changes consisted primarily of 
NASA/TM—2012-217140 24

increasing the tank length but keeping the tank diameter fixed. A propellant management device is 

included in tank mass roll-up.

A He pressurization system is necessary to maintain the propellant delivery pressure to all of the 

engines. This system is comprised to two spherical metallic Ti tanks and a pressurant distribution and 

control feed system. The pressurization subsystem is derived from current configuration designs or 

existing hardware to minimize or eliminate technology development activity. For example, the pressurant 

tanks are based on an Arde, Inc., high-pressure tank that is readily available. 

All components of propulsion subsystem are at a high technology readiness level, so little or no 

development costs required.

4.5.4 Propulsion and Propellant Management Trades

The primary trade study performed within the propulsion subsystem was to consider a Blowdown 

propellant storage and delivery approach versus an inert gas pressurization subassembly. Preliminary 

trade on Blowdown system found the tanks required for storage of the hydrazine propellant to be heavier

and require more volume within the spacecraft. While more complex, the He pressurization system was 

selected in order to save both weight and volume. Single metallic tank selected for storage of the 

hydrazine propellant. While this configuration is zero-fault tolerance, it has been the approach of 

Exploration to consider passive elements to have innately lower risk; consequently the approach for the 

tank, as well as the propellant feed lines is acceptable.

4.5.5 Propulsion and Propellant Management Analytical Methods

The analytical techniques used in the propulsion subsystem model are as follows:

1. Mass roll-up of the known components and assemblies within the model. This included the 

thrusters, the feed system control components, and instrumentation.

2. Hoop Stress analysis with material characteristics derived from commercial-off-the-shelf (COTS) 

high-pressure tanks were used to properly size the propellant storage and pressurant tanks in 

response to changes in propellant loads.

3. Feed lines were estimated from expected run lengths within the spacecraft and lineal density 

calculations developed for hypergol propellants.

4. Pressurant gas estimates were made with gas density estimates obtained with National Institute of 

Standards and Technology (NIST) databases.

4.5.6 Propulsion and Propellant Management Risk Inputs

Two potential risks were identified:

1. RCS Thruster Plume Impingement (LOW)

a. Degradation of spacecraft performance due to deposition on sensitive surfaces, such as the 

solar arrays or the radiators may be possible due to the RCS thruster plumes. Careful 

examination of potential interferences should be undertaken.

2. Propellant freezing within fluid lines (LO-MED)

a. Freezing of the hydrazine within the feed lines, particularly the lines going to the RCS 

thrusters, could lead to line ruptures or other performance failures. Careful modeling of the 

spacecraft thermal environment to understand where cold spots can form and proper 

configuration of thermal insulation is necessary to protect the integrity of the feedlines.

4.5.7 Propulsion and Propellant Management Recommendation

The propulsion subsystem used in the LRS vehicle analysis is a conventional assembly of mature and 

proven components. No new technologies are needed for this vehicle. Therefore, the best 
NASA/TM—2012-217140 25

recommendation to be made is that proper modeling and investigation of the thermal conditions on the 

spacecraft be performed to insure that propulsion subsystem can operate properly.

4.6 Thermal Control

4.6.1 Thermal Requirements

The thermal requirements for the LRS were to provide a means of cooling the spacecraft during 

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

throughout the mission.

The maximum heat load to be rejected by the thermal system was 480 W, and the desired operating 

temperature for the electronics and propellant was 300 K. 

4.6.2 Thermal Assumptions

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

operational environment. It was assumed that operation would take place within the lunar orbital 

environment. The following assumptions were utilized to size the thermal system. 

• The view factors for the radiator to the Earth, lunar surface and solar array 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.

4.6.3 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. Figure 

12 shows an illustration of the major items in the thermal system.

Excess heat is collected from a series of Al cold plates located throughout the interior of the 

spacecraft. 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 spacecraft body and are integrated to the radiator are protected with a micro meteor 

shield. The radiator has exterior louvers on it to provide some control over its heat transfer capability. 

The radiator was sized with approximately 75 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. 

Multilayer insulation is also utilized on the spacecrafts, and propellant system to regulate and 

maintain the desired temperatures. 

Figure 12 is a notional block diagram of the separate components of the LRS thermal subsystem.

Table 15 lists the MEL for the thermal control system sketched in Figure 12 used to perform the LRS 

Mission.
NASA/TM—2012-217140 26

Figure 12.—LRS thermal control system.

