Steven R. Oleson and Melissa L. McGuire

Glenn Research Center, Cleveland, Ohio

COMPASS Final Report: 

Lunar Network Satellite-High Rate (LNS-HR)

NASA/TM—2012-217139

July 2012

CD–2008–29
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Steven R. Oleson and Melissa L. McGuire

Glenn Research Center, Cleveland, Ohio

COMPASS Final Report: 

Lunar Network Satellite-High Rate (LNS-HR)

NASA/TM—2012-217139

July 2012

CD–2008–29

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135
Acknowledgments

The authors would like to thank the COMPASS Team members: Les Balkanyi, Laura Burke, Tony Colozza, Jon Drexler, Doug 

Fiehler, Josh Freeh, John Gyekenyesi, Joe Hemminger, Eric Knoblock, Tom Krivanek, Mike Martini, Nelson Morales, Tom Packard, 

Tom Parkey, Mark Poljak, Carl Sandifer, Tim Sarver-Verhey, Anita Tenteris, Joe Warner, Glenn Williams, and Jeff Woytach.

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 identification 

only. Their usage does not constitute an official 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 findings, 

subject to revision as analysis proceeds.
NASA/TM—2012-217139 iii

Contents

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

1.1 Frozen Orbit Case........................................................................................................................1

1.2 Lagrange Orbit Case....................................................................................................................1

2.0 Study Background and Assumptions....................................................................................................4

2.1 Introduction .................................................................................................................................4

2.2 Assumptions ................................................................................................................................5

2.2.1 Coverage Assumptions/Requirements.............................................................................5

2.2.2 Communications Assumptions/Requirements.................................................................6

2.2.3 Earth Based Telescopes for Laser Optical Trunk.............................................................7

2.2.4 Derived Requirements......................................................................................................7

2.3 Similar Past Missions ..................................................................................................................8

2.4 Margin and Growth Policy ..........................................................................................................9

2.4.1 Definition of Terms..........................................................................................................9

2.4.2 Definition of Mass Growth Allowance (MGA)...............................................................9

2.4.3 From the AIAA Standard—Terms and Definitions as applied to Dry Mass. ................11

2.4.4 Launch Margin...............................................................................................................12

2.4.5 Power Growth ................................................................................................................12

2.5 Mission Description—Frozen Orbits.........................................................................................12

2.5.1 Frozen Orbit Mission Analysis Assumptions ................................................................13

2.5.2 Frozen Orbit Mission Analysis Analytic Methods.........................................................13

2.5.3 Frozen Orbit Mission Analysis Event Timeline.............................................................16

2.5.1 Frozen Orbit Delta V Schedule......................................................................................16

2.5.2 Frozen Orbit Mission Trajectory and Orbit Details.......................................................17

2.5.3 Frozen Orbit Details.......................................................................................................17

2.5.4 Selected Design Frozen Orbit ........................................................................................18

2.5.5 Frozen Orbit Lunar Coverage ........................................................................................18

2.5.6 Frozen Orbit Solar Coverage .........................................................................................23

2.6 Mission Description—Earth/Moon Lagrange Orbits.................................................................24

2.6.1 Earth/Moon Lagrange Orbit Mission Analysis Assumptions........................................24

2.6.2 Earth/Moon Lagrange Orbit Mission Analysis Analytic Methods ................................24

2.6.3 Earth/Moon Lagrange Orbit Mission Analysis Event Timeline ....................................25

2.6.4 Earth/Moon Lagrange Orbit Delta V Assumptions........................................................25

2.6.5 Earth/Moon Lagrange Orbit Mission Trajectory Details...............................................26

2.6.6 Earth/Moon Lagrange Orbit Lunar Coverage ................................................................26

2.7 Launch Vehicle (LV) Details.....................................................................................................26

2.7.1 LV Data and Trades.......................................................................................................26

2.7.2 LV Performance and Cost..............................................................................................26

2.7.3 LV Details—Frozen Orbit Constellation .......................................................................27

2.7.4 LV Details—Lagrange Orbit Constellation ...................................................................28

2.8 S/C Pointing...............................................................................................................................28

2.8.1 Frozen Orbit S/C Pointing..............................................................................................28

2.8.2 Lagrange Orbit S/C Pointing..........................................................................................33

2.9 System Design Trade Space ......................................................................................................34

3.0 Baseline Design..................................................................................................................................34

3.1 Top Level Design (MEL and PEL)—Microwave LNS.............................................................34

3.1.1 Master Equipment List (MEL)—Microwave LNS (Frozen Orbit)................................34

3.1.2 Power Equipment List (PEL)—Microwave LNS (Frozen Orbit) ..................................35

3.2 Top Level Design (MEL and PEL)—Optical LNS (Frozen Orbit)...........................................36
NASA/TM—2012-217139 iv

3.2.1 MEL—Optical LNS (Frozen Orbit)...............................................................................36

3.2.2 PEL—Optical LNS (Frozen Orbit)................................................................................37

3.3 Top Level Design (MEL and PEL)—Microwave LNS (Lagrange Orbit).................................38

3.3.1 MEL—Microwave LNS (Lagrange Orbit) ....................................................................38

3.3.2 PEL—Microwave LNS (Lagrange Orbit)......................................................................39

3.4 Top Level Design (MEL and PEL)—Optical LNS (Lagrange Orbit).......................................40

3.4.1 MEL—Optical LNS (Lagrange Orbit)...........................................................................40

3.4.2 PEL—Optical LNS (Lagrange Orbit)............................................................................41

3.5 System Level Summary.............................................................................................................42

3.5.1 Frozen Orbit RF (Microwave) Concept Configuration..................................................42

3.5.2 Lagrange Orbit RF Concept Configuration....................................................................44

4.0 Conclusions, Lessons Learned, and Areas for Future Study..............................................................45

4.1 Conclusions ...............................................................................................................................45

4.2 Areas for Future Study...............................................................................................................47

5.0 Subsystem Breakdown .......................................................................................................................48

5.1 Microwave Communications.....................................................................................................48

5.1.1 Microwave Communications Requirements..................................................................48

5.2 Microwave Communications Assumptions...............................................................................48

5.2.1 Microwave Communications Design and MEL.............................................................49

5.2.2 Microwave Communications Analytical Methods.........................................................50

5.2.3 Microwave Communications Risk Inputs......................................................................53

5.2.4 Microwave Communications Recommendation ............................................................53

5.3 Optical Communications...........................................................................................................53

5.3.1 Optical Communications Requirements ........................................................................53

5.3.2 Optical Communications Assumptions..........................................................................53

5.3.3 Optical Communications Design and MEL...................................................................53

5.3.4 Optical Communications Trades....................................................................................55

5.3.5 Optical Communications Analytical Methods...............................................................55

5.4 Command and Data Handling (C&DH) ....................................................................................55

5.4.1 C&DH Requirements.....................................................................................................55

5.4.2 C&DH Assumptions......................................................................................................55

5.4.3 C&DH Design and MEL................................................................................................55

5.4.4 C&DH Trades................................................................................................................57

5.4.5 C&DH Risk Inputs.........................................................................................................57

5.5 Guidance, Navigation and Control (GN&C).............................................................................57

5.5.1 GN&C Requirements.....................................................................................................57

5.5.2 GN&C Assumptions......................................................................................................58

5.5.3 GN&C Design and MEL................................................................................................59

5.5.4 GN&C Trades................................................................................................................59

5.5.5 GN&C Analytical Methods............................................................................................59

5.6 Electrical Power System............................................................................................................60

5.6.1 Power Requirements......................................................................................................60

5.6.2 Power Assumptions........................................................................................................60

5.6.3 Power Design and MEL.................................................................................................60

5.6.4 Power Trades..................................................................................................................61

5.6.5 Power Analytical Methods.............................................................................................61

5.6.6 Power Recommendation ................................................................................................62

5.7 Structures and Mechanisms.......................................................................................................62

5.7.1 Structures and Mechanisms Requirements ....................................................................62

5.7.2 Structures and Mechanisms Assumptions......................................................................62

5.7.3 Structures and Mechanisms Design and MEL...............................................................62
NASA/TM—2012-217139 v

5.7.4 Structures and Mechanisms Analytical Methods...........................................................64

5.7.5 Structures and Mechanisms Risk Inputs........................................................................64

5.7.6 Structures and Mechanisms Recommendation ..............................................................64

5.8 Propulsion and Propellant Management....................................................................................64

5.8.1 Propulsion and Propellant Management Requirements.................................................64

5.8.2 Propulsion and Propellant Management Assumptions...................................................64

5.8.3 Propulsion and Propellant Management Design and MEL............................................64

5.8.4 Propulsion and Propellant Management Analytical Methods........................................67

5.8.5 Propulsion and Propellant Management Risk Inputs.......................................................68

5.8.6 Propulsion and Propellant Management Recommendation ...........................................68

5.9 Thermal Control ........................................................................................................................68

5.9.1 Objective ........................................................................................................................68

5.9.2 Thermal Requirements...................................................................................................68

5.9.3 Thermal Assumptions....................................................................................................69

5.9.4 Thermal Design and MEL..............................................................................................69

5.9.5 Thermal Trades..............................................................................................................70

5.9.6 Thermal Analytical Methods..........................................................................................70

5.9.7 Thermal Risk Inputs.......................................................................................................73

5.9.1 Thermal Recommendation.............................................................................................74

6.0 Software Cost Estimation...................................................................................................................74

6.1 Objectives..................................................................................................................................74

6.2 Assumptions ..............................................................................................................................74

6.3 Approach ...................................................................................................................................75

7.0 Cost, Risk and Reliability...................................................................................................................77

7.1 Costing.......................................................................................................................................77

7.2 Risk Analysis and Reduction.....................................................................................................81

7.2.1 Assumptions...................................................................................................................81

7.2.2 Risk List and Summary..................................................................................................81

7.2.3 Trade Space Iterations....................................................................................................82

7.3 Case 1—Frozen RF....................................................................................................................84

7.4 Case 2—Frozen Optical.............................................................................................................85

7.5 Case 3—Lagrange RF................................................................................................................86

7.6 Case 4—Lagrange Optical.........................................................................................................87

Appendix A.—Acronyms and Abbreviations.............................................................................................89

Appendix B.—Rendered Images ................................................................................................................93

Appendix C.—LNS Lagrange Orbit Rendered Design Drawings.............................................................97

Appendix D.—SNAP Input File for Frozen Orbit Trajectory Propagation Analysis.................................99

Appendix E.—Study Participants.............................................................................................................101

References.................................................................................................................................................102

NASA/TM—2012-217139 1

COMPASS Final Report:

Lunar Network Satellite-High Rate (LNS-HR)

Steven R. Oleson and Melissa L. McGuire

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

1.0 Executive Summary

Two design options were explored to address the requirement to provide lunar piloted missions with 

continuous communications for outpost and sortie missions. Two unique orbits were assessed, along with 

the appropriate spacecraft (S/C) to address these requirements (Figure 1). Both constellations (with only 

two S/C each) provide full time coverage (24 hr/7 d) for a south polar base and also provide continuous 7 

day coverage for sorties for specified sites and periodic windows (for a top level comparison see Table 1).

Thus a two-satellite system can provide full coverage for sorties for selected windows of opportunity 

without reconfiguring the constellation.

1.1 Frozen Orbit Case

The first design was a Lunar Network Satellite (LNS) system using a frozen orbit around the Moon. 

A frozen orbit has orbit parameters that change very little over time and therefore do not require stationkeeping propellant (Figure 2). For the Moon, frozen orbits exist in which inclination and right ascension 

oscillate between acceptable limits. Extensive work was done using Satellite Orbit Analysis Program 

(SOAP) to assess: frozen orbit stability, coverage times, and S/C pointing and stabilization. The final 

design consists of two S/C evenly spaced in a 45° inclined, 24 hr, 0.4 eccentric nearly identical orbits. No 

major station keeping is needed for the orbits because they are in slightly different altitudes (5 km) to 

offset secular drift caused by Earth perturbations. The S/C will be Earth facing to eliminate previous 

designs needs for flipping the S/C to access Earth twice a month. The solar array (SA) axis is parallel to 

the Moon’s axis. The two, 1 m lunar dishes will swing on an arm once an orbit to provide voice and high 

data rate communications for future crewed lunar outposts and sortie sites. The system would also be 

capable of servicing various science users. An optical trunk communications link back to Earth has also 

been made. The baseline design will launch two LNS S/C together on an Atlas 511 launcher and provide 

communications and navigation for 12 yr. 

1.2 Lagrange Orbit Case

The second design utilized Earth/Moon L1 and L2 Lagrange orbits. A large selection of Lagrange 

orbits exist that could host lunar communications S/C. A ‘figure eight’ halo orbit at L1 and one at L2 were 

chosen for the design but further optimization may find more optimal orbits. The selected 16 day orbits 

would require around 60 m/s station keeping ∆V. Due to the requirement of providing communications to 

a 1 W astronaut suit radio combined with the distance of the S/C to the Moon, two 7-m dishes are 

required on the S/C. These will necessarily be deployable dishes similar to the early Tracking and Data 

Relay Satellite System (TDRSS) S/C. While an approach to launch both on an Atlas 521 launch vehicle 

would reduce launch costs, the cost for the two S/C network is 25 percent higher than the frozen orbit 

design due to the need for substantial station keeping propellant and large, deployable antennas. This 

report attempts to compare the two designs, along with an optical communications Earth link option.
NASA/TM—2012-217139 2

Figure 1.—LNS-HR frozen orbit (left) and Lagrange orbit (right) designs.

Figure 2.—An artist’s rendition of the frozen orbit case.

TABLE 1.—LNS-HR FROZEN ORBIT VERSUS 

LAGRANGE ORBIT DESIGN TOP LEVEL COMPARISON

Orbit type Mass 

(kg) Dimensions Life cycle cost ($M)

Frozen Orbit (RF technology) 1490 5- by 10-m 899

Lagrange Orbit (RF technology) 2280 17.5- by 13-m 1,073

Table 2 collects the details of the subsystems at a top level in the frozen orbit Lunar Network Satellite 

(LNS) design using microwave technology for the communications links to the Moon and back to Earth.
NASA/TM—2012-217139 3

This is also referred to as radio frequency (RF) communications throughout this report. All data are for a 

single satellite. Two satellites will be used in the constellation.

Table 3 collects the details of the subsystems at a top level in the Lagrange orbit LNS design. Note that 

some of the subsystems are the same as the design for the Frozen orbit RF case since they are unaffected by 

the choice of orbit. All data are for a single satellite. Two satellites will be used in the constellation.

TABLE 2.—MISSION AND S/C SUMMARY—FROZEN ORBIT RF SINGLE SATELLITE

Subsystem area Details Total mass with growth

Top level system S/C to relay lunar communications and provide navigation support 1489 kg with growth per 

each satellite

Mission, Operations, 

Guidance, Navigation and 

Control (GN&C)

Two S/C in 24 hr, 45° inclined frozen elliptical lunar orbit: Minimal station-keeping

12 yr life, each LNS single fault tolerant

Attitude Control System 

(ACS)

14 N-m-s momentum storage in four reactions wheels

12 Sun sensors, two star trackers, two inertial measurement units (IMUs) 51 kg

Launch Atlas 511: >25% launch margin

Dual launch capability

3915 kg performance to 

C3 = –2 km2/s2

Power

1100 W average power load (including 30% margin)

One single axis SA, 28% efficient triple-junction cells, 5.9 m2 area, 1600 W beginning of 

life (BOL)

Li-ion batteries sized for 2.5 hr lunar eclipse at full power, reduced ops for infrequent 7 

hr Earth eclipses: 4.3 kW-hr

108 kg

Propulsion Pressure fed hydrazine, two 100 lbf and sixteen 1 lbf thrusters 61 kg

Structures and mechanisms Al-Li panel around a central thrust-tube, thrust-tube designed to carry second S/C above 

for launch 288 kg

Communications

Provides continuous voice communications to all users including suit radios

400 Mbps data from users, 50 Mbps from other lunar surface user using space-based 

router

Support two-way ranging to up to five users simultaneously

138 kg

Command and Data 

Handling (C&DH) 

Systems command, control, health management

24 hr communications storage array (0.3 TB) 97 kg

Thermal Heat pipe—radiator system, heaters and multilayer insulation (MLI) 70 kg

TABLE 3.—MISSION AND S/C SUMMARY—LAGRANGE ORBIT CASE RF SINGLE SATELLITE DETAILS

Subsystem area Details Total mass with growth

Top level system S/C to relay lunar communications and provide navigation support 2281 kg with growth 

per each satellite

Mission, Operations, 

GN&C

ACS 94 kg

Launch Atlas 521: >25% launch margin

Dual launch capability

4710 kg performance to 

C3 = –2 km2/s2

Power 125 kg

Propulsion 132 kg

Structures and 

mechanisms 

Al-Li panel around a central thrust-tube, thrust-tube designed to carry second S/C above 

for launch 363 kg

Communications

Provides continuous voice communications to all users including suit radios

400 Mbps data from users, 50 Mbps from other lunar surface user using Space Based 

router

Support two-way ranging to up to five users simultaneously

248 kg

C&DH Systems command, control, health management

24 hr communications storage array (0.3 TB) 97 kg

Thermal Heat pipe—radiator system, heaters and MLI 70 kg
NASA/TM—2012-217139 4

Figure 3.—Lunar architecture surface communications links.

2.0 Study Background and Assumptions

2.1 Introduction 

COMPASS was engaged by the NASA Space Communications and Navigation (SCaN) office to 

provide a LNS concept designs as a follow-on for the Lunar Relay Satellite (LRS) design completed May 

2007. The COMPASS team began the design of a LNS and constellation to provide voice and high data 

rate communications for future crewed lunar outposts and sortie sites (Figure 3). The system would also 

be capable of servicing various science users. Frozen lunar orbits and Lagrange orbits were traded in 

order to minimize propulsion requirements. A trade on microwave versus optical trunk communications 

back to Earth was also made.

A NASA Request for Information (RFI) was recently put out to the space communications 

community to solicit studies on the future technologies and ideas to accomplish the communications goals

of NASA. 

http://www.nasa.gov/home/hqnews/2008/aug/HQ_08207_RFI_Lunar_Comm.html

MEDIA ADVISORY: 08-207

NASA Seeks Input For Commercial Lunar Communications & Navigation

WASHINGTON -- NASA issued a Request for Information, or RFI, on Monday to gauge interest and 

solicit ideas from private companies in providing communications and navigation services that would 

support the development of exploration, scientific and commercial capabilities on the Moon over the next 

25 yr.

Responses should be submitted to Barbara Adde, NASA Headquarters, Mail Suite 7L70, 300 E. St., SW, 

Washington, D.C. 20546-0001, by 4 p.m. EDT on Sept. 15, 2008. 
NASA/TM—2012-217139 5

To view the Request for Information, visit: http://www.spacecomm.nasa.gov

2.2 Assumptions

2.2.1 Coverage Assumptions/Requirements

NASA’s Level 0 Requirement to the SCaN Program for communications and navigation coverage of 

the lunar surface is as follows: SCaN shall provide anytime/anywhere communication and navigation 

services as needed for lunar and Mars human missions. To satisfy this requirement with a two-satellite 

constellation of LNS, the system design will strive to provide continuous S-band coverage for a selected 

list of potential outpost and sortie sites. The expectation is to support two sites simultaneously—including 

the outpost continuously and another sortie site for 7 to 14 days. The selected list of sortie and outpost 

sites is described in Figure 4 and Table 4. Additionally, the LNS system will strive to provide periodic 

(8 hr/12 hr) high data rate coverage for the selected list of potential outpost and sortie sites. Table 5 lists 

the study assumptions and requirements and identified potential technology trade space. The third column 

is the areas of technology trades that could be done for each of the subsystems.

Figure 4.—Lunar Architecture SCaN Program Proposed Landing Sites

TABLE 4.—SELECTED LIST OF POTENTIAL LUNAR OUTPOST AND SORTIE SITES

Landing Site Latitude Longitude Notes

A South Pole 89.9 S 180.0 W Rim of Shackleton

B Far side SPA floor 54.0 S 162.0 W Near Bose

C Orientale basin floor 19.0 S 88.0 W Near Kopff

D Oceanus Procellarum 3.0 S 43.0 W Inside Flamsteed P

E Mare Smythii 2.5 S 86.5 W Near Peek

F W/NW Tranquilitatis 8.0 N 21.0 E North of Arago

G Rima Bode 13.0 N 3.9 W Near Bode vent system

H Aristarchus plateau 26.0 N 49.0 W North of Cobra Head

I Central far side highlands 26.0 N 178.0 E Near Dante

J North Pole 89.5 N 91.0 E Rim of Peary B
NASA/TM—2012-217139 6

TABLE 5.—STUDY ASSUMPTIONS

Subsystem 

area

Assumptions and study requirements Critical trades

Top-Level

Two LRS provide comm/nav for piloted lunar missions, 2018 and 2020 launches. 12 yr 

mission for dual launch, all lunar surface elements (South Pole outpost site focused), 

but, also global coverage as necessary, robotic, rover, hab, in situ resource utilization 

(ISRU), Solar Power Unit/Makeup Power Unit (SPU/MPU), LCT RF power limit 40 to 

80 W (determines antenna size of LRS), LCT 1 m Ka antenna, rover RF power limit 10 

W and antenna size 0.3 m Ka band (modulation limited 2 bits/Hz)

12 hr Frozen orbit versus L1/ L2

(two S/C at each) for each 

configuration, S-Band and KaBand coverage circles, 7 yr option 

for separate launch approach

System

Figures of merit (FOMs): Cost, mass, flexibility, antenna sweep speed and excursion 

(LRS and lunar assets), Beam spot size of LRS

Off-the-shelf (OTS) equipment where possible, Push technology to save mass/cost, 

Technology Readiness Level-6 (TRL-6) cutoff 2014, single fault tolerant , Mass 

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

Communications

LRS to Earth: Ka band (400 Mbps mono, 800 Mbps dual), S-band (150 kbps), LRS to 

LCT: S-Band (150kbps), Ka band (400 Mbps,1200 Mbps), No cross-link with other 

LRS, gimbaled To Earth (110°): LRS to surface links: Ka 23/26 (25 Mbps) and TT&C 

S band, Processing for comm/nav in payload, storage in C&DH

Concentric ka/S downlink 

antenna(s), phase-array, trade LRS 

antennas, omni-S/Ka band 

antennas, number of Ka band 

antennas

Mission, Ops, ACS 18 m Earth antennas (3), constant one-way nav signal, two-way nav signal to Altair 

(descent/ascent)

12 hr frozen versus L2 orbits 

disposal (is a ∆V of 50 m/s 

sufficient)

Launch vehicle 

(LV) Atlas 501 or Delta IV equiv. Launch Loads: ~5gs Dual launch manifest option

Power 

1000 W triple junction gallium arsenide (GaAs), battery for (frozen orbit: 6 hr max, 2 

hr avg. Earth eclipse ~once a year for L2), 1 hr lunar eclipse (frequency), 30% power 

margin

Trade eclipse outage coverage

C&DH 

(including 

software)

Store of data (0.3 TB, single fault tolerant), storage, command and telemetry (C&T), 

nav, avionics processor boxes, 54 krad parts, –20 to 50 °C, 100 W

Redundant boxes, amount of 

processing/software between 

communications and C&DH 

packages

Thermal and 

Environment

Lowest temperature, (frozen orbit 6 hr Earth eclipse) , louvers, ~0.5 m radiator with 

heat pipes, 54 krad behind 100 mil Al –20 to 50° C

Mechanical

Q Earth pt. antenna (110° gimbal(s), booms), K-, S-band LCT pointing antenna (dual 

axis, 90° from zenith gimbal where necessary), space frame (Al Li), 

thermal/micrometeoroid and orbital debris (MMOD)/dust cover. Antenna 

mount/deployment systems or fixed (launch loads) Secondary: 4% of stage 

components

Redundant gimbals, box location, 

configuration, antenna/boom 

locations, interface with Lunar 

Lander.

2.2.2 Communications Assumptions/Requirements

• No crosslinks to other LNS-HR

2.2.2.1 LNS-HR Microwave Trunk Link

• Ka-band spectrum: 37/40 GHz; bandwidth up to 500 MHz

• S-band spectrum: 2.2/2.0 GHz

• 400 Mbps from LNS-HR to Earth at Ka-band

• 100 Mbps from Earth to LNS-HR at Ka-band

• 150 kbps from LNS-HR to Earth at S-band

• Bit error rate: 10–8

• One dual Ka-band/S-band trunk link antenna

• Radiometric tracking of LNS-HR from Earth at Ka-band (37/40 GHz)

• Command, control, telemetry, contingency voice, and safing on S-band

• Modulation limited to 2 bits/Hz

2.2.2.2 LNS-HR Optical Trunk Link

• 1.2 Gbps from LNS-HR to Earth

• 25 Mbps from Earth to LNS-HR
NASA/TM—2012-217139 7

• Packet loss rate: 10–9

• Optometric tracking of LNS-HR from Earth

• S-band spectrum: 2.2/2.0 GHz

• Command, control, telemetry, contingency voice, and safing on S-band

• S/C attitude stability: 0.3°

• S/C attitude knowledge: 1/10 of stability metric

2.2.2.3 LNS-HR Microwave Access Links

• Ka-band spectrum: 26/23 GHz

• S-band spectrum: 2.2/2.0 GHz

• Modulation limited to 2 bits/Hz

• 400 Mbps maximum from LNS-HR to lunar surface at Ka-band

• 150 Kbps maximum from LNS-HR to lunar surface at S-band

• Two Dual Ka-band/S-band access link antennas

• Navigation, TT&C, contingency voice, and safing on S-band link

2.2.2.4 Lunar Surface Assets

• Lunar Communications Terminal (LCT) antenna size: 1 m

• Rover antenna size: 0.3 m

• Extra Vehicular Activity (EVA) antenna: dipole

• EVA maximum transmit power: 1 W

2.2.3 Earth Based Telescopes for Laser Optical Trunk

• Station locations: Madrid, Canberra, Goldstone, three additional geographically separated sites

○ Antenna sizes: 18 m

• Aperture sizes: ~2 m (in order to preserve the same S/C pointing capability as RF)

• Transmit power: 5 W (TBR)

• Photon counting receivers: (TBD)

• State-of-the art (SOA) differential phase-shift keying-return-to-zero (DPSK-RZ) modem with 

132 ms forward error correction (FEC)/interleaver

○ Most likely Pulse Position Modulation (PPM) instead of DPSK-RZ

• Tracking: 5 μrad coarse tracking, < 1 μrad fine tracking TBR

• Laser safety coordination: FAA and Laser Clearing House

• Near Sun pointing: 10° (TBR)

• Atmospheric monitoring: 15 dB fading (TBR)

• Data handling 2.5 Gbps (TBR); Boeing also has 10 and 40 Gbps archiving, and distribution due 

to high data rates

○ 10 to 20 Mbps up to the Moon

2.2.4 Derived Requirements

The following requirements are derived from the Level 0 constellation mission requirements. The 

wording choice of “shall” in the statements indicates that the statement is a requirement.

• The LNS System shall provide continuous voice coverage for the selected list of potential outpost 

and sortie sites.

○ Expected to support two sites continuously and simultaneously for 7 days

– Outpost and another sortie site

○ Satellites may be moved but must complete move in < 3 months
NASA/TM—2012-217139 8

○ It is assumed that no LNS-HR voice coverage is required for sites in view of the Earth

• The LRS system shall provide periodic (8 hr/12 hr) high data rate (400 Mbps/500 MHz BW) 

coverage for the list of potential outpost and sortie sites.

○ 400 Mbps for all RF system

○ 1.2 Gbps for optical (space trunk) system

• The LRS system shall provide communications to/from astronauts on EVA during contingency 

operations 

○ 1 W max RF power from suit

○ Equivalent isotropically radiated power (EIRP) of –0.6 dB-W

2.3 Similar Past Missions

The LNS needs to be able to point to three (and perhaps four places) places simultaneously: the Sun 

(SA), the Earth (0.5 m trunk antenna), and down to the surface for lunar sites (perhaps up to two). The 

COMPASS design of a LRS performed in the spring of 2007 and documented in report CD–2007–12, 

designed a lunar relay satellite in support of human lunar base requirements. The requirements of this 

study have been further expanded since that study. The LRS design was used as one of the starting points 

for this study. In addition, a survey was done of current and recent examples of S/C that point to three 

things simultaneously. These examples also used as references in this study include:

• TDRSS (Figure 5)

○ Geosynchronous Earth orbit (GEO), points to Sun, Earth and objects to sides (other TDRSS)

• Hubble

○ Points to stars for long periods, SAs and two communications antennas 

• Lunar Reconnaissance Orbiter (LRO) (Figure 6)

○ Polar orbit

○ Data return when on near side of Earth only

○ Earth antenna on long boom, dual gimbaled

○ Two axis gimbaled SA

Figure 5.—TDRSS Satellite.
NASA/TM—2012-217139 9

Figure 6.—An artists’ rendition of the LRO mission.

2.4 Margin and Growth Policy

In order to clear up the margin and contingency (growth) policy of the MEL modeling, it helpful to 

define what is meant by the terms margin and contingency in the workings of the COMPASS design 

session.

2.4.1 Definition of Terms

For the purpose of the COMPASS designs, margin is the mass that is not flown (i.e., ∆V does is not 

applied to this mass in the rocket equation) to the final destination. This is most often referred to as LV 

margin. See Section 2.7 for a definition and further description of LV margin as applied to COMPASS 

designs. The margin percentage value is a per mission calculation. In some missions, the launch vehicle 

chosen to fly the mission will require more margin as a percentage of published launch vehicle 

performance than in others based on the propulsion system and mission risk(s).

For COMPASS, margin is not the same as Contingency or growth mass. Contingency or growth mass 

is a factor applied to the best estimate mass of a component in the S/C design, in order to bridge the gap 

between the current technology readiness level of that component and the final flight technology maturity 

mass of that component. This contingency is applied to the best guess mass to give a total mass 

calculation. It is this total mass that is flown to the mission’s final destination and used with the rocket 

equation and mission ∆Vs to calculate propellant mass. Each subsystem applies a growth percentage to 

each line item in their subsystem MEL. The system integration sheet rolls up these growth masses and 

calculate the overall S/C dry mass system contingency percentage (growth mass/total dry mass).

2.4.2 Definition of Mass Growth Allowance (MGA)

The COMPASS team uses the AIAA S–120–2006, Standard Mass Properties Control for Space 

Systems (Ref. 1). Table 6 shows the percent mass growth separated into a matrix specified by level of 

design maturity and specific subsystem. MGA is defined as the predicted change to the basic mass of an 

item based on an assessment of the design maturity and fabrication status of the item, and an estimate of 

the in-scope design changes that may still occur.
NASA/TM—2012-217139 10

The schedule shown in Table 6 was used by the team as of September 2008. For most studies, the 

Major Category for all parts will be the E1 row, unless a component is significantly farther along in the 

technology development process.

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

design is calculated. The COMPASS design team designs to a total growth of 30 percent or less. An 

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

30 percent limit on the dry mass of the system. Note that for designs requiring propellant, growth in 

propellant is either carried in the propellant calculation itself or in the ∆V used to calculate the propellant 

required to fly a mission. 

Because the analysis performed by the COMPASS team has traditionally been in the pre-phase A 

study phase, things like customer reserve as illustrated in Figure 7, mass definition illustration, taken from 

the 2006 AIAA report and shown below are not taken into consideration in the design. A phase A design 

would begin by allocating customer reserve masses on top of the design developed in the pre-phase A 

study session. The purpose of a pre-phase A study is to come up with a design that will give the customer 

the Mass requirement or allowable mass as shown in Figure 7.

TABLE 6.—PERCENT MGA

Major category

Maturity code

Design maturity

(Basis for mass determination)

Percent MGA

Electrical/electronic 

components

StructureBrackets, clips, hardwareBatterySolar arrayThermal controlMechanismsPropulsionWire harnessInstrumentationECSLSS, crew systems

0-5 

kg

5-15 

kg

>15 

kg

E

1

Estimated

(1) An approximation based on rough sketches, 

parametric analysis, or undefined requirements 

(2) A guess based on experience (3) A value 

with an unknown basis or pedigree

30 25 20 25 30 25 30 25 25 25 55 55 23

2

Layout

(1) A calculation or approximation based on 

conceptual designs (equivalent to layout 

drawings) (2) Major modifications to 

existing hardware

25 20 15 15 20 15 20 20 15 15 30 30 15

C

3

Preliminary design

(1) Calculations based on a new design 

after initial sizing but prior to final structural 

or thermal analysis (2) Minor modifications 

of existing hardware

20 15 10 10 15 10 10 15 10 10 25 25 10

4

Released design

(1) Calculations based on a design after final 

signoff and release for procurement or 

production (2) Very minor modification of 

existing hardware (3) Catalog value

10 5 5 5 6 5 5 5 5 5 10 10 6

A

5

Existing hardware

(1) Actual mass from another program, 

assuming that hardware will satisfy the 

requirements of the current program with no 

changes (2) Values based on measured masses 

of qualification hardware

3 3 3 3 3 3 3 2 3 3 5 5 4

6 Actual mass

Measured hardware No mass growth allowance—use appropriate measurement uncertainty values

7 Customer furnished equipment or specification 

value

Typically a “not-to-exceed” value is provided; however, contractor has the option 

to include MGA if justified
NASA/TM—2012-217139 11

Figure 7.—Mass definition

Whether the total mass of the system designed by COMPASS becomes the requirement or the 

allowable mass depends on the requirements and goals of the customer in the outcome of the study. Some 

COMPASS studies provide the customer with the ceiling of what a certain launch vehicle would be able 

to send to a preferred destination. Other studies, with different goals, will give the customer various 

options of using different launch vehicles and provide the size and capabilities of the missions that can be 

performed if limitations are put on the launch vehicle, propulsion system, and/or reductions are made. The 

scope of allowable technology options in anything from launcher to propulsion system to power system 

will impact the total mass of the system achieved at the end of the design study.

2.4.3 From the AIAA Standard—Terms and Definitions as applied to Dry Mass. 

The following definitions apply only to hardware and are taken from the AIAA 2006 document 

(Ref. 2). Appendix C also shows a series of MGA tables used in various NASA programs.

Allowable mass The limits against which margins are calculated after accounting for basic 

masses of flight hardware, MGA, and uncertainties 

NOTE: Derived from the requirements early in the program, the allowable mass 

is intended to remain constant for the duration of the program. 

Basic mass The current mass data based on an assessment of the most recent baseline 

design.

NOTE 1: This design assessment includes the estimated, calculated, or measured 

(or actual) mass, and includes an estimate for undefined design details like 

cables, multi-layer insulation and adhesives. 

NOTE 2: The MGA and uncertainties are not included in the basic mass. 

COMPASS has referred to this as current best estimate (CBE) in past mission 

design.

Mass The measure of the quantity of matter in a body 

MGA The predicted change to the basic mass of an item based on an assessment of the 

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

design changes that may still occur 
NASA/TM—2012-217139 12

Predicted mass The basic mass plus the mass growth allowance

This is also referred to as the total mass by the COMPASS team.

Mass margin The difference between the space system allowable mass and predicted mass

For the COMPASS process, the total percentage on dry mass 

Mass Margin + Predicted Mass = CBE + 30 percent*CBE.

Therefore, mass margin will be carried as a percentage of dry mass rather 

than as a hard-targeted number based on launch vehicle performance, etc.

Because these studies are pre-Phase A and conceptual, unless the customer has 

other requirements for this total growth on the system, the 30 percent number is 

standard COMPASS operating procedure.

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

requirements in modeling the power system. See Sections 3.1.2, 3.2.2, 3.3.2, 3.4.2 

for the power system assumptions.

2.4.4 Launch Margin

Margin in terms of a LV from the S/C design perspective is that part of the LV performance that 

cannot be transferred to the S/C. This is uncertainty in the performance calculations. It is taken off the top 

of the launch vehicle’s performance to the target orbit or C3 (energy). The LV to S/C adaptor mass (i.e., a 

part left with the launch vehicle after separation) is typically taken out of this margin for the unusable LV 

performance capability. A conservative margin of 10 percent has been the usual rule of thumb. It not 

flown on the S/C, and not used in the inert mass calculations applied to trajectory parameters when sizing 

propellant requirements for the mission (but the adapter is part the full up LV mass of the pad and effects 

fuel, and performance). For this level of analysis, the COMPASS team uses a 10 percent factor on the 

quoted launch performance as margin.

Launch Margin = 10% * Specified LV Performance

The launch margin is treated two ways in a design. It is either an established parameter that cannot be 

traded, or it is used as a parameter to establish the closure or goodness of a design. If a S/C design allows 

for better than 10 percent launch margin, it is considered to be a robust design. If it allows for less than 10

percent margin, then more work must be done, and the design is not closed.

2.4.5 Power Growth

The COMPASS team uses a 30 percent margin on the bottoms up power requirements in modeling 

the power system. See Section 5.6 for the power system assumptions and design.

2.5 Mission Description—Frozen Orbits

Any initial orbit around a central body will begin to change primarily due to the nonuniform gravity 

field of the central body and 3rd body effects. Some characteristics/orbit elements of the initial orbit can 

be fixed, i.e. frozen, or controlled by specifying values and/or rates of change for the other orbit elements. 

The goal, in theory, is to place two S/C into the same orbit, but located 180° apart. In order to maintain 

even coverage of the south-pole, the idea is to freeze the eccentricity and semi-major axis (SMA). For 

operation, it was discovered that a slight difference in SMA between the two orbits would keep the 

satellites a constant 180° apart in their orbits.

Several authors note the existence of a family of “Frozen” orbits around the Moon for a specific 

combination of orbital elements (see Refs. 10, 11, and 12). This family of orbits includes inclinations 

above 39.6°, arguments of perigee of 90° or 270°, and eccentricities (e) that correspond to inclinations (i) 

by the equation 
NASA/TM—2012-217139 13

cos ( ) 1

3

5 2 2 e + i =

For eccentricity of a circular orbit, e = 0.

= ≅ °

=

=

39.23

5

3 cos( )

cos ( ) 1

3

5

0

2

i

i

e

It is easy to see from this equation that a circular orbit is ‘frozen’ at 39.6° with eccentricity increasing 

with inclination. In actuality these orbits do oscillate in eccentricity and inclination but repeat on a cycle 

of years. Thus, no station-keeping is needed if the extremes of inclination and eccentricity still meet the 

S/C’s coverage requirement.

Despite these circulations such orbits are very advantageous for single polar coverage with the 

combination of arguments of perigee of 90° or 270° and high ellipticity. Thus an outpost at the South Pole 

would benefit from a high inclination and an argument of perigee of 90°. For example two S/C in a 12 hr

orbit, spaced 180° apart with an inclination of 39° and SMA (9750.3 km and 9755.3 km offset to handle 

drift perturbations) gives good coverage of a South Pole outpost. However, such orbits do not provide 

good coverage of the northern regions of the Moon.

The requirements for this design study are to provide both continuous South Pole coverage and 

continuous sortie coverage, especially for the far side sites. Assuming that only one sortie occurs at any 

one time and that it is of a limited duration (less than 2 wk), a new frozen orbit could be designed which 

could provide continuous coverage of both the South Pole outpost and a specific far side sortie site for the 

2 wk period. This can be achieved by using a higher frozen orbit (24 hr period) at a lower inclination to 

allow for a higher perigee to cover both the South Pole and the sortie site for at least a 2 wk period. The 

orbit inclination selected was 45° and an eccentricity of 0.4. The tool SOAP was used to ensure full 

coverage of each of the desired sortie sites during Sunlit periods (for power needs).

In addition to the oscillations up and down of the eccentricity and inclination over time the SMA also 

changes slightly up and down over time. For a single S/C this is of no issue but for a constellation of two 

or more, their spacing apart in an orbit will change and they will no longer maintain an equal spacing in 

argument of perigee. Thus the S/C could drift together and coverage would no longer be continuous.

However, raising the SMA of one of the satellites slightly 5 km over the other has been shown to prevent 

this drifting of SMA from affecting the relative spacing of the two S/C. The orbits are maintainable 

without any orbit maintenance propellant necessary over the 12-yr lifetime requirement.

2.5.1 Frozen Orbit Mission Analysis Assumptions

While the orbits are ‘Frozen’ they do oscillate in SMA and eccentricity, but this oscillation is cyclic 

and repeats every ~ 2 yr. The effects on the orbit are:

• Argument of periapsis varies between 80° and 115°

• Eccentricity (between 0.35 and 0.5) will increase altitude and then decrease it again

• Inclination (between 39° and 55°) will change coverage

2.5.2 Frozen Orbit Mission Analysis Analytic Methods 

Propagation of the orbits over time was done by two methods using two different trajectory analysis 

programs used by the COMPASS team. Orbit was designed using SOAP and checked with Spacecraft NBody Analysis Program (SNAP).
NASA/TM—2012-217139 14

The first method was via SOAP’s low thrust propagation modeling capability. The orbital elements were 

input into SOAP and a propagation of those elements over time was run out for the expected 12-yr mission.

SOAP is used extensively by COMPASS to visualize orbits and communications link analysis as well 

as Sun angle for power analysis over the lifetime of a satellite in its orbit. SOAP, shown if Figure 8, is an 

interactive three-dimensional orbit visualization and analysis program for Windows, Macintosh, and Sun 

workstations. The software is capable of simultaneously propagating thousands of satellites, ground 

stations, aircraft, ships and planets using a diverse set of prediction methods SOAP was developed in 

1988 internally by the Engineering Applications Department within the Aerospace Corporation. The 

purpose of SOAP was to provide visualization and analysis for Space missions. Users can also use 

SOAP’s internal predefined variable types to perform analysis on a range of parameters in an orbit or 

architecture. Users can also import computer aided design (CAD) models of S/C and use visualization to 

gather further knowledge of pointing requirements for the various components: Antenna, arrays, thrusters.