TABLE 15.—LRS THERMAL CONTROL SYSTEM DETAILED MEL

Description Quantity Unit mass

(kg)

CBE mass

(kg)

Growth

(%)

Growth

(kg)

Total mass

(kg)

Lunar COMM Relay Satellite (LRS) ------- --------- 1030.5 8.7 90.1 1120.6

LRS-2 ------- --------- 1030.5 9 90.1 1120.6

Thermal Control (Non-Propellant) ------- --------- 47.3 14 6.7 54.0

Active Thermal Control ------- --------- 4.7 12.77 0.6 5.3

Heaters 10 0.143 1.4 10.00 0.1 1.6

Thermal control/heaters circuit 2 0.2 0.4 10.00 0.0 0.4

Data acquisition 1 1 1.0 15.00 0.2 1.2

Thermocouples 25 0.01 0.3 10.00 0.0 0.3

Radiator MMOD shielding 2.07 0.769 1.6 15.00 0.2 1.8

Miscellaneous no. 2 0 0 0.0 0.00 0.0 0.0

Passive Thermal Control ------- --------- 34.9 14.32 5.0 39.8

Heat sinks 4 3.463 13.9 15.00 2.1 15.9

Heat pipes 1 1.98 2.0 15.00 0.3 2.3

Radiators 1 12.12 12.1 15.00 1.8 13.9

MLI 1 5.497 5.5 12.00 0.7 6.2

Temperature sensors 50 0.01 0.5 10.00 0.1 0.6

Phase change devices 0 0 0.0 0.00 0.0 0.0

Thermal coatings/paint 1 0.907 0.9 10.00 0.1 1.0

Miscellaneous no. 1 0 0 0.0 0.00 0.0 0.0

Miscellaneous no. 2 0 0 0.0 0.00 0.0 0.0

Semi-Passive Thermal Control ------- --------- 7.8 14.49 1.1 8.9

Louvers 1 6.977 7.0 15.00 1.0 8.0

Thermal switches 4 0.2 0.8 10.00 0.1 0.9

4.6.4 Thermal Analytical Methods

The analysis performed to size the thermal system is based on first principle heat transfer from the 

spacecraft to the surroundings. This analysis takes into account the design and layout of the thermal 

system and the thermal environment to which heat is being rejected or the vehicle is being insulated from. 

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

System Sizing” was produced.
NASA/TM—2012-217140 27

4.6.5 Thermal Risk Inputs

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

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, micrometeor impact or 

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

loss of all or a portion of the spacecraft’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.

4.6.6 Thermal Recommendation

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

system design changes can be made.

• Redundant heat pipes 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 spacecraft to insure that the interior components do not drop below their desired 

minimum temperature based on a known shadow period of operation. 

5.0 Cost and Risk

5.1 Costing

Costing results and assumptions will be documented in a subsequent report.

5.2 Risk Analysis and Reduction

• Risk Management Requirements for LRS

○ The management of risk is a foundational issue in the design, development and extension of 

technology.

○ Each subsystem was tasked to write a risk statement regarding any concerns, issues and ‘ah 

ha’s’. Mitigation plans would focus on recommendations to alleviate, if not eliminate the risk.

5.2.1 Risk Assumptions

• Assumptions for Risk Management

○ Risk attributes are based on Crew Exploration Vehicle (CEV) risk values.

○ Risk List is not based on trends or criticality.

○ Some mitigation plans are offered as suggestions.

○ Based on the derived model of this study only.

5.2.2 Risk Matrix and Risk List
NASA/TM—2012-217140 28

TABLE 16.—LRS RISK MATRIX

No. L×C Team Risk title

1 4×4 Design Overly Complex 

Spacecraft Design

2 3×4 Avionics Avionics System 

Functions Not 

Completely Defined

3 3×4 Avionics Data Recorder Design 

Based on Future 

Technology

4 2×5 Structure Potential Structure 

Failure

5 3×3 Avionics Atomic Clock at Low 

TRL

6 3×3 Communications Health Management

TABLE 17.—LRS RISK LIST AND DETAILED DATA TABLE
NASA/TM—2012-217140 29
NASA/TM—2012-217140 30

TABLE 17.—CONCLUDED.

5.2.3 Risk Summary

• Areas of concern for this study include:

○ Overly complex design

○ Design of Avionics

○ Available technology for communications

○ Impact damage to the structures

• These risks, with proper pro-active planning can be mitigated early to avoid becoming problems 