To corroborate the behavior of the frozen orbits that were being produced by SOAP, a comparison of 

a representative frozen orbit was performed with SNAP. Developed at the NASA Glenn Research Center

(GRC), SNAP is a high fidelity trajectory propagation program that can propagate the trajectory of a S/C

about virtually any body in the solar system. The equations of motion can include the effects of central 

body gravitation with N x N harmonics, other body gravitation with N x N harmonics, solar radiation 

pressure, atmospheric drag (for Earth orbits) and S/C thrusting. The equations are solved using an eighthorder Runge-Kutta Fehlberg single step method with variable step size control. The input deck for this 

case is shown in Appendix D.

The representative orbit was chosen to be a 24 hr circular lunar orbit with an initial inclination of 40°

and an initial right ascension of the ascending node (RAAN) of 50°. Including the Earth and the Sun as 

perturbing bodies, the orbits were propagated for 7 yr in both SNAP and SOAP and the time history of 

the orbital elements between the two programs was compared. As can be seen in Figure 9, while the 

initial comparison showed a similar trend in the behavior of the inclination, obviously the orbits produced 

by the programs differed greatly over time. 

It has been seen that there are many different definitions of inertial frames, especially when it comes 

to lunar orbits, and it was decided to compare the definitions used by SNAP and SOAP. Upon

investigation, it was discovered that the definition for the lunar inertial frame was indeed different, 

specifically the direction of the X-axis in inertial space. To obtain the direction of the X -axis for a Moon 

centered inertial frame, SNAP used the orbit normal of the Earth crossed with the Moon’s North Pole 

whereas SOAP used a [0,0,1] vector in inertial space crossed with the Moon’s North Pole. Since RAAN

is measured from this X-axis, the initial orbits of the two programs were not actually aligned in inertial 

space even though the same initial conditions were used. The definition of the lunar inertial frame used by 

SOAP was implemented in SNAP, and the results of this change can be seen in Figure 10.

Figure 8.—SOAP graphic.
NASA/TM—2012-217139 15

Figure 9.—Initial comparison of a representative frozen orbit between SNAP and SOAP.

Figure 10.—Final comparison of a representative frozen orbit between SNAP and SOAP.
NASA/TM—2012-217139 16

Figure 11.—LNS frozen orbit case CONOPs timeline.

After aligning the lunar inertial frames, the two programs showed very good agreement in the 

behavior of the representative frozen orbit.

2.5.3 Frozen Orbit Mission Analysis Event Timeline

Most of the ∆V is taken from the NASA Goddard Space Flight Center’s (GSFC) I M Design Center

(IMDC) study from the summer of 2006 for the sections of the mission.

Table 7 lists the ∆V assumptions used in sizing the propellant budget for performing the propulsion 

maneuvers throughout the frozen orbit mission. An additional allocation of 50 m/s has been added on for 

attitude control and station keeping maneuvers while in lunar orbit, and an additional 60 m/s has been 

added for end of life disposal of the S/C (Figure 11).

2.5.1 Frozen Orbit Delta V Schedule

TABLE 7.—LNS FROZEN ORBIT ∆V SCHEDULE

Phase name

OMS 

ΔV,

m/s

RCS 

ΔV,

m/s

Heritage

Launch from Earth 

Checkout

Loiter to TLI window opening TLI opening to ignition 

TLI burn 

Correct for ELV Dispersions 85 GSFC IMDC 

Trans-Earth mid-course corrections trans-lunar coast 25 GSFC IMDC 

Lunar orbit capture burn 410 GSFC IMDC 

Lunar orbit adjustments 250 GSFC IMDC 

Phasing maneuvers 50 GSFC IMDC 

Lunar telecom orbit 50 50 Added 2v margin for attitude control and station keeping 

Disposal 80 Estimate for disposal on lunar surface near Pericynthion 

Total ΔV 950 50
NASA/TM—2012-217139 17

Figure 12.—LNS frozen orbit LOI maneuvers.

2.5.2 Frozen Orbit Mission Trajectory and Orbit Details

The LNS is injected to TLI from the Atlas V launch vehicle. It departs Earth on a coast trajectory in 

the (same state injected into by the launch vehicle on TLI). It coasts to the Moon and at the Moon three 

burns are completed to achieve the final frozen orbit. The first burn is lunar orbit insertion 1 (LOI 1), with 

the highest ∆V of the three burns. LOI 1 is an in-plane burn whose function is to capture the S/C into a 

highly elliptical orbit about the Moon. LOI 2 performs two functions. It is both a high plane change burn 

and also raises periapse to result in the desired final inclination. Next, there is another coast back to 

periapse where LOI 3 is performed (Figure 12). This third burn adjusts apoapse after LOI 2 in order to 

ensure reaching the final orbital parameters of the desired frozen orbit.

Verification of the LOI ∆V was performed using Copernicus. Copernicus is a high precision system 

for trajectory design and optimization where trajectories are modeled as a series of segments and was 

developed at the University of Texas at Austin by Dr. Cesar Ocampo.

2.5.3 Frozen Orbit Details

COMPASS design recommends a 24 hr orbit to provide better visibility the Moon. The 24 hr period 

orbits provide a simple duty cycle for ground controllers and users.

Using a frozen orbit design with two S/C 12 yr of constant South Pole coverage is possible with no 

station keeping. The SMA includes the radius of the Moon at 1738 km. The eccentricity of 0.4 yields an 

orbit that varied in altitude above the surface.

TABLE 8.—LNS FROZEN ORBIT ORBITAL PARAMETERS

Frozen orbit variable Values

Date .............................................................................................................January 1, 2018, 12:00 am

SMA.........................................................9750.3 and 9755.3 km (offset to handle drift perturbations)

Eccentricity.............................................................................................................................0.408248

Inclination........................................................................................................................................45°

Argument of periapsis......................................................................................................................90°

RAAN..............................................................................................................................................50°

Mean anomaly ................................................................................................................135° and 315°

Figure 13 shows the two orbits described in Table 8 as illustrated using the trajectory tool, SOAP.

The orbit planes are inclined 45° to the lunar equator.
NASA/TM—2012-217139 18

Figure 13.—Frozen orbit two satellite constellation illustration.

Figure 14.—Frozen orbit orbital element perturbation history.

2.5.4 Selected Design Frozen Orbit

Constellation spaced 180° in mean anomaly to provide constant coverage. Orbits for S/C LNS A and 

LNS B do oscillate but repeat roughly every 3 yr. No orbit maintenance is required.

2.5.5 Frozen Orbit Lunar Coverage 

Requirements were established that 100 percent Sun visibility and constant coverage by at least one 

of the LNS-HR satellites was necessary at all times during long duration lunar outpost (South Pole) 

missions as well as 7 days continuous satellite communication and Sun visibility for sortie missions at the 
NASA/TM—2012-217139 19

sites listed in Table 4. Coverage analysis for the LNS-HR frozen orbit was performed to verify that the 

lunar outpost received continuous coverage and that opportunities existed at each of the sortie sites for a 

7-day mission.

The coverage analysis over a 4-wk period is presented in Figure 15. Four different Cartesian views of 

the lunar surface show contours of the percent coverage over a 7-day period. The color gradient bar 

corresponds to percentage coverage over each 7-day period with deep red representing 100 percent

coverage, meaning at least one of the LNS-HR satellites and the Sun is visible during the entire 7 days. 

The beginning of this particular 4-wk period corresponds to the single opportunity available at which the 

Central Far Side Highlands site has continuous coverage for a full week, beginning August 30, 2021,

10:00:00. A further discussion of sortie opportunities is presented later in this section. In each view there 

is a concentrated area that receives full coverage surrounded by areas of less coverage and then an area 

that is not covered at all during the timeframe. The high percentage areas are located on regions of the 

lunar surface that is beneath the apocynthion of the frozen orbit. Since the satellites remain in this 

segment of the orbit for the longest period of time, longer coverage times are possible. The frozen orbit is 

inclined such that the LNS-HR satellites, when in the area of the apocynthion, are able to view the South 

Pole of the Moon. This orientation results in the southern sites on the Moon being more easily viewed by 

the satellites and in extended periods of coverage to be more numerous and frequent. The area of 

concentrated coverage moves approximately 180° in 14 days as the Moon rotates beneath the frozen orbit. 

While maintaining constant coverage of the outpost, this allows for periods of high coverage ratios at 

most of the sortie sites. Coverage ratio is defined as the total time the Sun and at least one of the LNS-HR 

satellites are in view during the total time of the opportunity (total time in view/total time of opportunity).

Although there is constant satellite coverage of the South Pole, there are occurrences when Sun 

visibility is interrupted. When there is full Sun visibility, 100 percent coverage lasts on average for 

114 days. Figure 16 in which the opportunity for a 7-day 100 percent coverage mission to the Central Far 

Side Highlands is presented, the South Pole is being shaded for the entire month long period. Since the 

Central Far Side Highlands is the most northern sortie site examined, for the most part, longer periods of 

Sun visibility only occur when the position of the Sun is such that the South Pole received little to no 

visibility. However, there are limited opportunities for 7-day missions with above 99 percent coverage.

Most of the Moon is visible to the LNS-HR constellation, and the satellites are able to provide 

66 percent coverage to the surface for high data rate access. The Sun and at least one of the frozen 

satellites is in view for 7-day periods. The Central Far Side Highlands has 100 percent coverage for 

7 days beginning August 30, 2021, 10:00:00. Note in Figure 16 that the center of ridge moves ~180° in 14 

days. This means that most sortie sites could be serviced for >95 percent of time for 14 days.

See Section 2.9 for S/C pointing analysis.

The orbital period drift is offset by increasing the SMA of one satellite with respect to the other by 

5 km. Figure 17 shows the zenith view from selected lunar surface sites of interest as listed in Table 4.

The lunar South Pole has Continuous coverage from the LNS-HR constellation. The Far Side SPA has 

long periods of continuous coverage. The Central Far Highlands have daily coverage with at least one set 

of continuous 7-day coverage. The North Pole has periods of daily coverage.

Continuing the analysis of the sites of interest from Table 4 and in Figure 17 are the coverage ratios 

from the LNS-HR constellation to the lunar surface.

Several of the sortie sites from Table 4 were examined to determine the opportunity availability based 

on their Sun-satellite coverage ratios. Although it is desirable to have 100 percent coverage during the 

entire 7 day sortie mission for communication and data transfer capabilities, because near side and LIM 

sortie sites can use the Deep Space Network (DSN) for voice communication that 66.6 percent coverage 

(8 hr/12 hr ratio) would be acceptable to perform the necessary data transmissions. For the two far side 

site selected, Far Side SPA (South Pole-Aitken) Floor and Central Far Side Highlands, 100 percent

coverage is required because communication using the DSN network is not available. 
NASA/TM—2012-217139 20

Figure 15.—Lunar surface coverage over 1 month.

Figure 16.—Frozen orbit drift and South Pole coverage.
NASA/TM—2012-217139 21

Figure 17.—Zenith view from selected sites.

Figure 18.—Visibility coverage ratio of sortie sites.

The visibility ratios for all sortie opportunities 7-days or longer for the sites investigated are shown in 

Figure 18. Below is a summary of opportunities for each site:

• Far-side sites

○ Central Far Side Highlands
NASA/TM—2012-217139 22

– 544 sortie opportunities with visibility ratios greater than 66.6 percent (8 hr/12 hr)

♦ Maximum mission duration: 13.5 days

♦ Average mission duration: 7.8 days

– One opportunity with 100 percent Sun-sat visibility

○ Far Side SPA Floor

– 25 sortie opportunities, all have 100 percent Sun-sat visibility

♦ Maximum mission duration: 14.4 days

♦ Average mission duration: 10.7 days

• LIM-sites

○ South Pole

– 12 sortie opportunities, all have 100 percent Sun-sat visibility

♦ Maximum mission duration: 171.2 days

♦ Average mission duration: 114.2 days

○ North Pole

– No opportunities for a 7-day period having a visibility ratio > 66.6 percent (8 hr/12 hr)

○ Mare Smythii

– 226 sortie opportunities with visibility ratios > 66.6 percent (8 hr/12 hr)

♦ Maximum mission duration: 14.5 days

♦ Average mission duration: 8.9 days

– 31 opportunities that have 100 percent Sun-sat visibility

• Near-side sites

○ Aristarchus Plateau

– 553 sortie opportunities with visibility ratios > 66.6 percent (8 hr/12 hr)

♦ Maximum mission duration: 13.0 days

– Average mission duration: 7.7 days

– No opportunities with 100 percent Sun-sat visibility, highest visibility ratio is 0.9943

– 104 sortie opportunities with visibility ratios > 0.95 for 7-days

The LNS-HR Two-S/C Constellation fulfills 7-day Sortie requirement for far-side sites while 

providing 24/7 Outpost (South Pole) service. The constellation also provides alternate 8 /12 hr direct to 

Earth (DTE) link for voice and data for Near Side and LIM sites (except for North Pole).

The lack of 100 percent 7-day opportunities for Aristarchus Plateau and only one such opportunity for 

Central Far Side Highlands prompted an examination of whether 100 percent coverage could be achieved 

for slightly shorter sortie missions. Figure 2.17 shows the available coverage opportunities with durations 

greater than 2 days. Both 100 percent Sun and at least one of the LNS-HR Satellites are visible during 

these opportunities.

• Aristarchus Plateau

○ Forty-one 100 percent opportunities

– Five opportunities with durations longer than 6 days

– Maximum duration is 6.39 days

• Central Far Side Highlands

○ Forty 100 percent opportunities

– One opportunity with duration 7 days or more (7.12)

○ Fourteen opportunities with durations between 6 and 7 days
NASA/TM—2012-217139 23

Figure 19.—One hundred percent coverage of sortie sites.

TABLE 9—SUN SHADE DURATION STATISTICS

Maximum ............................................................6.91 hr

Minimum...........................................................0.047 hr

Average .............................................................1.036 hr

Figure 20.—Satellite Sun shading durations for frozen orbit satellites.

2.5.6 Frozen Orbit Solar Coverage 

During the lifetime of the frozen orbit communications satellite, there are 569 Sun-Shade cycles. The 

periodically occurring short-duration Sun shading periods are caused by the LNS satellite being shaded by 

the Moon. These shading durations are less than 2 hr long, averaging slightly over an hour (Table 9 and 

Figure 20. On occasion both the Moon and the Earth shade the S/C resulting in longer shade durations. 
NASA/TM—2012-217139 24

Figure 21.—Lagrange system illustration.

2.6 Mission Description—Earth/Moon Lagrange Orbits

The Lagrange points are the positions around a two body system where the gravitational forces of the 

two large masses (such as the Earth-Moon system or the Earth-Sun system) equal the centripetal force 

required to rotate with them. These provide relatively stable orbits where another, smaller body, may orbit 

and have the same stationary view of the two body system over time. These orbits will require 

stationkeeping propulsion to maintain the position over sometimes short and sometimes long periods of 

time (Refs. 8 and 9)

The L1 point of the Earth-Sun system affords an uninterrupted view of the Sun and is currently home 

to the Solar and Heliospheric Observatory (SOHO) Satellite. The L2 point of the Earth-Sun system is 

home to the Wilkinson Microwave Anisotropy Probe (WMAP) S/C and (perhaps by the year 2011) the 

James Webb Space Telescope.

Figure 21 shows the five Lagrange points in a two-body system with one body far more massive than 

the other (e.g., the Sun and the Earth or the Earth and the Moon). In such a system, L3–L5 will appear to 

share the secondary body's orbit, although in fact they are situated slightly outside it.

2.6.1 Earth/Moon Lagrange Orbit Mission Analysis Assumptions

A representative Lagrange orbit was chosen from the literature at the time of the study. An actual 

Lagrange orbit desired from trajectory analysis was not found by the time of the S/C design and is 

ongoing. Two Lagrange orbits, one at L1 and one at L2, were chosen from Reference 12 in order to 

provide a starting point for the power and communications system design. Complete lunar South-Pole 

coverage can be obtained with two S/C in the constellation.

2.6.2 Earth/Moon Lagrange Orbit Mission Analysis Analytic Methods

Due to the limited time for this study, rather than running an independent Lagrange trajectory, the 

Lagrange Orbit selection is from the technical report “Multibody Orbit Architectures for Lunar South 

Pole Coverage” Grebow, et al. Grebow looks for an orbit that would provide optimal South 

Pole/Shackelton crater coverage. The “optimal” was defined as continuous South Pole—and 7 day sortie 

site line of site coverage. Best Lagrange constellation option so far consists of a Vertical L1 and a Vertical 

L2 16 day halo orbit each. This is the only family that provides ~100 percent South Pole (North too) but 

also can provide extended (~7 day) coverage of sorties in northern lunar hemisphere.
NASA/TM—2012-217139 25

2.6.3 Earth/Moon Lagrange Orbit Mission Analysis Event Timeline

Figure 22.—Lagrange orbit constellation CONOPS.

2.6.4 Earth/Moon Lagrange Orbit Delta V Assumptions

The ∆V values used in Table 10 were taken from recent analysis published in technical literature on 

station-keeping for and orbit capture into Lagrange orbits. Approximately 500 m/s was allocated for orbit 

insertion with approximately 60 m/s/yr of station keeping for 12 yr, yields 720 m/s total station keeping 

∆V. With larger launch vehicle for Lagrange orbits, propellant loading is, again, not a driving design 

parameter.

TABLE 10.—LAGRANGE ORBIT ΔV SCHEDULE BY PHASE

Phase Name OMS 

∆V 

(m/s)

RCS 

∆V, 

m/s

Heritage

Launch from Earth 

Checkout 

Loiter to TLI window opening 

TLI opening to ignition 

TLI burn 

Correct for ELV dispersions 85 GSFC IMDC 

Trans-Earth mid-course corrections trans-lunar coast 25 GSFC IMDC 

Lunar orbit capture burn 500 Gordon Thesis, Purdue University 

Lunar orbit adjustments 50 Estimate 

Phasing maneuvers 720 Folta, Vaughn Paper, AIAA 2004-4741 

Lunar telecom orbit 50 50 Added 2v margin for attitude control and station keeping 

Disposal 80 Estimate for disposal on lunar surface near Pericynthion 

Total ΔV 1550 50
NASA/TM—2012-217139 26

Figure 23.—Manifold insertion ∆V for short time of flight transfers to

halo orbits between Az=1,000 and Az = 10,000 km.

2.6.5 Earth/Moon Lagrange Orbit Mission Trajectory Details

Table 11 captures the parameters used in assuming the Lagrange Orbit for the second type of the 

lunar communication satellite constellation examined in this study. Most of these parameters were 

obtained from literature.

TABLE 11.—LAGRANGE ORBIT PARAMETERS

Lagrange orbit parameters Value

L1 and L2 Orbits........................................................................................................16 day

Station keeping.................................................................................................. ~60 m/s/yr

C3...................................................................................................~ -2 km2

/s2 lunar centric

Lagrange Points orbit insertion ............................................................................~500 m/s

Distance to lunar users .....................................................................................~72,000 km

L1 and L2 S/C ....................................................................................................Dual launch

2.6.6 Earth/Moon Lagrange Orbit Lunar Coverage 

See Section 2.9 for S/C pointing analysis.

2.7 Launch Vehicle (LV) Details

2.7.1 LV Data and Trades

Table 12 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. The range on the Delta 

IV medium is related to the choice of payload fairing diameter.

TABLE 12.—LAUNCH SERVICE COSTS FROM 

THE NASA KENNEDY SPACE CENTER (KSC)

ELV 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

0 to 4,711 to 5,400 kg ~$224M Delta IV Medium +

2.7.2 LV Performance and Cost

Assumptions:

• Performance range of the ELV was based on a targeted C3 = –2 km2

/s2 (lunar). 

• Used a launch date of 1 January 2018 and an eastern range.
NASA/TM—2012-217139 27

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

Strategy for the Evolved Expendable Launch Vehicles (EELV) Launch Service. Note that this is a 

slightly reduced amount of conservatism based on some recent acquisitions.

• All costs are estimated in real-year dollars (2008) 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 LSP.

• 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 (USAF) infrastructure cost allocations, the 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.

2.7.3 LV Details—Frozen Orbit Constellation

The Atlas V 401 was initially baselined as the launch vehicle choice for the LNS-HR mission based 

on the LRS design. The Atlas V provides ~66 percent mass margin for LNS-HR (i.e., it could launch 

almost three LNS-HR) 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 24

which 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. After sizing the LNS-HR S/C from bottoms 

up, in order to fit two LNS satellites onto one launch vehicle, the frozen orbit cases required the Atlas V 

511 performance (the blue line in Figure 24).

C3 (km2/sec2)

Figure 24.—Atlas V 501, 401, 511 performance curve for C3 = –2 km2/sec2 (lunar).
NASA/TM—2012-217139 28

Figure 25.—Atlas V 511 fairing dimensions with notional two LNS launch packaging.

Figure 25 shows how two LNS-HR Frozen orbit satellites would fit onto the Atlas 511 launch 

vehicle.

2.7.4 LV Details—Lagrange Orbit Constellation

After sizing the LNS-HR S/C from bottoms up, in order to fit two LNS satellites onto one launch 

vehicle, the Lagrange orbit constellation required the Atlas V 521 performance (Figure 26 and Figure 27).

2.8 S/C Pointing

2.8.1 Frozen Orbit S/C Pointing

Four systems on the S/C must point continuously to four different locations (Figure 28)

• Lunar antennas to two separate sites

• Earth antenna to Earth

• SA to Sun

The following are a series of S/C to target pointing options to provide pointing between the 

following: data from lunar surface to LNS, from LNS to Earth and from LNS SAs to the Sun. In all cases, 

the pointing refers to the communications system and how it is pointed.

• Point S/C at Moon

• Point S/C at Sun

• Point S/C at Earth
NASA/TM—2012-217139 29

Performance plot results:

Important note: The data contained in these curves are based on ground rules and assumptions located below the plot. Please read this information 

carefully. This information is intended for NASA customers only.

NASA ELV Performance Curve(s)

High Energy Orbits

Please note the ground rules and assumptions below

C3 (km2/sec2)

Assumptions:

Atlas V (521)

 This performance does not include the effects of orbital debris compliance, which must be 

evaluated on a mission-specific basis. This could result in a significant performance impact 

for mission in which launch vehicle hardware remains in Earth orbit.

 3-sigma mission required margin, plus additional reserves as determined by the LSP.

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

 Performance values assume harness, logo, reradiating antenna, three payload fairing doors.

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

accommodations.

 Type B2 payload adapter plus type C2 spacer.

 5-m Short Payload Fairing

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

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

 Performance shown is applicable to declinations between 

28.5° and –28.5°.

Last updated: May 22, 2006, 8:11:53 am.

Figure 26.—Atlas V 521 performance curve for C3 = –2 km2/sec2 (lunar).

Figure 27.—Lagrange constellation S/C inside Atlas V 521 fairing.
NASA/TM—2012-217139 30

Figure 28.—Frozen orbit constellation pointing notional illustration.

Past designs have pointed the S/C at the Moon and provided an antenna on an arm to point at the 

Earth. At issue was the flipping that the S/C would need to perform to point the Earth antenna above or 

below the orbit plane based on the relation of the orbit plane to the Earth. Such flipping maneuvers would 

require propellant and would incur risks. One way to eliminate this flipping maneuver would be to point 

the S/C to the Earth. This has an additional benefit since the Sun is always very close to the Earth’s orbit 

plane when viewed from an orbit around the Moon. Such an approach would require very little pointing 

for the Earth antenna, the SA will only require a single gimbal that will make one revolution each month

(Figure 29). The lunar pointing antennas regardless of approach need to point to different portions of the 

Moon. With the Earth pointing S/C approach the antennas will need to rotate once a day and be able to 

point up and down 45° from the orbit plane. 

The proposed pointing of LNS for the frozen orbit is summarized as follows

• Point 0.5 m antenna mounted on the S/C body at Earth 

• Single (one axis gimbal) SA perpendicular to Moon/Earth orbit plane

• Point lunar antennas down/up to Moon (±45° elevation) and rotate about their axis 360° per orbit 

(once a day)

• The S/C will not have to flip over (as with past designs)

• GN&C

○ Satellite held in ‘fixed’ pointing to Earth

○ Lunar gravity gradient torques may be an issue (but should be periodic and can be taken out 

by wheels)

Since the Moon’s orbit is only inclined ~6° from ecliptic, for the Moon, the Earth and the Sun are 

roughly in the same plane when viewed from an orbit about the Moon.

• Pointing approach

○ Align SA axis parallel to lunar axis, rotate SA about one axis of the S/C body once a month

○ Point the antenna mounted on the body of the S/C at Earth continuously

○ Rotate lunar antenna boom once an orbit

○ Gimbal the lunar antennas to individual lunar users

○ Eliminates S/C flip or additional Earth antenna from previous designs
NASA/TM—2012-217139 31

Figure 29.—Frozen orbit pointing notional illustration for single satellite.

The S-band and a Ka-band antenna are pointed at the lunar surface and rotate in order to continuously 

point to a desired site on the surface under the LNS as long as that point is in sight as shown in Figure 30. 

For example, the antenna points to the South Pole site (purple beam), until that site has gone out of range 

of the LNS through its orbit. A second LNS in the same orbit will pick up the line of site coverage with 

the South Pole as the first LNS goes out of range.

Figure 31 shows that the SAs rotate independently of the antenna in order to track the Sun. The 

yellow beam is the notional modeling of the line of site between the SA panels and the Sun. The purple 

beam is the antenna pointed at the South Pole of the lunar surface. The yellow beam representing viewing 

of the Sun from the S/C can also be seen in Figure 30. When looking from the point of view of the SAs, 

the Sun, the plane of the area of the SAs is perpendicular to the vector connecting the Sun to the arrays.

From the graphic in Figure 30, the yellow line goes off to the top left where the Sun would be located 

with respect to the Moon.

The DTE antenna is located on the body of the LNS S/C bus, and is set to point at the Earth. This 

antenna can be seen in Figure 32 as the dish located on the Earth facing portion of the S/C bus. The Solar 

array is shown as the blue rectangular panel located at the bottom of the S/C bus. The purple beam is the 

link between the antenna and the lunar South Pole and the green beam is the link between the second 

antenna and a site chosen from the list of lunar surface sites of interest in Table 4.
NASA/TM—2012-217139 32

Figure 30.—Frozen orbit pointing SOAP modeling.

Figure 31.—Frozen orbit pointing modeled in SOAP.
NASA/TM—2012-217139 33

Figure 32.—Frozen orbit pointing modeled in SOAP.

Figure 33.—Lagrange orbit pointing notional diagram.

2.8.2 Lagrange Orbit S/C Pointing

Options considered for pointing

• Earth pointing S/C

• Moon Pointing S/C
NASA/TM—2012-217139 34

Having the S/C point parallel to chosen Earth-Moon vector was easier on both Earth antenna and the 

SAs. The 7 m lunar antennas need to gimbal ~ 160° along the y-axis. The S/C was slewed about the z-axis 

to point to Moon. For this maneuver, the 0.5 m Earth antenna needed to gimbal ~ 15° in z and y-axis

(Figure 33).

2.9 System Design Trade Space

Two major trades were made in the design session. First, the trade was made between constellation 

orbits: frozen circular lunar orbits versus Lagrange orbit constellations. Second, the trade was made in 

each constellation on type of communications technology: microwave (RF) versus Optical.

3.0 Baseline Design 

3.1 Top Level Design (MEL and PEL)—Microwave LNS

3.1.1 Master Equipment List (MEL)—Microwave LNS (Frozen Orbit)

Table 13 lists the top level of the MEL of the design with all the subsystem line elements hidden such 

that only the top level masses are shown. The total growth on the dry mass of the S/C is then rolled up to 

find a total growth mass and growth percentage. Engineers enter in the CBE mass for each of their line 

elements, as well as quantity. Then the Growth column is where each subsystem lists the recommended 

growth factor on each line items following the AIAA WGA schedule outlined in Table 6 in Section 2.4.

The MEL takes all of the items and racks them up into totals and calculates a total CBE mass, a Total 

mass and a total growth mass.

Where the MEL (Table 13) captures the bottoms up estimation of CBE and growth percentage line 

item by item from the subsystem designer, Table 15 wraps up 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. This additional system level mass is assumed as part of the inert 

mass that is flown along the required trajectory. Therefore, the additional system level growth mass 

impacts the total propellant loading for the mission design.

TABLE 13.—MASTER EQUIPMENT LIST—LNS RF FROZEN ORBIT CASE

WBS 

no.

Description

LNS-HR (September 2008) Frozen, RF

Quantity Unit

mass, 

kg

Basic

mass, 

kg

MGA, 

%

Growth,

kg

Predicted

mass,

kg

6.0 Lunar Network Satellite-High Rate - - 1288.1 11 142.8 1430.9

6.0.1 RF Communications Package - - 123.1 12 14.9 138.0

6.0.2 Optical Communications Package - - 0.0 0 0.0 0.0

6.0.3 C&DH - - 77.2 26 20.2 97.4

6.0.4 GN&C - - 42.5 20 8.5 51.0

6.0.5 Structures & Mechanical Systems - - 233.7 23 54.7 288.4

6.0.6 Power System - - 83.0 30 24.9 107.9

6.0.10 Propulsion (Chemical) - - 10.6 10 1.1 11.7

6.0.11 Propellant Management (Chemical) - - 43.9 13 5.7 49.5

6.0.12 Propellant (Chemical) - - 617.6 0 0.0 617.6

6.0.13 Thermal Control (Non-Propellant) - - 56.5 23 13.0 69.5
NASA/TM—2012-217139 35

TABLE 14.—SYSTEM INTEGRATION—LNS RF FROZEN ORBIT CASE

3.1.2 Power Equipment List (PEL)—Microwave LNS (Frozen Orbit)

The power listing for nominal loads is shown in Table 15. 

• Power system tracks its own power

• Power system must provide:

○ ~1250 W (with 30 percent growth) during communications operations

○ ~1000 W (with 30 percent growth) during S/C shadowing (seasonal) assuming Ka not used

• Generated thermal waste heat includes:

○ ~700 W (with 30 percent growth) during communications operations

○ ~550 W (with 30 percent growth) during S/C shadowing (seasonal) assuming Ka not used

GLIDE 

container: LunarRelaySat: LNS_Frozen_RF

Spacecraft Master Equipment List Rack-up (Mass)

COMPA SS S/C 

Design 

WBS Main Subsystems

Basc (CBE) 

Mass (kg)

Growth 

(kg)

Predicted (Total) 

Mass (kg)

Aggregate 

Growth (%)

6.0 Lunar Network Satellie-High Rate 1288.1 142.8 1430.9

6.0.1 RF Communications Package 123.1 14.9 138.0 12%

6.0.2 Optical Communications Package 0.0 0.0 0.0 TBD

6.0.3 Command & Data Handling (C&DH) 77.2 42.5 97.4 55%

6.0.4 Guidance, Navigation and Control System (GN&C)42.5 8.5 51.0 20%

6.0.5 Structures & Mechanical Systems 233.7 54.7 288.4 23%

6.0.6 Power System 83.0 24.9 107.9 30%

6.0.7 Propulsion (Electric) 0.0 0.0 0.0 TBD

6.0.8 Propellant Management (EP) 0.0 0.0 0.0 TBD

6.0.9 Propellant (EP) 0.0

6.0.10 Propulsion (Chemical) 10.6 1.1 11.7 10%

6.0.11 Propellant Management (Chemical) 43.9 5.7 49.5 13%

6.0.12 Propellant (Chemical) 617.6

6.0.13 Thermal Control (Non-Propellant) 56.5 13.0 69.5 23%

Estimated Spacecraft Dry Mass 670 143 813 21%

Estimated Spacecraft Wet Mass 1288 165 1431

Estimated Spacecraft Inert Mass (for traj.) 689 201 889.7

System LeveL Growth Calculations Total Growth

Dry Mass Desired System Level Growth 670 201 871.6 30%

Mass Margin (carried a t system level) 58 9%

Total We t Mass with Growth 1288 201 1489.2

Available Launch Perform ance to C3 (kg) 3915.0 kg

Total Num ber o f LNS-HR Sa tellites 2

Total m ass o f Two LNS-HR sa tellites 2978.5 kg

Launch m argin available (kg) 936.5 kg

Launch m argin available (%) 23.9 %

TABLE 15.—PEL—FROZEN ORBIT—RF

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

Launch 10 m 20 25 20 5 48 0.0 0.0 118 35.34 153 

S/C checkout 24 h 75 152 589 30 112 0.0 0.0 958 287.43 1246 

Cruise 4 d 75 152 20 30 112 0.0 0.0 389 116.73 506 

Cruise Communications 75 152 387 30 112 0.0 0.0 756 226.83 983 

Orbit Injection 30 m 75 152 387 30 168 0.0 0.0 812 243.63 1056 

Lunar Sunlit Communications Continuous 75 152 589 30 112 0.0 0.0 958 287.43 1246 

Lunar Eclipse Communications 2h/d (max.) 75 152 387 60 112 0.0 0.0 786 235.83 1022 

Waste Heat

Launch 0.0 25 10.0 0.0 32.0 0.0 0.0 66.5 20.0 86.5 

S/C checkout 0.0 152 294.5 0.0 92.0 0.0 0.0 538.6 161.6 700.2 

Cruise 0.0 152 10.0 0.0 92.0 0.0 0.0 254.1 76.2 330.3 

Cruise Communications 0.0 152 193.5 0.0 92.0 0.0 0.0 437.6 131.3 568.9 

Orbit Injection 0.0 152 193.5 0.0 92.0 0.0 0.0 437.6 131.3 568.9 

Lunar Sunlit Communications 0.0 152 294.5 0.0 92.0 0.0 0.0 538.6 161.6 700.2 

Lunar Eclipse Communications 0.0 152 193.5 0.0 92.0 0.0 0.0 437.6 131.3 568.9 
NASA/TM—2012-217139 36

3.2 Top Level Design (MEL and PEL)—Optical LNS (Frozen Orbit)

3.2.1 MEL—Optical LNS (Frozen Orbit)

Table 16 lists the top level of the MEL the frozen orbit case using optical communications as well as 

microwave. Figure 17 wraps up those total masses, CBE and total mass after applied growth percentage 

for the frozen orbit optical communications case.

TABLE 16.—MEL—FROZEN ORBIT—OPTICAL

WBS 

no.

Description

LNS-HR (September 2008) Frozen, Optical

Quantity Unit

mass, 

kg

Basic

mass, 

kg

MGA,

%

Growth, 

kg

Predicted

mass, 

kg

6.0 Lunar Network Satellite-High Rate - - 1261.9 11 144.2 1406.1

6.0.1 RF Communications Package - - 92.3 11 10.5 102.8

6.0.1.1 Ka band (40/37) DTE Link - - 0.0 0 0.0 0.0

6.0.1.2 Ka band (23/26) Relay Link - - 15.5 10 1.6 17.1

6.0.1.3 Ka band (40/37) TT&C and Ranging - - 0.0 0 0.0 0.0

6.0.1.4 S band Relay Link - - 24.4 10 2.4 26.8

6.0.1.5 Data Router and SSR - - 4.0 10 0.4 4.4

6.0.1.6 Antennas (Booms excluded) - - 48.4 13 6.1 54.5

6.0.2 Optical Communications Package - - 18.1 32 5.8 23.8

6.0.2.1 Optical Head - - 6.7 34 2.3 9.0

6.0.2.2 Electronics Box - - 10.8 30 3.2 14.0

6.0.2.3 Optical Head-to-E.Box Interface - - 0.5 55 0.3 0.8

6.0.2.4 Interface to Spacecraft - - 0.1 30 0.0 0.1

6.0.3 C&DH - - 77.2 26 20.2 97.4

6.0.3.1 C&DH Hardware - - 74.2 25 18.5 92.7

6.0.3.2 Instrumentation & Wiring - - 3.0 55 1.7 4.7

6.0.4 GN&C - - 42.5 20 8.5 51.0

6.0.4.1 GN&C Hardware - - 42.5 20 8.5 51.0

6.0.5 Structures & Mechanical Systems - - 225.0 23 52.5 277.5

6.0.5.1 Primary Structures - - 199.1 25 49.8 248.9

6.0.5.2 Secondary Structures - - 7.3 25 1.8 9.1

6.0.5.3 Installation - - 3.6 25 0.9 4.4

6.0.5.4 Mechanisms - - 15.0 0 0.0 15.0

6.0.6 Power System - - 83.0 30 24.9 107.9

6.0.6.1 Battery System - - 43.0 30 12.9 55.9

6.0.6.2 Solar Array - - 25.0 30 7.5 32.5

6.0.6.3 Power Management & Distribution - - 15.0 30 4.5 19.5

6.0.10 Propulsion (Chemical) - - 10.6 10 1.1 11.7

6.0.10.1 Main Engine - - 4.6 10 0.5 5.1

6.0.10.2 Reaction Control Sys tem - - 6.0 10 0.6 6.6

6.0.11 Propellant Management (Chemical) - - 51.5 15 7.8 59.2

6.0.11.1 OMS Propellant Management - - 0.0 0 0.0 0.0

6.0.11.2 RCS Propellant Management - - 51.5 15 7.8 59.2

6.0.12 Propellant (Chemical) - - 605.4 0 0.0 605.4

6.0.12.1 Main Engine Propellant - - 0.0 0 0.0 0.0

6.0.12.2 RCS Propellant - - 605.4 0 0.0 605.4

6.0.13 Thermal Control (Non-Propellant) - - 56.5 23 13.0 69.5

6.0.13.1 Active Thermal Control - - 4.2 21 0.9 5.1

6.0.13.2 Passive Thermal Control - - 41.4 23 9.5 50.9

6.0.13.3 Semi-Passive Thermal Control - - 10.9 24 2.6 13.5
NASA/TM—2012-217139 37

TABLE 17.—MEL—FROZEN ORBIT—OPTICAL

3.2.2 PEL—Optical LNS (Frozen Orbit)

The power listing for nominal loads is listed in Table 18.

• Power system tracks its own power

• Power system must provide:

○ ~1250 W (with 30 percent growth) during communications operations

○ ~1000 W (with 30 percent growth) during S/C shadowing (seasonal) assuming Ka not used

• Generated thermal waste heat includes:

○ ~700 W (with 30 percent growth) during communications operations

○ ~550 W (with 30 percent growth) during S/C shadowing (seasonal) assuming Ka not used

TABLE 18.—PEL—FROZEN ORBIT—OPTICAL Notional Time Propulsion,

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

Launch 10 m 20 76 20 5 48 0.0 0.0 169 50.79 220 

S/C checkout 24 h 75 150 589 30 112 0.0 0.0 956 286.65 1242 

Cruise 4 d 75 150 20 30 112 0.0 0.0 387 115.95 502 

Cruise Communications 75 150 387 30 112 0.0 0.0 754 226.05 980 

Orbit Injection 30 m 75 150 387 30 168 0.0 0.0 810 242.85 1052 

Lunar Sunlit Communications Continuous 75 150 589 30 112 0.0 0.0 956 286.65 1242 

Lunar Eclipse Communications 2 h/d (max.) 75 150 387 60 112 0.0 0.0 784 235.05 1019 

Waste Heat

Launch 0.0 76 10.0 0.0 32.0 0.0 0.0 118.0 35.4 153.4

S/C checkout 0.0 150 294.5 0.0 92.0 0.0 0.0 536.0 160.8 696.8

Cruise 0.0 150 10.0 0.0 92.0 0.0 0.0 251.5 75.5 327.0

Cruise Communications 0.0 150 193.5 0.0 92.0 0.0 0.0 435.0 130.5 565.5

Orbit Injection 0.0 150 193.5 0.0 92.0 0.0 0.0 435.0 130.5 565.5

Lunar Sunlit Communications 0.0 150 294.5 0.0 92.0 0.0 0.0 536.0 160.8 696.8

Lunar Eclipse Communications 0.0 150 193.5 0.0 92.0 0.0 0.0 435.0 130.5 565.5

COMPASS study: Lunar Network Satellite - High Rate (LNS-HR) Study Date 9/22/08

GLIDE 

container: LunarRelaySat: LNS_Frozen_Optical

Spacecraft Master Equipment List Rack-up (Mass)

COMPA SS S/C 

Design 

WBS Main Subsystems

Basc (CBE) 

Mass (kg)

Growth 

(kg)

Predicted (Total) 

Mass (kg)

Aggregate 

Growth (%)

6.0 Lunar Network Satellie-High Rate 1261.9 144.2 1406.1

6.0.1 RF Communications Package 92.3 10.5 102.8 11%

6.0.2 Optical Communications Package 18.1 5.8 23.8 32%

6.0.3 Command & Data Handling (C&DH) 77.2 42.5 97.4 55%

6.0.4 Guidance, Navigation and Control System (GN&C)42.5 8.5 51.0 20%

6.0.5 Structures & Mechanical Systems 225.0 52.5 277.5 23%

6.0.6 Power System 83.0 24.9 107.9 30%

6.0.7 Propulsion (Electric) 0.0 0.0 0.0 TBD

6.0.8 Propellant Management (EP) 0.0 0.0 0.0 TBD

6.0.9 Propellant (EP) 0.0

6.0.10 Propulsion (Chemical) 10.6 1.1 11.7 10%

6.0.11 Propellant Management (Chemical) 51.5 7.8 59.2 15%

6.0.12 Propellant (Chemical) 605.4

6.0.13 Thermal Control (Non-Propellant) 56.5 13.0 69.5 23%

Estimated Spacecraft Dry Mass 657 144 801 22%

Estimated Spacecraft Wet Mass 1262 167 1406

Estimated Spacecraft Inert Mass (for traj.) 674 197 871.3

System LeveL Growth Calculations Total Growth

Dry Mass Desired System Level Growth 657 197 853.5 30%

Mass Margin (carried a t system level) 53 8%

Total We t Mass with Growth 1262 197 1458.9

Available Launch Perform ance to C3 (kg) 3915.0 kg

Total Num ber o f LNS-HR Sa tellites 2

Total m ass o f Two LNS-HR sa tellites 2917.8

Launch m argin available (kg) 997.2
NASA/TM—2012-217139 38

3.3 Top Level Design (MEL and PEL)—Microwave LNS (Lagrange Orbit)

3.3.1 MEL—Microwave LNS (Lagrange Orbit)

Table 19 lists the top level of the MEL the Lagrange orbit case using optical communications as well 

as microwave. Figure 20 wraps up those total masses, CBE and total mass after applied growth 

percentage for the Lagrange orbit microwave communications case.