late in the development life cycle.
NASA/TM—2012-217140 31

Appendix A.—Acronyms and Abbreviations

AF U.S. Air Force

Al aluminum

AO Announcement of Opportunity

BAE British Aerospace

C&DH Command and Data Handing

CBE current best estimate

CEV Crew Exploration Vehicle

Comm Communications

COMPASS COllaborative Modeling and Parametric Assessment of Space Systems

COTS commercial off the shelf

DMR Design for Minimum Risk

DSN Deep Space Network

DTE direct to Earth

EELV Evolved Expendable Launch Vehicle

EP Electric Propulsion

ESPA EELV Secondary Payload Adaptor

FEA finite element analysis 

FOM figure of merit

GSFC NASA Goddard Space Flight Center

GLIDE GLobal Integrated Design Environment

GN&C Guidance, Navigation and Control

GRC NASA Glenn Research Center

hab habitat

HQ NASA Headquarters

IMDC I M Design Center

IP internet protocol

ISRU in situ resource utilization

He helium

JPL NASA Jet Propulsion Laboratory

KSC NASA Kennedy Space Center

LCT Lunar Communications Terminal

LRS Lunar Relay Station

LSP Launch Service Program

LSTO Launch Service Task Order

MAC Media Access Control

MEL Master Equipment List

MPU Makeup Power Unit

NASA National Aeronautics and Space Administration

Nav navigation

NIST National Institute of Standards and Technology
NASA/TM—2012-217140 32

NLS NASA Launch Services

OMS Orbital Maneuvering System

PEL Power Equipment List

PN pseudo-noise

RCS Reaction Control System

S/C spacecraft

SADA Solar Array Drive Assembly

SEP Solar Electric Propulsion

SN signal-to-noise

SPACE System Power Analysis for Capability Evaluation

SPU Solar Power Unit

Ti titanium

TDRSS Tracking and Data Relay Satellite System

TWTA Traveling Wave Tube Amplifier
NASA/TM—2012-217140 33

Appendix B.—Rendered Design Drawings

Figure B.1 and Figure B.2 show two views of the stowed configuration of the LRS with arrays folded 

at the sides, and the Q-band antenna on the top face also showed in the stowed position with major 

spacecraft dimensions.

Figure B.1.—LRS stowed configuration K/Ka band antenna face.

Figure B.2.—LRS stowed configuration, solar array panel side.
NASA/TM—2012-217140 34

Figure B.3 and Figure B.4 show the LRS in deployed configuration with both the two solar arrays and 

the Q band antenna on the top face of the satellite deployed.

Figure B.3.—LRS deployed configuration, K/Ka band antenna face.

Figure B.4.—LRS deployed configuration.
NASA/TM—2012-217140 35

Appendix C.—Study Participants

Lunar Relay Satellite (LRS) Design Session

Subsystem Name Center Email

LAT2 Comm. and Nav Lead Jim Shier HQ James.Schier-1@nasa.gov

LRS Customer Mark Flanegan GSFC Mark.C.Flanegan@nasa.gov

Design Session Lead Steve Oleson GRC Steven.R.Oleson@nasa.gov

Communications Payload Interface 

(GRC) Joe Warner GRC Joseph.D.Warner@nasa.gov

Communications Sub team Lead 

(GRC) Lynn Anderson GRC Lynn.M.Anderson@nasa.gov

System Integration, Mass properties, 

MEL, final documentation Melissa McGuire GRC Melissa.L.Mcguire@nasa.gov

ELV, Integration and Test, Operations Jeff Woytach GRC Jeffery.M.Woytach@nasa.gov

Mission and ACS Doug Fiehler GRC Douglas.I.Fiehler@nasa.gov

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

Structures and Mechanisms John Gyekenyesi GRC John.P.Gyekenyesi@nasa.gov

Structures and Mechanisms Amy Stalker GRC Amy.R.Stalker@nasa.gov

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

Propulsion Tim Sarver-Verhey GRC Timothy.R.Verhey@grc.nasa.gov

Power Josh Freeh GRC Joshua.E.Freeh@nasa.gov

Command and Data Handling Jeff Juergens GRC Jeffrey.R.Juergens@nasa.gov

Communications Sub team (GRC) O. Scott Sands GRC Obed.S.Scott@nasa.gov

LRS Comm. interface Biren Shah JPL Biren.N.Shah@nasa.gov

Configuration Mark Poljack GRC Mark.D.Poljak@nasa.gov

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

Configuration Dave Peters GSFC

Software TC Nguyen GRC Thanh.C.Nguyen@nasa.gov

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

Risk/Reliability Anita Tenteris GRC Anita.D.Tenteris@nasa.gov

Risk/Reliability Bill Strack GRC bstrack@wowway.com
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13. SUPPLEMENTARY NOTES

14. ABSTRACT

The Lunar Relay Satellite (LRS) COllaborative Modeling and Parametric Assessment of Space Systems (COMPASS) session was tasked to 

design a satellite to orbit in an elliptical lunar polar orbit to provide relay communications between lunar South Pole assets and the Earth. 

The design included a complete master equipment list, power requirement list, configuration design, and brief risk assessment and cost 

analysis. The LRS is a half-TDRSS sized box spacecraft, which provides communications and navigation relay between lunar outposts (via 

Lunar Communications Terminals (LCT)) or Sortie parties (with user radios) and large ground antennas on Earth. The LRS consists of a 

spacecraft containing all the communications and avionics equipment designed by NASA Jet Propulsion Laboratory’s (JPL) Team X to 

perform the relay between lunar-based assets and the Earth. The satellite design is a standard box truss spacecraft design with a thermal 

control system, 1.7 m solar arrays for 1 kWe power, a 1 m diameter Ka/S band dish which provides relay communications with the LCT, 

and a Q-band dish for communications to/from the Earth based assets. While JPL’s Team X and Goddard Space Flight Center’s (GSFC) I M 

Design Center (IMDC) have completed two other LRS designs, this NASA Glenn Research Center (GRC) COMPASS LRS design sits 

between them in terms of physical size and capabilities. 

15. SUBJECT TERMS

Lunar communication; Parabolic reflectors; Lunar satellites; Relay satellites; Aerospace systems; Satellite orbits; Elliptical orbits; 

Lunar orbits 

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