TABLE 19.—MEL—LAGRANGE ORBIT—MICROWAVE

WBS 

no.

Description

LNS-HR (September 2008) Lagrange RF

Quantity Unit

mass,

kg

Basic mass,

kg

MGA,

%

Growth,

kg

Predicted

mass,

kg

6.0 Lunar Network Satellite-High Rate - - 2005.2 10 207.9 2213.1

6.0.1 RF Communications Package - - 204.1 22 44.1 248.2

6.0.1.1 Ka band (40/37) DTE Link - - 17.9 10 1.8 19.7

6.0.1.2 Ka band (23/26) Relay Link - - 15.4 10 1.5 16.9

6.0.1.3 Ka band (40/37) TT&C and Ranging - - 14.9 18 2.7 17.6

6.0.1.4 S band Relay Link - - 24.4 10 2.4 26.8

6.0.1.5 Data Router and SSR - - 4.0 10 0.4 4.4

6.0.1.6 Antennas (Booms excluded) - - 127.5 28 35.2 162.7

6.0.2 Optical Communications Package - - 0.0 0 0.0 0.0

6.0.2.1 Optical Head - - 0.0 0 0.0 0.0

6.0.2.2 Electronics Box - - 0.0 0 0.0 0.0

6.0.2.3 Optical Head-to-E.Box Interface - - 0.0 0 0.0 0.0

6.0.2.4 Interface to Spacecraft - - 0.0 0 0.0 0.0

6.0.3 C&DH - - 77.2 26 20.2 97.4

6.0.3.1 C&DH Hardware - - 74.2 25 18.5 92.7

6.0.3.2 Instrumentation & Wiring - - 3.0 55 1.7 4.7

6.0.4 GN&C - - 78.5 20 15.7 94.2

6.0.4.1 GN&C Hardware - - 78.5 20 15.7 94.2

6.0.5 Structures & Mechanical Systems - - 291.0 25 71.8 362.8

6.0.5.1 Primary Structures - - 248.3 25 62.1 310.4

6.0.5.2 Secondary Structures - - 9.4 25 2.4 11.8

6.0.5.3 Installation - - 18.3 25 4.6 22.9

6.0.5.4 Mechanisms - - 15.0 18 2.8 17.8

6.0.6 Power System - - 96.0 30 28.8 124.8

6.0.6.1 Battery System - - 43.0 30 12.9 55.9

6.0.6.2 Solar Array - - 35.0 30 10.5 45.5

6.0.6.3 Power Management & Distribution - - 18.0 30 5.4 23.4

6.0.10 Propulsion (Chemical) - - 18.7 10 1.9 20.5

6.0.10.1 Main Engine - - 9.3 10 0.9 10.3

6.0.10.2 Reaction Control System - - 9.3 10 0.9 10.2

6.0.11 Propellant Management (Chemical) - - 98.6 13 12.6 111.2

6.0.11.1 OMS Propellant Management - - 0.0 0 0.0 0.0

6.0.11.2 RCS Propellant Management - - 98.6 13 12.6 111.2

6.0.12 Propellant (Chemical) - - 1084.6 0 0.0 1084.6

6.0.12.1 Main Engine Propellant - - 0.0 0 0.0 0.0

6.0.12.2 RCS Propellant - - 1084.6 0 0.0 1084.6

6.0.13 Thermal Control (Non-Propellant) - - 56.5 23 13.0 69.5

6.0.13.1 Active Thermal Control - - 4.2 21 0.9 5.1

6.0.13.2 Passive Thermal Control - - 41.4 23 9.5 50.9

6.0.13.3 Semi-Passive Thermal Control - - 10.9 24 2.6 13.5
NASA/TM—2012-217139 39

TABLE 20.—MEL—LAGRANGE ORBIT—MICROWAVE LNS

3.3.2 PEL—Microwave LNS (Lagrange Orbit)

The following requirements were used in sizing the power for the Lagrange orbit constellation using 

RF communications (Table 21).

• Power system tracks its own power

• Power system must provide:

○ ~1200 W (with 30 percent growth) during communications operations

○ ~800 W (with 30 percent growth) during S/C shadowing (seasonal) assuming Ka not used

• Generated thermal waste heat includes:

○ ~850 W (with 30 percent growth) during communications operations

○ ~650 W (with 30 percent growth) during S/C shadowing (seasonal) assuming Ka not used

TABLE 21.—PEL—MICROWAVE LNS-HR LAGRANGE ORBIT—RF Notional Time Propulsion,

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

Launch 10 m 20 79 20 3 48 0.0 0.0 170 51.12 222 

S/C checkout 24 h 75 152 531 30 112 0.0 0.0 900 270.03 1170 

Cruise 2 d 75 152 20 30 112 0.0 0.0 389 116.73 506 

Cruise Communications 75 152 385 30 112 0.0 0.0 754 226.23 980 

Orbit Injection 1 h 75 152 385 30 168 0.0 0.0 810 243.03 1053 

Lunar Sunlit Communications Continuous 75 152 531 30 112 0.0 0.0 900 270.03 1170 

Lunar Eclipse Communications 2 h/d (max.) 75 152 203 60 112 0.0 0.0 602 180.63 783 

Waste Heat

Launch 0.0 79 10.0 0.0 32.0 150.0 0.0 270.7 81.2 351.9 

S/C checkout 0.0 152 265.5 0.0 92.0 150.0 0.0 659.6 197.9 857.5 

Cruise 0.0 152 10.0 0.0 92.0 150.0 0.0 404.1 121.2 525.3 

Cruise Communications 0.0 152 192.5 0.0 92.0 150.0 0.0 586.6 176.0 762.6 

Orbit Injection 0.0 152 192.5 0.0 92.0 150.0 0.0 586.6 176.0 762.6 

Lunar Sunlit Communications 0.0 152 265.5 0.0 92.0 150.0 0.0 659.6 197.9 857.5 

Lunar Eclipse Communications 0.0 152 101.5 0.0 92.0 150.0 0.0 495.6 148.7 644.3 

Spacecraft Master Equipment List Rack-up (Mass) COMPA SS S/C 

Design 

WBS Main Subsystems

Basc (CBE) 

Mass (kg)

Growth 

(kg)

Predicted (Total) 

Mass (kg)

Aggregate 

Growth (%)

6.0 Lunar Network Satellie-High Rate 2005.2 207.9 2213.1

6.0.1 RF Communications Package 204.1 44.1 248.2 22%

6.0.2 Optical Communications Package 0.0 0.0 0.0 TBD

6.0.3 Command & Data Handling (C&DH) 77.2 20.2 97.4 26%

6.0.4 Guidance, Navigation and Control System (G 78.5 15.7 94.2 20%

6.0.5 Structures & Mechanical Systems 291.0 71.8 362.8 25%

6.0.6 Power System 96.0 28.8 124.8 30%

6.0.7 Propulsion (Electric) 0.0 0.0 0.0 TBD

6.0.8 Propellant Management (EP) 0.0 0.0 0.0 TBD

6.0.9 Propellant (EP) 0.0

6.0.10 Propulsion (Chemical) 18.7 1.9 20.5 10%

6.0.11 Propellant Management (Chemical) 98.6 12.6 111.2 13%

6.0.12 Propellant (Chemical) 1084.6

6.0.13 Thermal Control (Non-Propellant) 56.5 13.0 69.5 23%

Estimated Spacecraft Dry Mass 921 208 1129 23%

Estimated Spacecraft Wet Mass 2005 208 2213

Estimated Spacecraft Inert Mass (for traj.) 952 276 1228.1

System LeveL Growth Calculations Total Growth

Dry Mass Desired System Level Growth 921 276 1196.8 30%

Mass Margin (carried a t system level) 68 7%

Total We t Mass with Growth 2005 276 2281.4

Available Launch Perform ance to C3 (kg) 4710.0 kg

Total Num ber o f LNS-HR Sa tellites 2

Total m ass o f Two LNS-HR sa tellites 4562.7

Launch m argin available (kg) 147.3 103
NASA/TM—2012-217139 40

3.4 Top Level Design (MEL and PEL)—Optical LNS (Lagrange Orbit)

3.4.1 MEL—Optical LNS (Lagrange Orbit)

Table 22 lists the top level of the MEL the Lagrange orbit case using optical communications as well 

as microwave. Table 23 wraps up those total masses, CBE and total mass after applied growth percentage 

for the Lagrange orbit optical communications case.

TABLE 22.—MEL—LAGRANGE ORBIT—OPTICAL

WBS no. Description

LNS-HR (September 2008) Lagrange , Optical

Quantity Unit

mass,

kg

Basic

mass,

kg

MGA, 

%

Growth,

kg

Predicted

mass,

kg

6.0 Lunar Network Satellite-High Rate - - 1915.4 11 206.3 2121.7

6.0.1 RF Communications Package - - 147.0 25 37.1 184.1

6.0.1.1 Ka band (40/37) DTE Link - - 0.0 0 0.0 0.0

6.0.1.2 Ka band (23/26) Relay Link - - 15.4 10 1.5 16.9

6.0.1.3 Ka band (40/37) TT&C and Ranging - - 0.0 0 0.0 0.0

6.0.1.4 S band Relay Link - - 12.6 10 1.3 13.9

6.0.1.5 Data Router and SSR - - 4.0 10 0.4 4.4

6.0.1.6 Antennas (Booms excluded) - - 115.0 29 33.9 148.9

6.0.2 Optical Communications Package - - 18.1 32 5.8 23.8

6.0.3 C&DH - - 77.2 26 20.2 97.4

6.0.3.1 C&DH Hardware - - 74.2 25 18.5 92.7

6.0.3.2 Instrumentation & Wiring - - 3.0 55 1.7 4.7

6.0.4 GN&C - - 78.5 20 15.7 94.2

6.0.4.1 GN&C Hardware - - 78.5 20 15.7 94.2

6.0.5 Structures & Mechanical Systems - - 289.3 25 71.3 360.6

6.0.5.1 Primary Structures - - 248.3 25 62.1 310.4

6.0.5.2 Secondary Structures - - 9.4 25 2.4 11.8

6.0.5.3 Installation - - 16.5 25 4.1 20.7

6.0.5.4 Mechanisms - - 15.0 18 2.8 17.8

6.0.6 Power System - - 96.0 30 28.8 124.8

6.0.6.1 Battery System - - 43.0 30 12.9 55.9

6.0.6.2 Solar Array - - 35.0 30 10.5 45.5

6.0.6.3 Power Management & Distribution - - 18.0 30 5.4 23.4

6.0.10 Propulsion (Chemical) - - 18.7 10 1.9 20.5

6.0.10.1 Main Engine - - 9.3 10 0.9 10.3

6.0.10.2 Reaction Control System - - 9.3 10 0.9 10.2

6.0.11 Propellant Management (Chemical) - - 98.6 13 12.6 111.2

6.0.11.1 OMS Propellant Management - - 0.0 0 0.0 0.0

6.0.11.2 RCS Propellant Management - - 98.6 13 12.6 111.2

6.0.12 Propellant (Chemical) - - 1035.6 0 0.0 1035.6

6.0.12.1 Main Engine Propellant - - 0.0 0 0.0 0.0

6.0.12.2 RCS Propellant - - 1035.6 0 0.0 1035.6

6.0.13 Thermal Control (Non-Propellant) - - 56.5 23 13.0 69.5

6.0.13.1 Active Thermal Control - - 4.2 21 0.9 5.1

6.0.13.2 Passive Thermal Control - - 41.4 23 9.5 50.9

6.0.13.3 Semi-Passive Thermal Control - - 10.9 24 2.6 13.5
NASA/TM—2012-217139 41

TABLE 23.—MEL—LAGRANGE ORBIT—OPTICAL LNS

3.4.2 PEL—Optical LNS (Lagrange Orbit)

• Power system tracks its own power

• Power system must provide (Table 24):

○ ~1200 W (with 30 percent growth) during communications operations

○ ~900 W (with 30 percent growth) during S/C shadowing (seasonal) assuming Ka not used

• Generated thermal waste heat includes:

○ ~850 W (with 30 percent growth) during communications operations

○ ~700 W (with 30 percent growth) during S/C shadowing (seasonal) assuming Ka not used

TABLE 24.—PEL—ORBITAL LNS-HR LAGRANGE ORBIT—OPTICAL Notional 

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

Launch 10 m 20 79 20 3 48 0.0 0.0 170 51.12 222 

S/C checkout 24 h 75 152 531 30 112 0.0 0.0 900 270.03 1170 

Cruise 2 d 75 152 20 30 112 0.0 0.0 389 116.73 506 

Cruise Communications 75 152 337 30 112 0.0 0.0 706 211.68 917 

Orbit Injection 1 h 75 152 337 30 168 0.0 0.0 762 228.48 990 

Lunar Sunlit Communications Continuous 75 152 531 30 112 0.0 0.0 900 270.03 1170 

Lunar Eclipse Communications 2h/day (max) 75 152 294 60 112 0.0 0.0 693 207.93 901 

Waste Heat

Launch 10 m 0.0 79 10.0 0.0 32.0 150.0 0.0 270.7 81.2 351.9

S/C checkout 24 h 0.0 152 265.5 0.0 92.0 150.0 0.0 659.6 197.9 857.5

Cruise 2 d 0.0 152 10.0 0.0 92.0 150.0 0.0 404.1 121.2 525.3

Cruise Communications 0.0 152 168.3 0.0 92.0 150.0 0.0 562.4 168.7 731.1

Orbit Injection 1 h 0.0 152 168.3 0.0 92.0 150.0 0.0 562.4 168.7 731.1

Lunar Sunlit Communications Continuous 0.0 152 265.5 0.0 92.0 150.0 0.0 659.6 197.9 857.5

Lunar Eclipse Communications 2h/d (max) 0.0 152 147.0 0.0 92.0 150.0 0.0 541.1 162.3 703.4

Spacecraft Master Equipment List Rack-up (Mass)

COMPA SS S/C 

Design 

WBS Main Subsystems

Basc (CBE) 

Mass (kg)

Growth 

(kg)

Predicted (Total) 

Mass (kg)

Aggregate 

Growth (%)

6.0 Lunar Network Satellie-High Rate 1915.4 206.3 2121.7

6.0.1 RF Communications Package 147.0 37.1 184.1 25%

6.0.2 Optical Communications Package 18.1 5.8 23.8 32%

6.0.3 Command & Data Handling (C&DH) 77.2 20.2 97.4 26%

6.0.4 Guidance, Navigation and Control System (G 78.5 15.7 94.2 20%

6.0.5 Structures & Mechanical Systems 289.3 71.3 360.6 25%

6.0.6 Power System 96.0 28.8 124.8 30%

6.0.7 Propulsion (Electric) 0.0 0.0 0.0 TBD

6.0.8 Propellant Management (EP) 0.0 0.0 0.0 TBD

6.0.9 Propellant (EP) 0.0

6.0.10 Propulsion (Chemical) 18.7 1.9 20.5 10%

6.0.11 Propellant Management (Chemical) 98.6 12.6 111.2 13%

6.0.12 Propellant (Chemical) 1035.6

6.0.13 Thermal Control (Non-Propellant) 56.5 13.0 69.5 23%

Estimated Spacecraft Dry Mass 880 206 1086 23%

Estimated Spacecraft Wet Mass 1915 206 2122

Estimated Spacecraft Inert Mass (for traj.) 910 264 1173.8

System LeveL Growth Calculations Total Growth

Dry Mass Desired System Level Growth 880 264 1143.7 30%

Mass Margin (carried a t system level) 58 7%

Total We t Mass with Growth 1915 264 2179.3

Available Launch Perform ance to C3 (kg) 4710.0 kg

Total Num ber o f LNS-HR Sa tellites 2

Total m ass o f Two LNS-HR sa tellites 4358.7

Launch m argin available (kg) 351.3 246
NASA/TM—2012-217139 42

3.5 System Level Summary

3.5.1 Frozen Orbit RF (Microwave) Concept Configuration

The desire to launch two S/C in one launch without the use of a Dual Payload Attach Fitting (DPAF)

was the primary driver in the sizing and layout of the S/C bus. By utilizing a thrust tube structure that 

extends beyond the faces of the S/C bus, two S/C can be stacked on top of one another, with only the need 

for a separation mechanism, allowing the launch and stacked loads to be carried primarily through the 

thrust tube. The stowed and stacked configuration of the Frozen Orbit RF S/C is shown back in Figure 25.

The 1.1 m diameter flange on the propellant tank drove the diameter of the thrust tube. The height 

was selected to completely enclose the propellant and pressurant tanks, allow the main thrusters to stay 

outside of the S/C yet allow for stacking in the fairing, and to maintain the proper clearances near the 

Launch Vehicle Adaptor (LVA). A conical S/C adaptor is required to mate the thrust tube to the standard 

LVA interface. The overall bus size was a resultant of the thrust tube dimensions, and the ability to

completely enclose all internal components within the bus. These subsystem components are mounted on 

structural panels that radiate from the thrust tube diagonally out the bus edges, and are grouped together 

by subsystem to allow for easier assembly. Layout of the bus structure and subsystem components is 

show in Figure 34 and Figure 35. Overall deployed dimensions are shown in Figure 36.

Figure 34.—Frozen orbit RF S/C components.
NASA/TM—2012-217139 43

Locations of the lunar antennas, Earth antenna, SA, and radiators were driven by the requirement to 

have 24-hr communication capabilities between either of the two S/C in the constellation and Earth, lunar 

outpost at the South Pole, another lunar sortie site, as well as the ability to track the Sun with the arrays.

In order to minimize gimbal requirements for tracking, the Earth antenna was placed on the side of the 

S/C bus that is always pointed back towards Earth. This “fixed” orientation resulted in minimum two-axis

gimbal angle requirements for the Earth antenna and allowed the SA to have a single axis gimbal since 

the Sun and Earth are essentially in the same plane. The array is located on a boom extended from the 

face of the S/C perpendicular to the face containing the Earth antenna, and opposite the face which 

contains the two lunar pointing antennae, to allow for the single axis Sun tracking. The radiators are 

located on the face containing the SA boom to ensure they will maintain a view that is perpendicular to 

both Sun and Earth thus increasing their effectiveness in thermal rejection. A T-shaped boom is used for 

the two lunar antennas to allow for proper tracking of the lunar surface. With the antennas located on each 

end of the T-shaped boom, the entire boom can utilize a single axis gimbal rotating 360° per orbital 

period to ensure that both antennas are pointed towards the lunar surface. Each antenna then has a twoaxis gimbal to allow each to track two separate locations on the lunar surface. The Traveling Wave Tube 

Amplifier (TWTA) for each lunar antenna is mounted to the back of the dish and contains its own radiator 

for heat rejection. Locations of all components are shown in Figure 34 and Figure 35.

Star trackers are internally mounted and have a view out of the remaining open face of the S/C. They 

are oriented in such a way that they are pointed 90° from each other and angled 45° from the face of the 

bus in the direction of the SA. This orientation ensures that the star trackers never have a direct view of 

the Sun, Earth, or lunar surface. Figure 35 illustrates the location and orientation of the star trackers.

Figure 35.—Frozen Orbit RF S/C components.
NASA/TM—2012-217139 44

Figure 36.—LNS RF frozen orbit S/C dimensions.

3.5.2 Lagrange Orbit RF Concept Configuration

Similarly to the Frozen Orbit configuration, a thrust tube was implemented in the bus design to allow 

for stacking two S/C in the payload fairing without the use of a DPAF. Again, the diameter of the thrust 

tube was driven by the propellant tank diameter, while the length was increased to include an additional 

propellant tank. All internal components were mounted in the same fashion as the Frozen Orbit 

configuration.

With the S/C in a Lagrange Orbits (one about L1 and one about L2, see Figure 22 in Section 2.6.3

timeline), the Moon and Earth are located on opposite sides of the S/C. This dictated the location and 

pointing of both lunar antennas and the Earth antenna. The lunar antennas are located on booms placed on 

opposite sides of the S/C to help balance out the forces from solar pressure considering their large 

diameter. Each has a two-axis gimbal to allow for accurate and individual pointing. The Earth antenna is 

pointed opposite of the lunar antennas and has a two-axis gimbal for accurate pointing. The SAs are 

located on booms perpendicular to the lunar antenna booms, and perpendicular to the direction of the 

Earth antenna, again to allow for single axis tracking of the Sun, which is essentially in the same plane as 

the Earth. In this case, the arrays are located on two separate booms on opposite sides of the bus. This 

location was selected to help balance out the force from the solar pressure. Dimensions of the fully 

deployed Lagrange Point S/C are shown in Figure 37.
NASA/TM—2012-217139 45

Figure 37.—LNS RF Lagrange orbit S/C dimensions.

4.0 Conclusions, Lessons Learned, and Areas for Future Study

4.1 Conclusions

Both the Frozen orbit and Lagrange orbit S/C can be dual launched on an atlas class S/C. The 

Lagrange S/C is much larger due to its large antennas and larger station keeping requirements. The 

antennas are very large solely based upon the 1 W suit radio and the distance to the Lagrange point. This 

points to making the frozen orbit constellation preferred over the Lagrange point network.

A trade of laser communications for the Earth link was made. The pointing accuracy of the S/C was 

made the same as the RF option by utilizing larger ground telescopes. Thus the differences in the S/C

design were minimal. A complete trade including the ground stations would need to be made to 

demonstrate the advantages and disadvantages of the laser communications.

The frozen orbit S/C was half the height of the Lagrange S/C due to much less propellant (much 

smaller station keeping requirements). The Frozen orbit S/C also requires a much smaller lunar antenna 

(1.2 m versus 7 m) due to much lower lunar orbit.
NASA/TM—2012-217139 46

Figure 38.—LNS RF Lagrange versus frozen S/C dimensions—Front view.

Figure 39.—LNS RF Lagrange versus frozen S/C dimensions—Side view.
NASA/TM—2012-217139 47

Figure 40.—LNS RF frozen versus Lagrange S/C predicted wet mass.

Figure 41.—LNS RF frozen versus Lagrange S/C predicted wet mass by subsystem.

4.2 Areas for Future Study

The following areas were identified during this study as future work or other items that should be 

examined in more detail.

• Trajectory analysis on actual halo Lagrange orbit in order to model the RCS propellant loading 

and pointing rather than an assumed orbit.

• Perform Risk and Reliability analysis of the Frozen orbit S/C (both RF and Optical).

• Refine the Frozen orbits and S/C design aspects
NASA/TM—2012-217139 48

• Verify the SOAP frozen orbit 12-yr perturbations with SNAP high fidelity integration code 

(orbital modeling).

5.0 Subsystem Breakdown 

5.1 Microwave Communications

5.1.1 Microwave Communications Requirements

All options must provide 100 percent 24/7 coverage to outpost, 7 day of both Sunlight and 

communications coverage at other sortie sites, and voice and global coverage as necessary. DTE 

communication for voice, available 24/7, and LNS using Ka Band will have about 28 min of 

communication time per orbit.

5.2 Microwave Communications Assumptions

Links

• LNS to Earth: Ka band (400 Mbps mono, 800 Mbps dual), S-band (150 kbps)

• LNS to outpost: S-Band (150 kbps), Ka band (400 Mbps,1200 Mbps)

• LNS to surface links: Ka 23/26 (25 Mbps) and Telemetry, Tracking and Command (TT&C)

S-band, processing for communication/navigation in payload, storage in C&DH

LNS-HR Microwave Trunk Line

• LNS-HR to Earth link (Ka-band)

○ Trunk data rate: 400 Mbps per polarization

○ Spectrum: 37/40 GHz 

○ Bandwidth: 500 MHz

○ Modulation: QPSK

○ Coding: Reed-Solomon or LDPC (TBR)

○ Code rate: 1⁄2 (TBR for each orbit)

○ Trunk antenna size: 0.5 m (TBR)

○ Bit error ratio (BER): 10×10–8

• LNS-HR to Earth link (S-band)

○ Data Rate: 150 kbps

Figure 42.—Frozen orbit communication trunk line.
NASA/TM—2012-217139 49

Earth-Based Ground Systems

• Antennas: 18 m

• Locations: DSN sites (Goldstone, Canberra, Madrid)

TT&C of LNS-HR From Earth

• Radiometrics for position/velocity knowledge using S-band

• Star tracking for orientation

General RF Aspects

• The highest modulation allowed for S-band and K/Ka-band is a form of quadrature phase-shift 

keying (QPSK).

• The forward error correction code is Reed-Solomon with Viterbi rate 1⁄2 (204, 255).

• The maximum K/Ka-band data rate is 400 Mbps uncoded.

• The maximum S-band data rate is 150 kbps uncoded.

• The digital part of the radio is software designed radio with an average power need of 37.5 W

• The high power amplifier is a TWTA for RF transmit power of 10 W or more, and solid state for 

less than 10 W of power.

• The efficiency of the TWTA is 40 percent for transmit powers less than 25 W and 50 percent for 

transmit power more than 25 W.

• The efficiency of the solid-state is 10 percent for powers greater than 4 W and 20 percent for 

power less than 5 W. (This reflects that above 4 W one generally needs to do power combining to 

achieve the power needed and below 4 W on three-stage power amplifier is all that is needed.)

• The mass and size of the TWTA does not change with power level up to several watts of power.

• The LRS be a full network node with a high-speed router, and memory for store-and-forward and 

for file serving. 

Lunar Links

• The links from the surface of the Moon to the LNS in terms of frequency and protocols are to be 

the same as from the Moon to Earth.

• The size of the antenna is sized to receive 8 kbps from the EVA suit. The maximum transmit 

power for the EVA suit is 1 W and the EVA suit uses a simple dipole antenna. The suits transmit 

linearly polarized signals and the LRS receives circular polarized signals. (Increase margin from 

3 to 6 dB.)

• The size of the antenna on the lunar communication terminal is 1 m. 

• The size of the antenna on the rover is 0.3 m.

• The two lunar coverage antennae can be backup of each other; therefore additional redundant RF 

components are not necessary

Earth Links

• The antennae looking down to Earth is a dual band capability receiving at 25.5 to 27 GHz and at 

2.20 to 2.29 GHz, and receiver at 23 to 23.5 GHz and at 2.00 to 2.09 GHz

• The system noise temperature of the antenna and the LNA system is 306 K at 25.5 GHz and 

139 K at 2.2 GHz. (This is based on a G/T of 26.5 dB/K for the 25.5 to 27 GHz band, and a G/T 

of 28.8 dB/K for the LRS 18.3 m antenna.)

5.2.1 Microwave Communications Design and MEL

The RF signals at intermediate frequency (IF) will be used to modulate individual optical 

wavelengths. The IF frequency is below 5 GHz and may differ for the K/Ka band, Ka band, and S-band 

signals. These signals will be multiplexed together and sent up a single optical fiber. A space qualified 

optical fiber rotary joint will be used in each gimbal assembly. For the antenna systems looking down at 

the lunar system with two antennas, the optical wavelength at the “T” will be split with the appropriate 

wavelength going to each antenna. At the antenna, the signals will be de-multiplexed and converted back 

to the IFs. These intermediate frequencies will be upconverted to the frequencies needed.
NASA/TM—2012-217139 50

Figure 43.—Microwave communications flow diagram.

5.2.2 Microwave Communications Analytical Methods

5.2.2.1 Lunar Communication From Frozen Orbit to Lunar Surface 

Table 25 gives the power needed to complete the links for the different forward error correction 

technique. As you can see the transmitted power is the same for the two forward error correction 

techniques, the reason being, the antenna size was adjusted to give the same bit error rate.

The antenna size to give a gain of 28.5 dBi for the receive frequency of 2.2 is 1.56 m, while for 

26.2 dBi it is 1.2 m. These antenna sizes are chosen to allow astronauts on EVA to be able to talk using 

voice over IP to the satellite.

TABLE 25.—LRS TO ELLIPTICAL ORBIT LINK BUDGET—LPDC

LRS in Elliptical Orbit d = 12400 km LPDC

Direction

Rx = 26 GHz Tx = 23 GHz for LRS Rx = 2.2 GHz Tx = 2.0 GHz for LRS

Gain, 

db Data rate,

Mbps

Tx power,

W

Gain, 

db Data rate,

Mbps

Tx power,

W Tx Rx Tx Rx

LCT→DTE 46.1 71.4 200 7.4 24.7 50 150/3000 0.22/4.4

Rover→DTE 35.6 71.4 200 9.0 14.2 50 150 2.5

EVA→LRS 1 26.2 8 1

LRS→EVA 25.4 1 150 14

LRS→ LCT 46.7 45 100 0.65 25.4 23.8 3000 0.98

LRS→Rover 46.7 34.6 10 0.72 25.4 13.4 3000 10.9

TABLE 26.—LRS TO ELLIPTICAL ORBIT LINK BUDGET—VITERBI

LRS in Elliptical Orbit d = 12400 km Viterbi

Direction

Rx = 26 GHz Tx = 23 GHz for LRS Rx = 2.2 GHz Tx = 2.0 GHz for LRS

Gain, 

db Data rate,

Mbps

Tx power,

W

Gain, 

db Data rate,

Mbps

Tx power,

W Tx Rx Tx Rx

LCT→DTE 46.1 71.4 200 7.4 24.7 50 150/3000 0.22/4.4

Rover→DTE 35.6 71.4 200 9.0 14.2 50 150 2.5

EVA→LRS 1 28.5 8 1

LRS→EVA 27.7 1 150 14

LRS→ LCT 48.9 45 100 0.65 27.7 23.8 3000 0.98

LRS→Rover 48.9 34.6 10 0.72 27.7 13.4 3000 10.9
NASA/TM—2012-217139 51

When choosing the power of the transmitter on LNS to communicate with the surface, chose the 

maximum power needed. As one can see the power needed to communicate with the rover is the 

maximum amount of power. This assumes the LCT has a 1 m antenna on it and the rover a 1/3 m antenna.

The transmit data rate from LNS to the LCT and rover is estimated but needs to be reviewed. The 

transmit frequency is 23 GHz. The data rate from LRS to LCT is 100 Mbps. This is data rate is higher 

than that for the International Space Station (ISS) and will support the activities in the habitat and in the 

habitat area. The data rate from LNS to a rover away from the habitat area is 10 Mbps. For clarity purpose 

the needed power to close those links when the temperature of the lunar surface at the equator is 425 K is 

shown for both the LCT and the rover. But in building the satellite, the large transmit power at 23 GHz is 

chosen.

Lunar Communication From Lagrange Orbit to Lunar Surface

Going out to the L2 point and allowing the astronauts to communicate at 8 kbps only changes the 

antenna size on the S/C if the losses between the antenna and the LNA and high power antenna (HPA)

remains constant. The antenna size for the Viterbi FEC case is 9.06 m in diameter and for the LPDC (lowdensity parity check) the size is 6.95 m in diameter (Table 27). All the output power from the HPAs at 

either the L1 and L2 halo orbits as well as for the elliptical frozen orbit on the lunar surface assets remain 

the same.

TABLE 27.—LRS TO LAGRANGE ORBIT LINK BUDGET—LPDC AND VITERBI

LRS in L2 Orbit d = 7202400 km LPDC

Direction

Rx = 26 GHz Tx = 23 GHz for LRS Rx = 2.2 GHz Tx = 2.0 GHz for LRS

Gain, 

db Data rate,

Mbps

Tx power,

W

Gain, 

db Data rate,

Mbps

Tx power,

W Tx Rx Tx Rx

LCT→DTE 46.1 71.4 200 7.4 24.7 50 150/3000 0.22/4.4

Rover→DTE 35.6 71.4 200 9.0 14.2 50 150 2.5

EVA→LRS 1 41.5 8 1

LRS→EVA 40.7 1 150 14.0

LRS→ LCT 61.9 45 100 0.65 40.7 23.8 3000 0.69

LRS→Rover 61.9 34.6 10 0.72 40.7 13.4 3000 10.9

LRS in L2 Orbit d = 72000 km Viterbi

Direction

Rx = 26 GHz Tx = 23 GHz for LRS Rx = 2.2 GHz Tx = 2.0 GHz for LRS

Gain, 

db Data rate,

Mbps

Tx power,

W

Gain, 

db Data rate,

Mbps

Tx power,

W Tx Rx Tx Rx

LCT→DTE 46.1 71.4 200 7.4 24.7 50 150/3000 0.22/4.4

Rover→DTE 35.6 71.4 200 9.0 14.2 50 150 2.5

EVA→LRS 1.0 43.8 8 1.0

LRS→EVA 43.0 1.0 150 14.0

LRS→ LCT 64.2 45.0 100 0.65 43.0 23.8 3000 0.69

LRS→Rover 64.2 34.6 10 0.72 43.0 13.4 3000 10.9

In conclusion on the power needed by the HPA for 23 GHz is at least 72 W while for the 2.0 GHz 

HPA the power needed is 14 W. In all cases the power can be adjusted if the assumed data rates change

(Table 28).
NASA/TM—2012-217139 52

TABLE 28.—LUNAR COMMUNICATION LINK BUDGET ASSUMPTIONS AND CALCULATIONS

Vertibi suit 

to LRS

LPDC suite 

to LRS

Tx frequency, GHz 2.2 2.2

Tx wavelength, λ = c/f, mm 136.27 136.27

Antenna efficiency, η, percent 55 55

Efficiency loss, Lη = 10log10(η/100), dB 2.60 2.60

Tx antenna diameter, DTX 0.30 0.30

Tx antenna half-power beamwidth, α(°) = 72λ/(DTX), degree 32.70 32.70

Tx antenna gain, 20log10[π(DTX)/λ] – Lη, dBi 1.00 1.00

TWTA output, W 1.00 1.00

TWTA out, dB 0.00 0.00

Output power delivery loss, dB 1.5 1.5

TWTA out, W 0.71 0.71

Tx antenna feedhorn output power, dBW –1.50 –1.50

Pointing (gimbal) loss, dB 0.1 0.1

EIRP, dBW –0.60 –0.60

Path length/range, Rb, km 12,400 12,400

Path loss, PL = 20log10 [4π(Rb)/λ], dB 181.16 181.16

Receive pointing (gimbal) loss, dB 0.1 0.1

Rain fade, 26 GHz, dB 0.00 0.00

Rain fade, 23 GHz, dB 0.00 0.00

Received power, dBW –181.86 –181.86

Rx illuminating flux density, dBW/m2 –153.46 –153.46

Rx antenna, DRX 1.56 1.20

Rx antenna gain, 20log10[π(DRX)/λ] – Lη, dBi 28.52 26.22

Rx antenna half-power beamwidth, α(°) = 72λ/(DRX) 6.29 8.20

Rx LNA gain, dB 23 23

Rx LNA noise figure, NF, (dB) wrt To=290 K, dB 0.00 0.00

Rx LNA noise temperature, TLNA = (F – 1)*To, K 0.00 0.00

Rx diplexer loss, dB 0.00 0.00

Rx waveguide/mismatch loss, dB 0.00 0.00

Rx waveguide temperature, TL, K 0.00 0.00

SLR 1 m Rx antenna (26 GHz) at Equator, K 0.00 0.00

SLR 1 m Rx antenna (2.26 GHz) at South Pole, K 425 425

Antenna galactic noise temperature, TGAL, K 0.00 0.00

Antenna resistive loss, TA,RES (0.4 emissivity), K 0.00 0.00

Combined input loss in front of LNA, dB 0.00 0.00

Input loss (lumped lossy elements), L 1.00 1.00

(L – 1)*TL, K 0.00 0.00

LNA noise contribution, L*TLNA, K 140.00 140.00

System temperature, TSYS = TA,RES + TGAL + TL + L*TLNA, K 373.75 373.75

G/T (dBi/K), dB/K 2.80 0.49

Boltzmann’s constant, k, dBW/K-Hz –228.60 –228.60

No = k*TSYS, dBW/Hz –202.88 –202.88

Power level at Rx antenna feedhord, dBW –153.34 –155.64

Rx carrier/no, dBHz 49.53 47.23

Rx bit rate, Mbps 0.008 0.008

Rx noise bandwidth, B = 2.0*Rx bit rate, MHz 0.016 0.016

Rx noise BW, dBHz 42.04 42.04

Input noise, N = k*B*T, dBW –160.83 –160.83

LRS input SNR, C/N, dB 7.49 5.19

Required Eb/No OQPSK 10–8 BER, dB 4.5 2.2

Required SNR OQPSK 10–8 BER, dB 1.49 –0.81

Link margin, dB 6.00 6.00

Required link margin, dB 6.00 6.00

Excess, dB 0.00 0.00
NASA/TM—2012-217139 53

5.2.3 Microwave Communications Risk Inputs

5.2.4 Microwave Communications Recommendation

The Lagrange orbit configuration levies large requirements for communications distances from the 

Moon to the LNS as well as from the LNS at L2 to the Earth.

5.3 Optical Communications

5.3.1 Optical Communications Requirements

All options must provide 100 percent 24/7 coverage to outpost, 7-day of both Sunlight and 

communications coverage at other sortie sites, and voice and global coverage as necessary. DTE 

communication for voice, available 24/7, and LNS using Ka Band will have about 28 min of 

communication time per orbit.

5.3.2 Optical Communications Assumptions

• Data rates

○ Downlink from LNS-HR: 1.2 Gbps (TBR)

○ Uplink to LNS-HR: 25 Mbps (TBR)

• Packet loss rates: 1×10–9 (from the NASA Jet Propulsion Laboratory (JPL))

• Size and mass

○ Mass: ~ 20 kg (from JPL) 

– Includes entire package, e.g., pointing mechanisms, tracking, etc.

– Does not include redundancy

– For single-fault tolerance—redundant diodes, redundant amplifier chain, backup 

processor

○ Physical size and dimensions: TBR 

– (6 in. diameter cylinder—4 in. pure aperture)

– Electronics box

♦ Processors, lasers, command and control, I/F to S/C bus

♦ Optical fiber and electrical connections to transceiver

• GEO Optical Relay

○ Station locations: GEO 

○ Aperture sizes: 50 cm (TBD)

– Arrayed versus single aperture

○ Transmit power: TBD

○ Photon counting receivers: TBD

– Signaling format: PPM

○ Tracking of LRS: TBD

○ Near Sun pointing: TBD

○ Satellite station keeping

– Platform stability and attitude for acquiring and pointing optical link

○ RF between Earth and optical relay

○ Data handling, archiving, and distribution due to high data rates

○ Some outages with one optical relay

5.3.3 Optical Communications Design and MEL

• Power

○ Total power allocation: 40 to 60 W (from JPL)

○ Survival power: TBD

○ Standby power: TBD
NASA/TM—2012-217139 54

○ Transmit power: ~ 1 to 2 W radiated power (TBR)

• Power efficiency:

○ Laser: 8 to 10 percent (typical) for 1550 nm (from JPL)

○ Overall system: TBD

• Optical transceiver

○ Aperture size: 5 to 10cm (from JPL)

○ Receive/transmit isolation: TBD

○ Receive/transmit wavelength: ~1550 nm with 30 nm separation between receive and transmit 

(from JPL)

○ Tracking sensor

○ Uplink detector

○ Mounted on two-axis gimbal

○ Fine pointing control

○ Solar protection filter

Figure 44.—LNS optical payload.

Figure 45.—Optical communication flow diagram.
NASA/TM—2012-217139 55

5.3.4 Optical Communications Trades

The use of the optical communications system itself was the trade performed.

5.3.5 Optical Communications Analytical Methods

Optical communications system designed from requirements and all details are experimental.

5.4 Command and Data Handling (C&DH)

The C&DH system is also often referred to as the Avionics system.

5.4.1 C&DH Requirements

• Storage array for 24 hr of storage (0.38 TB or better)

• Avionics for systems command, control, and health management

• Use of highly stable oscillators in conjunction with atomic clocks

• Functionality of the Avionics Systems is provided on a following chart

• S/C GN&C—includes interfacing with IMUs, star trackers, and Sun sensors

5.4.2 C&DH Assumptions

• Single fault tolerant avionics

• Storage array will use next generation radiation hardened storage, 128 GB or better per card

• 100 krad avionics provides operation for ~ 12 yr

• Cabling is estimated as 10 percent of the avionics hardware

• All spares are cold spares, except data recorder, which is a hot spare

5.4.3 C&DH Design and MEL

Model Summary

• One FT redundant system design

• Routine avionics cabling

• Growth factor chosen based on AIAA table

• Each C&DH Processor can access up to 32 (TBR) channels of I/O based on presently available 

space rated hardware 

Avionics Box contains a radiation tolerant PowerPC 750 processor and storage card for general LNS 

command and control. A communications card is included for communication with the LNS router. The 

package includes any necessary DC-DC converters, filter, and EMI shielding. There are two independent 

avionics strings for single fault tolerance. All avionics components based on COTS components. All 

avionics assume 3U-160 form factor cards.

An ultra-stable oscillator (USO) is included in the avionics package. There are two independent USOs 

for single fault tolerance. Oscillator drift nominally < 7×10–13/day.
NASA/TM—2012-217139 56

Figure 46.—General Avionics Processor.

A rubidium atomic clock is included in the avionics package. Clock drift assumed at nominally 5×10–14/d.

Data Recorder is based on 128 GB or better storage cards. A processor is included for managing the 

storage array. A comm. card is included for communications to the LNS router.

• General Avionics Processor (Figure 46)

○ System initialization

○ Antenna deployment

○ Antenna positioning

○ SA deployment

○ SA positioning

○ Satellite navigation—includes interfacing with IMUs, star trackers, and Sun sensors

○ Satellite guidance

○ Propulsion system control

○ Systems health and status management

○ Power management, control, distribution, and load shedding

○ Battery regulation and management

○ Thermal system management – includes control of pumps, valves, and heaters

○ System fault detection and correction

○ Time synchronization via atomic clock

○ Time stamping

○ Router management

○ Communications system management

• Data Recorder

○ Storage array monitoring and health management, including fault detection and correction

○ Synchronization of data between redundant data recorders

○ Buffering of data

○ Power down during eclipses if necessary

○ Risk assumed: technology advancements will have progressed far enough to permit 128 GB 

or better per PMC card

• Time generation unit

○ System time generation, used by the General Avionics Processor for time synchronization
NASA/TM—2012-217139 57

Both an atomic clock and a time generation unit were used in the design per recommendations from

JPL and the Optical Communications system engineers.

The components mentioned above of the C&DH system are listed in Table 29.

TABLE 29.—LNS-HR FROZEN ORBIT—C&DH MEL

WBS 

no.

Description

LNS-HR (September 2008) Frozen, RF

Quantity Unit

mass,

kg

Basic

mass,

kg

MGA,

%

Growth,

kg

Predicted

mass,

kg

6.0 Lunar Network Satellite-High Rate - --- 1288.1 11 142.8 1430.9

6.0.1 RF Communications Package - --- 123.1 12 14.9 138.0

6.0.2 Optical Communications Package - --- 0.0 0 0.0 0.0

6.0.3 C&DH - --- 77.2 26 20.2 97.4

6.0.3.1 C&DH Hardware - --- 74.2 25 18.5 92.7

6.0.3.1.a General Avionics Processor 2 4 8.0 10 0.8 8.8

6.0.3.1.b Time Generation Unit 2 0.4 0.8 20 0.2 1.0

6.0.3.1.c Command and Control Harness (data) 2 4 8.0 55 4.4 12.4

6.0.3.1.d Data Recorder 2 4 8.0 20 1.6 9.6

6.0.3.1.e Atomic Clock 2 6.6 13.2 20 2.6 15.8

6.0.3.1.f Command and Control Harness (data) 0 0 0.0 0 0.0 0.0

6.0.3.1.g Shared DPU (From APL Science Instruments) 0 0 0.0 0 0.0 0.0

6.0.3.1.h Avionics enclosure 4 8.8 35.2 25 8.8 44.0

6.0.3.1.i APL Ultra Stable Oscillator 2 0.4 0.8 10 0.1 0.9

6.0.3.1.j Router Printed Circuit Board 1 0.2 0.2 10 0.0 0.2

6.0.3.1.k Miscellaneous 3 0 0 0.0 0 0.0 0.0

6.0.3.2 Instrumentation & Wiring - --- 3.0 55 1.7 4.7

6.0.3.2.a Operational Instrumentation, sensors 8 0.3 2.4 55 1.3 3.7

6.0.3.2.b Data Cabling 2 0.3 0.6 55 0.3 0.9

6.0.3.2.c Miscellaneous 1 0 0 0.0 0 0.0 0.0

6.0.3.2.d Miscellaneous 2 0 0 0.0 0 0.0 0.0

6.0.4 GN&C - --- 42.5 20 8.5 51.0

6.0.5 Structures & Mechanical Systems - --- 233.7 23 54.7 288.4

6.0.6 Power System - --- 83.0 30 24.9 107.9

6.0.10 Propulsion (Chemical) - --- 10.6 10 1.1 11.7

6.0.11 Propellant Management (Chemical) - --- 43.9 13 5.7 49.5

6.0.12 Propellant (Chemical) - --- 617.6 0 0.0 617.6

6.0.13 Thermal Control (Non-Propellant) - --- 56.5 23 13.0 69.5

5.4.4 C&DH Trades

No trades performed.

5.4.5 C&DH Risk Inputs

The 0.3 TB storage array is designed 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.

The functions of all the avionics systems are not detailed to all lower levels. Exact packaging details 

have not been completely documented.

5.5 Guidance, Navigation and Control (GN&C)

5.5.1 GN&C Requirements

The GN&C system provides full 6-DOF control of the vehicle from launch through end of mission.
NASA/TM—2012-217139 58

Figure 47.—LNS avionics block diagram.

Figure 48.—GN&C flow diagram.

5.5.2 GN&C Assumptions

Parts assembled from off the shelf (OTS) hardware and components.

Router

Time Gen . Unit #1

Atomic

Clock Oscillator

Time Gen . Unit #2

Atomic

Clock Oscillator

Data Recorder Š0.3

TB

128 GB Memory

128 GB Memory

128 GB Memory

Processor

Comm . Card

Data Recorder Š0.3

TB

128 GB Memory

128 GB Memory

128 GB Memory

Processor

Comm . Card

Avionics Box #1

Memory Card

Dig. Cont . Card

Ana . Cont . Card

Processor

Comm . Card

Avionics Box #2

Memory Card

Dig. Cont . Card

Ana . Cont . Card

Processor

Comm . Card
NASA/TM—2012-217139 59

5.5.3 GN&C Design and MEL

• Model summary

○ Two Star Trackers

– Goodrich HD-1003

○ Two IMUs

– Honeywell MIMU

○ Sun sensors to aide in Earth acquisition

– Adcole Sun sensors, two units, three sensor heads, each

○ Four reaction wheel assemblies for attitude control and fine pointing

– 14 Nms of momentum storage for ~0.3° pointing accuracy

○ GN&C Software run on main C&DH computers

The components of the GN&C design of the Frozen orbit spacecraft is listed in Table 30.

TABLE 30.—LNS-HR FROZEN ORBIT—GN&C MEL

WBS 

no.

Description

LNS-HR (September 2008) Frozen, RF

Quantity Unit

mass,

kg

Basic

mass,

kg

MGA,

%

Growth,

kg

Predicted

mass,

kg

6.0 Lunar Network Satellite-High Rate - ------ 1288.1 11 142.8 1430.9

6.0.1 RF Communications Package - ------ 123.1 12 14.9 138.0

6.0.2 Optical Communications Package - ------ 0.0 0 0.0 0.0

6.0.3 C&DH - ------ 77.2 26 20.2 97.4

6.0.4 GN&C - ------ 42.5 20 8.5 51.0

6.0.4.1 GN&C Hardware - ------ 42.5 20 8.5 51.0

6.0.4.1.a Reaction Wheel Assembly 4 5 20.0 20 4.0 24.0

6.0.4.1.b RWA Mount 1 0.2 0.2 15 0.0 0.2

6.0.4.1.c Star Camera 2 3.402 6.8 20 1.4 8.2

6.0.4.1.d Star Camera Mount 1 0.1 0.1 15 0.0 0.1

6.0.4.1.e Inertial Measurement Units 2 4.7 9.4 20 1.9 11.3

6.0.4.1.f Course Sun Sensor Suite 6 1 6.0 20 1.2 7.2

6.0.5 Structures & Mechanical Systems - ------ 233.7 23 54.7 288.4

6.0.6 Power System - ------ 83.0 30 24.9 107.9

6.0.10 Propulsion (Chemical) - ------ 10.6 10 1.1 11.7

6.0.11 Propellant Management (Chemical) - ------ 43.9 13 5.7 49.5

6.0.12 Propellant (Chemical) - ------ 617.6 0 0.0 617.6

6.0.13 Thermal Control (Non-Propellant) - ------ 56.5 23 13.0 69.5

5.5.4 GN&C Trades

No trades performed.

5.5.5 GN&C Analytical Methods

Station-keeping Budget—Lagrange Orbit

From Table 31 the vertical cases L1, L2 are not as stable but provide better global coverage of the 

lunar surface.

For budgeting the ∆V for station keeping, use the details in the Folta, AIAA-2004-4741 technical 

report in Table 32. This gives a more general station keeping solution. Total ∆Vs of ~60 m/s/yr can be 

expected using dLQR (design linear-quadratic regulator) control method expressed in Table 32. Station

keeping burns need to be done on the order of every orbit.
NASA/TM—2012-217139 60

TABLE 31.—LAGRANGE ORBIT STATION-KEEPING 

Orbit type Liberation

point

Period,

day

Stability 

index

Average

3σ║δVi║,

cm/s

Number

of 

maneuvers

Average

time 

between,

days

Average

║∆νi║,

m/s

Average

║∆ν║,

m/s

Near-rectilinear halo L2 7.0 1.00 2.06 86 4.20 0.057 4.82

Near-rectilinear halo L1 8.0 1.25 1.52 55 6.40 0.101 5.54

Near-rectilinear halo L2 8.0 1.00 2.18 55 6.40 0.086 4.69

Halo L1 12.0 60 3.82 60 6.00 1.106 66.33

Halo L2 14.0 115 2.77 156 2.33 0.183 28.47

Vertical L1 14.0 690 3.13 68 5.25 2.527 171.82

Butterfly L2 14.0 11.3 9.78 78 4.67 0.409 31.86

Vertical L1 16.0 370 2.81 91 4.00 0.347 31.55

Vertical L2 16.0 515 2.75 60 6.00 1.472 88.32

TABLE 32.—DLQR STATIONKEEPING YEARLY COST (M/S) 

Small 

Lissajous

Small 

Halo

Large 

Lissajous

Large 

Halo

L1 no errors 6.41 6.11 5.61 5.99

L1 with errors 61.26 61.13 60.22 60.48

L2 no errors 5.37 5.38 5.38 5.61

L2 with errors 60.87 61.00 59.88 59.86

5.6 Electrical Power System

5.6.1 Power Requirements

• 12 yr operational life

• Solar arrays provide 1100 W end of life (EOL) net power including 30 percent margin

5.6.2 Power Assumptions

• Eclipse power is limited to 605 W net S/C power, including 30 percent margin, for maximum 

eclipse time of nearly 7 hr

5.6.3 Power Design and MEL

• Solar array

○ 5.9 m2 total area

○ 28.5 percent SOA triple-junction solar cells

○ Rectangular array structure

○ 15° cosine loss assumed

○ Single axis gimbal

○ 1600 W BOL net power including 30 percent margin

• Battery

○ 4.3 kWh total net energy

○ SOA Li-ion battery cells

• Power Management and Distribution

○ Switching unit

○ Battery charge control unit

○ Cabling

• Technology status

○ No special technology requirements

○ SOA SA, batteries, Power Management and Distribution (PMAD)

• SA mass: 32 kg

• Battery system mass: 56 kg

• Additional power system mass: 20 kg

The components of the power system design of the Frozen orbit spacecraft is listed in Table 33.
NASA/TM—2012-217139 61

TABLE 33.—LNS-HR FROZEN ORBIT—POWER SYSTEM MEL

WBS 

no.

Description

LNS-HR (September 2008) Frozen, RF

Quantity Unit

mass,

kg

Basic

mass,

kg

MGA,

%

Growth,

kg

Predicted

mass,

kg

6.0 Lunar Network Satellite-High Rate - --- 1288.1 11 142.8 1430.9

6.0.1 RF Communications Package - --- 123.1 12 14.9 138.0

6.0.2 Optical Communications Package - --- 0.0 0 0.0 0.0

6.0.3 C&DH - --- 77.2 26 20.2 97.4

6.0.4 GN&C - --- 42.5 20 8.5 51.0

6.0.5 Structures & Mechanical Systems - --- 233.7 23 54.7 288.4

6.0.6 Power System - --- 83.0 30 24.9 107.9

6.0.6.1 Battery System - --- 43.0 30 12.9 55.9

6.0.6.1.a Battery Assembly-Primary 1 43 43.0 30 12.9 55.9

6.0.6.1.b Battery Assembly-Secondary 0 0 0.0 0 0.0 0.0

6.0.6.2 Solar Array - --- 25.0 30 7.5 32.5

6.0.6.2.a Solar Array Panel 1 18 18.0 30 5.4 23.4

6.0.6.2.b Solar Array Structure 1 3 3.0 30 0.9 3.9

6.0.6.2.c Solar Array Drive Assembly 1 4 4.0 30 1.2 5.2

6.0.6.2.d Solar Array Interface 0 0 0.0 30 0.0 0.0

6.0.6.3 Power Management & Distribution - --- 15.0 30 4.5 19.5

6.0.6.3.a Main Bus Switching Unit 1 4 4.0 30 1.2 5.2

6.0.6.3.b Battery Charge Control Unit 1 1 1.0 30 0.3 1.3

6.0.6.3.c Power Cabling 1 10 10.0 30 3.0 13.0

6.0.6.3.d Miscellaneous 1 0 0 0.0 0 0.0 0.0

6.0.6.3.e Miscellaneous 2 0 0 0.0 0 0.0 0.0

6.0.10 Propulsion (Chemical) - --- 10.6 10 1.1 11.7

6.0.11 Propellant Management (Chemical) - --- 43.9 13 5.7 49.5

6.0.12 Propellant (Chemical) - --- 617.6 0 0.0 617.6

6.0.13 Thermal Control (Non-Propellant) - --- 56.5 23 13.0 69.5

5.6.4 Power Trades

More S/C eclipse power versus more battery mass.

5.6.5 Power Analytical Methods

Figure 49 shows the variation of eclipse time for the frozen orbit spacecraft throughout the 8 year 

mission lifetime. The eclipse times are used to size the battery systems.

Figure 49.—Eclipse time for frozen orbit power system modeling.
NASA/TM—2012-217139 62

5.6.6 Power Recommendation

• Further safe-mode-like operations for longest (but rare) eclipses will reduce battery mass further

• Ultraflex SA may reduce mass at potentially higher cost

5.7 Structures and Mechanisms

5.7.1 Structures and Mechanisms Requirements

• Contain necessary hardware for research instrumentation, avionics, communications, propulsion 

and power

• Withstand applied loads from launch vehicle and provide minimum deflections, sufficient 

stiffness, and vibration damping

• Maximum longitudinal loads: 6g 

• Maximum lateral loads: 4g

• Ability to have two probes stacked in launch vehicle

• Minimize weight

• Fit within confines of launch vehicle

• Accommodate landing and takeoff on terrestrial body

5.7.2 Structures and Mechanisms Assumptions

• Material: Aluminum

• Central thrust tube

• Space frame with square tubular members

• Shear panels

• Welded and threaded fastener assembly

5.7.3 Structures and Mechanisms Design and MEL

The thrust tube carries another S/C above as well as propellant tanks. The S/C volume was sufficient 

for all systems. Each system populates a single side of a stiffener panel. 

Frozen Orbit Main Structures

• Installation mass was calculated as 4 percent of the CBE mass of mounted unit masses

• Negligible stress, 140 psi, in thrust tube with 4g lateral load and stacked

Figure 50 shows the preliminary wire model of the LNS Frozen orbit spacecraft primary structural 

elements. This preliminary model was used to determine the size of the primary structures of the 

spacecraft bus. The structural components and masses are listed in Table 34. Figure 51 shows the 

preliminary wire model of the LNS Lagrange orbit spacecraft primary structural elements. This 

preliminary model was used to determine the size of the primary structures of the spacecraft bus. 

Figure 50.—LNS frozen orbit structure wire model. Figure 51.—LNS Lagrange orbit structure wire model.
NASA/TM—2012-217139 63

TABLE 34.—LNS-HR FROZEN ORBIT—STRUCTURES MEL

WBS 

no.

Description

LNS-HR (September 2008) Frozen, RF

Quantity Unit

mass,

kg

Basic

mass,

kg

MGA,

%

Growth,

kg

Predicted

mass,

kg

6.0 Lunar Network Satellite-High Rate ---- ------- 1288.1 11 142.8 1430.9

6.0.1 RF Communications Package ---- ------- 123.1 12 14.9 138.0

6.0.2 Optical Communications Package ---- ------- 0.0 0 0.0 0.0

6.0.3 C&DH ---- ------- 77.2 26 20.2 97.4

6.0.4 GN&C ---- ------- 42.5 20 8.5 51.0

6.0.5 Structures & Mechanical Systems ---- ------- 233.7 23 54.7 288.4

6.0.5.1 Primary Structures ---- ------- 199.1 25 49.8 248.9

6.0.5.1.a Separation Ring 1 18.72 18.7 25 4.7 23.4

6.0.5.1.b Top/Bottom Deck 2 15.54 31.1 25 7.8 38.9

6.0.5.1.c Vertical Inner Panels 4 3.942 15.8 25 3.9 19.7

6.0.5.1.d Side Panels 4 14.09 56.4 25 14.1 70.5

6.0.5.1.e Corner post 4 0.668 2.7 25 0.7 3.3

6.0.5.1.f Top Cover 0 0 0.0 25 0.0 0.0

6.0.5.1.g Thrust tube 1 74.51 74.5 25 18.6 93.1

6.0.5.1.h Miscellaneous clips/fasteners 0 0 0.0 25 0.0 0.0

6.0.5.1.i Miscellaneous 1 0 0 0.0 15 0.0 0.0

6.0.5.1.j Miscellaneous 2 0 0 0.0 15 0.0 0.0

6.0.5.2 Secondary Structures ---- ------- 7.3 25 1.8 9.1

6.0.5.2.a Antenna Boom 1 5.126 5.1 25 1.3 6.4

6.0.5.2.b Solar Array Boom (s ) 1 2.148 2.1 25 0.5 2.7

6.0.5.2.c EP Thruster Boom s 0 0.537 0.0 25 0.0 0.0

6.0.5.2.d Chemical Thruster Boom s 0 0.537 0.0 25 0.0 0.0

6.0.5.2.e Miscellaneous 1 0 0 0.0 0 0.0 0.0

6.0.5.2.f Miscellaneous 2 0 0 0.0 0 0.0 0.0

6.0.5.3 Installation ---- ------- 12.3 25 3.1 15.4

6.0.5.3.a RF Communications Installation 1 4.804 4.8 25 1.2 6.0

6.0.5.3.b Optical Communications Installation 0 0 0.0 25 0.0 0.0

6.0.5.3.c C&DH Installation 1 0.916 0.9 25 0.2 1.1

6.0.5.3.d GN&C Installation on 1 1.7 1.7 25 0.4 2.1

6.0.5.3.e Power Installation 1 2.92 2.9 25 0.7 3.7

6.0.5.3.f Propulsion (EP) Installation 0 0 0.0 25 0.0 0.0

6.0.5.3.g Propellant (EP) Storage Installation 0 0 0.0 25 0.0 0.0

6.0.5.3.h Propulsion (Chemical) Installation 1 0.48 0.5 25 0.1 0.6

6.0.5.3.i Propellant (Chemical) Storage Installation 0 1.663 0.0 25 0.0 0.0

6.0.5.3.j Thermal Installation 1 1.494 1.5 25 0.4 1.9

6.0.5.3.k Miscellaneous 1 0 0 0.0 0 0.0 0.0

6.0.5.3.l Miscellaneous 2 0 0 0.0 0 0.0 0.0

6.0.5.3.m Miscellaneous 3 0 0 0.0 0 0.0 0.0

6.0.5.4 Mechanisms ---- ------- 15.0 0 0.0 15.0

6.0.5.4.a Solar array deployment mechanism 1 5 5.0 0 0.0 5.0

6.0.5.4.b radiator deployment mechanism (if applicable) 0 0 0.0 0 0.0 0.0

6.0.5.4.c Separation mechanism (pyros ) 0 0 0.0 0 0.0 0.0

6.0.5.4.d Bi-axial antenna gimbal 1 6 6.0 0 0.0 6.0

6.0.5.4.e Antenna deployment 1 4 4.0 0 0.0 4.0

6.0.5.4.f Miscellaneous 3 0 0 0.0 0 0.0 0.0

6.0.6 Power System ---- ------- 83.0 30 24.9 107.9

6.0.10 Propulsion (Chemical) ---- ------- 10.6 10 1.1 11.7

6.0.11 Propellant Management (Chemical) ---- ------- 43.9 13 5.7 49.5

6.0.12 Propellant (Chemical) ---- ------- 617.6 0 0.0 617.6

6.0.13 Thermal Control (Non-Propellant) ---- ------- 56.5 23 13.0 69.5
NASA/TM—2012-217139 64

Lagrange Point Orbit Main Structures

• Structure width and thrust tube diameter maintained relative to frozen orbit structure

• Length increased relative to frozen orbit structure

• Installation mass, 4 percent of mounted unit mass

• Negligible stress, 182 psi, in thrust tube with 4g lateral load and stacked

5.7.4 Structures and Mechanisms Analytical Methods

Main structure is designed using launch loads and rough stresses and analysis at a top level. 

Preliminary structural analysis with given launch loads. Scaling equations are used for some mass 

numbers. Modeling is done in close coordination with configuration in order to get accurate dimensions 

for the main structure for mass calculations.

5.7.5 Structures and Mechanisms Risk Inputs

• Potential impact with foreign object or due to nearby operations.

5.7.6 Structures and Mechanisms Recommendation

• Current configuration based on heritage design

• Accommodate longitudinal and lateral launch loads

• Thrust tube bears majority structural loads

5.8 Propulsion and Propellant Management

5.8.1 Propulsion and Propellant Management Requirements

• Subsystem is single-fault tolerant, excluding DMR (Design for Minimum Risk) elements such as 

the propellant tank and main propellant lines.

5.8.2 Propulsion and Propellant Management Assumptions

• Thruster Operation

○ Thruster operation constant over duration of burn

• Propellant Storage

○ Propellant delivery assumed to be isothermal

○ Same with He pressurant into main tank

• Design for Minimum Risk Applied to Tanks and Feed Lines

○ Following Exploration lead

○ Limit DMR to Passive elements

5.8.3 Propulsion and Propellant Management Design and MEL

The monoprop propulsion subsystem used for station-keeping is comprised of (Figure 52):

• Two x 100-lbf-thrust MR-104 Hydrazine Monopropellant Engines—one operating, one spare

• Four pods of three each 0.5-lbf-thrust MR-111E Hydrazine Monopropellant Engines—all 

operating

• One metallic Ti hydrazine storage tank
NASA/TM—2012-217139 65

• One metallic Ti high pressure He storage tank 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

The components of the Propulsion system design and Propellant of the Frozen orbit spacecraft is 

listed in Table 35.

Figure 52.—Mono-Prop Propulsion System schematic.
NASA/TM—2012-217139 66

TABLE 35.—LNS-HR FROZEN ORBIT—PROPULSION AND PROPELLANT MEL

WBS 

no.

Description

LNS-HR (September 2008) Frozen, RF

Quantity Unit 

mass,

kg

Basic

mass,

kg

MGA,

%

Growth,

kg

Predicted

mass,

kg

6.0 Lunar Network Satellite-High Rate - ------- 1288.1 11 142.8 1430.9

6.0.1 RF Communications Package - ------- 123.1 12 14.9 138.0

6.0.2 Optical Communications Package - ------- 0.0 0 0.0 0.0

6.0.3 C&DH - ------- 77.2 26 20.2 97.4

6.0.4 GN&C - ------- 42.5 20 8.5 51.0

6.0.5 Structures & Mechanical Systems - ------- 233.7 23 54.7 288.4

6.0.6 Power System - ------- 83.0 30 24.9 107.9

6.0.10 Propulsion (Chemical) - ------- 10.6 10 1.1 11.7

6.0.10.1 Main Engine - ------- 4.6 10 0.5 5.1

6.0.10.1.a Main Engine 2 2.313 4.6 10 0.5 5.1

6.0.10.1.b Main Engine Gimbal 0 0 0.0 10 0.0 0.0

6.0.10.2 Reaction Control System - ------- 6.0 10 0.6 6.6

6.0.10.2.a RCS Engine 4 1.495 6.0 10 0.6 6.6

6.0.11 Propellant Management (Chemical) - ------- 43.9 13 5.7 49.5

6.0.11.1 OMS Propellant Management - ------- 0.0 0 0.0 0.0

6.0.11.1.a Fuel Tanks 0 0 0.0 10 0.0 0.0

6.0.11.1.b Fuel Lines 0 0 0.0 25 0.0 0.0

6.0.11.1.c Oxidizer Tanks 0 0 0.0 0 0.0 0.0

6.0.11.1.d Oxidizer Lines 0 0 0.0 0 0.0 0.0

6.0.11.1.e Pressurization System - tanks , panels , lines 0 0 0.0 25 0.0 0.0

6.0.11.1.f Feed System - regulators, valves, etc. 0 0 0.0 25 0.0 0.0

6.0.11.2 RCS Propellant Management - ------- 43.9 13 5.7 49.5

6.0.11.2.a Fuel Tanks 1 26.33 26.3 10 2.6 29.0

6.0.11.2.b Fuel Lines 0 0 0.0 0 0.0 0.0

6.0.11.2.c Pressurization System - tanks, panels, lines 1 13.63 13.6 15 2.0 15.7

6.0.11.2.d Feed System - regulators, valves, etc. 1 3.92 3.9 25 1.0 4.9

6.0.12 Propellant (Chemical) - ------- 617.6 0 0.0 617.6

6.0.12.1 Main Engine Propellant - ------- 0.0 0 0.0 0.0

6.0.12.1.a Fuel - ------- 0.0 0 0.0 0.0

6.0.12.1.a.a Fuel Usable 0 0 0.0 0 0.0 0.0

6.0.12.1.a.b Fuel Boiloff 0 0 0.0 0 0.0 0.0

6.0.12.1.a.c Fuel Residuals (Unused) 0 0 0.0 0 0.0 0.0

6.0.12.1.b Oxidizer - ------- 0.0 0 0.0 0.0

6.0.12.1.b.a Oxidizer Usable 0 0 0.0 0 0.0 0.0

6.0.12.1.b.b Oxidizer Boiloff 0 0 0.0 0 0.0 0.0

6.0.12.1.b.c Oxidizer Residuals (Unused) 0 0 0.0 0 0.0 0.0

6.0.12.1.c Main Engine Pressurant 0 0 0.0 0 0.0 0.0

6.0.12.2 RCS Propellant - ------- 617.6 0 0.0 617.6

6.0.12.2.a Fuel - ------- 614.5 0 0.0 614.5

6.0.12.2.a.a Fuel Usable 1 599.5 599.5 0 0.0 599.5

6.0.12.2.a.b Fuel Boiloff 0 0 0.0 0 0.0 0.0

6.0.12.2.a.c Fuel Residuals (Unused) 1 14.99 15.0 0 0.0 15.0

6.0.12.2.b Oxidizer - ------- 0.0 0 0.0 0.0

6.0.12.2.b.a Oxidizer Usable 0 0 0.0 0 0.0 0.0

6.0.12.2.b.b Oxidizer Boiloff 0 0 0.0 0 0.0 0.0

6.0.12.2.b.c Oxidizer Residuals (Unused) 0 0 0.0 0 0.0 0.0

6.0.12.2.c RCS Pressurant 1 3.123 3.1 0 0.0 3.1

6.0.13 Thermal Control (Non-Propellant) - ------- 56.5 23 13.0 69.5
NASA/TM—2012-217139 67

5.8.4 Propulsion and Propellant Management Analytical Methods

Hydrazine Thrusters

• Aerojet MR-104 thruster characteristics (Figure 53)

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

○ Specific impulse: 239; 239 to 223 sec

○ Mass: 1.86 kg

○ Power = 43.1 W

○ Lifetime = 2,654 sec (cumulative)

• Aerojet MR-111E thruster characteristics (Figure 54)

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

○ Specific Impulse: 224; 224 to 213 sec

○ Mass: 0.33 kg

○ Power = 13.64 W

○ Lifetime = 26.7 hr (cumulative)

• Storage tank based on COTS unit from ATK-PSC, Inc. (Figure 55)

○ Titanium metallic tank

○ Derived from Model No. 80352-1

– Size: 1.06 m diam. spherical (42 in. diam.)

– Internal volume = 0.59 m3 (20.3 ft3

)

○ Minor size changes to match propellant load

○ Propellant management device included in tank mass

Figure 53.—Aerojet MR-104 Thruster. Figure 54.—Aerojet MR-111E Thruster.

Figure 55.—Propellant Storage Tank.
NASA/TM—2012-217139 68

• Helium Pressurization System

○ One metallic Ti spherical tank, storing pressurant at 2800 psia

• Propellant Distribution System Components

○ Derived from existing technology

○ Did not need to include technology in development

○ Minimal redundancy to reduce system mass

• All Components of Propulsion Subsystem at High TRL

○ Little or no development costs required

5.8.5 Propulsion and Propellant Management Risk Inputs

• Potential risks

○ RCS thruster plume impingement (Low)

– Potential degradation due to deposition on sensitive surfaces

○ Propellant Freezing within fluid lines (Low-Medium)

– Freezing could lead to line ruptures

5.8.6 Propulsion and Propellant Management Recommendation

• Propellant storage and delivery

○ Single metallic tank selected for hydrazine storage

○ He pressurization selected over blowdown approach to save mass

– Preliminary trade on blowdown system found the tanks required for storage of the 

hydrazine propellant to be significantly larger and therefore heavier

○ He pressurization and propellant distribution systems are configured to be single-fault 

tolerant

– Excluding tanks and lines which are single string

5.9 Thermal Control

5.9.1 Objective

To provide spreadsheet based models capable of estimating the mass and power requirements of the 

various thermal systems. The thermal modeling provides power and mass estimates for the various 

aspects of the vehicle thermal control system based on a number of inputs related to the vehicle geometry, 

flight environment and component size. The system consists of the following elements 

• Electric heaters 

• MLI

• Thermal paint

• Radiator with louvers

• Thermal Control System (sensors, switches, data acquisition)

5.9.2 Thermal Requirements

The thermal requirements for the LNS 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. Figure 56 shows the Frozen orbit configuration with the thermal components identified.

The maximum heat load to be rejected by the thermal system was 656 W, and the desired operating 

temperature for the electronics and propellant was 300 K. 
NASA/TM—2012-217139 69

Figure 56.—LNS Frozen Orbit Thermal Systems.

5.9.3 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 SA were assumed to be 0.1, 0.25 

and 0.1 respectively. 

• The maximum angle of the radiator to the Sun was 30°.

• The radiator temperature was 320 K.

5.9.4 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 aluminum cold plates located throughout the interior of the 

S/C. These cold plates have heat pipes integrated into them. The heat pipes transfer heat from the cold 

plates to the radiator, which radiates the excess heat to space. The portions of the heat pipes that extend 

from the S/C 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. 

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

temperatures.

The components of the Thermal system design of the Frozen orbit spacecraft is listed in Table 36.
NASA/TM—2012-217139 70

TABLE 36.—LNS-HR FROZEN ORBIT—THERMAL MEL

WBS 

no.

Description

LNS-HR (September 2008) Frozen, RF

Quantity Unit 

mass,

kg

Basic

mass,

kg

MGA,

%

Growth,

kg

Predicted

mass,

kg

6.0 Lunar Network Satellite-High Rate -- -------- 1288.1 11 142.8 1430.9

6.0.1 RF Communications Package -- -------- 123.1 12 14.9 138.0

6.0.2 Optical Communications Package -- -------- 0.0 0 0.0 0.0

6.0.3 C&DH -- -------- 77.2 26 20.2 97.4

6.0.4 GN&C -- -------- 42.5 20 8.5 51.0

6.0.5 Structures & Mechanical Systems -- -------- 233.7 23 54.7 288.4

6.0.6 Power System -- -------- 83.0 30 24.9 107.9

6.0.10 Propulsion (Chemical) -- -------- 10.6 10 1.1 11.7

6.0.11 Propellant Management (Chemical) -- -------- 43.9 13 5.7 49.5

6.0.12 Propellant (Chemical) -- -------- 617.6 0 0.0 617.6

6.0.13 Thermal Control (Non-Propellant) -- -------- 56.5 23 13.0 69.5

6.0.13.1 Active Thermal Control -- -------- 4.2 21 0.9 5.1

6.0.13.1.a Heaters 10 0.143 1.4 15 0.2 1.6

6.0.13.1.b Thermal Control/Heaters Circuit 2 0.2 0.4 25 0.1 0.5

6.0.13.1.c Data Acquisition 1 1 1.0 25 0.3 1.3

6.0.13.1.d Thermocouples 25 0.01 0.3 15 0.0 0.3

6.0.13.1.e Radiator MMOD Shielding 1 1.109 1.1 25 0.3 1.4

6.0.13.2 Passive Thermal Control -- -------- 41.4 23 9.5 50.9

6.0.13.2.a Heat Sinks 4 3.463 13.9 20 2.8 16.6

6.0.13.2.b Heat Pipes 1 1.98 2.0 25 0.5 2.5

6.0.13.2.c Radiators 1 17.49 17.5 25 4.4 21.9

6.0.13.2.d MLI 1 5.497 5.5 25 1.4 6.9

6.0.13.2.e Temperature sensors 50 0.01 0.5 15 0.1 0.6

6.0.13.2.f Phase Change Devices 0 0 0.0 0 0.0 0.0

6.0.13.2.g Thermal Coatings /Paint 1 0.907 0.9 15 0.1 1.0

6.0.13.2.h Antenna Radiator 2 0.586 1.2 25 0.3 1.5

6.0.13.3 Semi-Passive Thermal Control -- -------- 10.9 24 2.6 13.5

6.0.13.3.a Louvers 1 10.07 10.1 25 2.5 12.6

6.0.13.3.b Thermal Switches 4 0.2 0.8 15 0.1 0.9

5.9.5 Thermal Trades

None

5.9.6 Thermal Analytical Methods

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

S/C to the surroundings. This analysis takes into account the design and layout of the thermal system and 

the thermal environment to which heat is being rejected to or insulated from. For more detailed 

information on the thermal analysis a summary white paper titled “Preliminary Thermal System Sizing” 

was produced. 

5.9.6.2 Environmental Models

Solar intensity modeling was based on S/C location. Components were sized for worst case operating 

conditions, heat rejection: near Earth, minimum temperature: lunar orbital location.

5.9.6.3 Systems Modeled

• Micrometeor shielding on radiator

• Radiator panels (placement, sizing)

• Thermal control of propellant lines and tanks

• S/C insulation (layers of MLI, MLI placement)
NASA/TM—2012-217139 71

• Avionics and PMAD cooling (number of cold plates, heat pipe length)

• Component heating (electric heaters and/or radioisotope heater units (RHU))

Table 37 captures the inputs and outputs used in the design of the thermal system.

TABLE 37.—THERMAL SYSTEM INPUTS AND OUTPUTS DATA PASSING

Thermal system input Thermal modeling output

S/C dimensions (length, diameter) Heat pipe length and mass

Power management and electronics dimensions Cold plate size and mass

Waste heat load to be rejected

View factor to the Earth, lunar surface and SAs

Solar flux

Radiator size and mass

S/C insulation mass and thickness

Thermal system components mass

Propellant tank dimensions and operating temperature Propellant tanks insulation mass and heater power level

Propellant line lengths and operating temperature Propellant line insulation mass and heater power level

Component minimum temperature requirements Heating mass and power requirement

5.9.6.4 Radiator Sizing

The radiator panel area has been modeled along with an estimate of its mass. The model was based on 

first principles analysis of the area needed to reject the identified heat load to space. From the area, a 

series of scaling equations were used to determine the mass of the radiator within the lunar environment. 

Lunar orbit 1 AU thermal environment was used to size the radiator.

Louvers are active or passive devices that regulate the amount of heat rejected by the radiator. Active 

controlled louvers use temperature sensors and actuators to control the louver position. Passive controlled 

louvers commonly use a bimetallic spring that opens and closes the louver based on temperature. The 

louver specific mass is 4.5 kg/m2

. Table 38 captures the assumptions used in the design of the radiator 

thermal system. Figure 57 shows an artist’s representation of a louver prototype.

TABLE 38.—THERMAL SYSTEM

RADIATOR SIZING ASSUMPTIONS

Variable Value

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

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

Radiator Sun angle ....................................................................... 70°

Radiator operating temperature..................................................320 k

Total radiator dissipation power ........................................... 656.5 W

View Factor to Solar Array..........................................................0.10

View Factor to Earth....................................................................0.10

View Factor to Moon...................................................................0.25

Figure 57.—Schematic view of the louver prototype.
NASA/TM—2012-217139 72

Figure 58.—Heat loss (heater power) and MLI mass as a function of MLI layers for a 1 m diameter 

propellant tank.

5.9.6.5 Thermal Analysis Propellant Lines and Tanks

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

heaters and propellant line insulation and heaters (Figure 58). 

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

propellant lines through the S/C structure to space. The heat loss through the insulation set the power 

requirement for the tank and line heaters. The 1 AU thermal environment was used to calculate the heat 

loss. Table 39 shows the assumptions used.

TABLE 39.—THERMAL SYSTEM TANK INSULATION SIZING ASSUMPTIONS

Variable Value

Tank surface emissivity (εt).......................................................................................................0.1

MLI emissivity (εi) ..................................................................................................................0.07

MLI material...............................................................................................................................Al

MLI material density (ρi)............................................................................................ 2,770 kg/m3

Internal tank temperature (Ti) ...............................................................................................300 K

MLI layer thickness (ti)...................................................................................................0.025 mm

Number of insulation layers (ni) .................................................................................................10

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

Tank immersion heater mass 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

5.9.6.6 Thermal Analysis—S/C Insulation

The mass of the S/C MLI insulation was modeled to determine the mass of the insulation and heat 

loss. The model was based on a first principles analysis of the heat transfer from the S/C through the 

insulation to space. Nighttime lunar surface thermal environment was used to size the insulation. Two 

types of heaters were considered, RHU, and electrical heaters. Table 40 shows the assumptions used.
NASA/TM—2012-217139 73

TABLE 40.—THERMAL SYSTEM TANK

INSULATION SIZING ASSUMPTIONS

Variable Value

S/C MLI material............................................................................................ Al

S/C MLI material density (ρisc)........................................................ 2,770 kg/m3

MLI layer thickness (ti)....................................................................... 0.025 mm

Number of insulation layers (ni).................................................................... 100

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

S/C Radius (rsc) .................................................................................... 0.825 m

5.9.6.7 Thermal Analysis—PMAD Cooling

Thermal control of the electronics and Active Thermal Control System (ATCS) is accomplished 

through a series of cold plates and heat pipes to transfer the excess heat to the radiators. The model for 

sizing these components was based on a first principles analysis of the area needed to reject the identified 

heat load to space. From the sizing a series of scaling equations were used to determine the mass of the 

various system components. Table 41 shows the assumptions used.

TABLE 41.—THERMAL SYSTEM PMAD 

COOLING SIZING ASSUMPTIONS

Variable Value

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

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

Number of cooling plates......................................................................................4

Cooling plate lengths.....................................................................................0.5 m

Cooling plate widths......................................................................................0.5 m

Cooling plate thickness................................................................................. 5 mm

Heat pipe specific mass...........................................................................0.15 kg/m

5.9.7 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, 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.—A failure of the louvers used to regulate the heat dissipation of the 

radiators can cause the radiator to partially or fully fail. The louvers are however, built of separate 

slats that are individually controlled by a passive temperature sensitive spring mechanism. Failure 

of one or more of these spring mechanisms would degrade the performance of the radiator. The 

margin built into the radiator sizing can accommodate some louver failure without impacting the 

required heat dissipation capability of the radiator. 
NASA/TM—2012-217139 74

5.9.1 Thermal Recommendation

To improve the reliability of the system and compensated 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 the known shadow period of operation. 

• Building margin into the radiator sizing will provide a means of accommodating any failure that 

may occur with the louver system. 

6.0 Software Cost Estimation

6.1 Objectives

• To understand the functional requirements that the LRS is designed to meet (software is strongly 

affected by hardware design) 

• To develop software cost estimates based on the LRS functional requirements, especially the LRS 

avionics functionality.

6.2 Assumptions

• Real-time operating system (RTOS) (such as VxWorks or RT Linux) for a single-board computer 

or a multipurpose computer to be ported to the LRS system/subsystems

• Inter-network Operating System (Similar to Cisco IOS) to be ported to LRS router

• IP-based networks to be supported

• Comply with all security protocols defined by Constellation C3I

• Any new software to be developed will be C/C++ or Ada or a combination of these languages for 

mission critical such as avionics. For this software cost estimate, the study team assumed C.

• Source lines of code (SLOC) is a count of the text of the source code including comment lines, 

neither “physical” nor “logical” is considered as the types of SLOC measures. Software sizing 

and estimation are based on mainly on the LRS hardware functionality for avionics and 

communication and navigation. Other sources and references are also used for estimating number 

of SLOC. For this study, see Table 42 for the SLOC estimate is used based on inputs from the 

COMPASS team

• The variation of the provided SLOC from –10 to 30 percent is consistent with other software 

estimates generated for COMPASS.
NASA/TM—2012-217139 75

TABLE 42.—ASSUMED SLOC FOR LNS

LRS Avionics Functionality SLOC –10% Provided 30% Application

General Avionics System initialization 400 360 400 520 System Device Utilities

Processor

Antenna deployment and positioning 800 720 800 1,040 Flight Systems

Solar array deployment and positioning 800 720 800 1,040 Flight Systems

Satellite navigation 11,400 10,260 11,400 14,820 Flight Systems

Satellite guidance 2,000 1,800 2,000 2,600 Flight Systems

Propulsion system control 300 270 300 390 Flight Systems

Systems health and status management 3,000 2,700 3,000 3,900 Network Management

Power management and monitoring 500 450 500 650 Diagnostics

Battery regulation and management 500 450 500 650 Diagnostics

Thermal system management 300 270 300 390 System Device Utilities

System fault detection and correction 2,500 2,250 2,500 3,250 Diagnostics

Time synchronization and stamping 200 180 200 260 Process Control

Router management (assuming no new 

software has to be written) 0 ------ ------ ------ OS/Executive

Communication system management 800 720 800 1,040 Communications

Data Recorder

Storage array monitoring and health 

management 300 270 300 390 Network Management

Synchronization of data between redundant 

data recorders 200 180 200 260 Testing Software

Buffering data from WAN and LAN for 

transmission to Earth 300 270 300 390 Communications

Estimated total SLOC—Avionics 24,300 ------ ------ ------

Communications

Command processing 400 360 400 520 Command/Control

Telemetry processing 400 360 400 520 Communications

Navigation Attitude Determination and Control 11,000 9,900 11,000 14,300 Flight Systems

Estimated total SLOC—Comm and nav 11,800 ------ ------ ------

Total SLOC 36,100 32,490 36,100 46,930

6.3 Approach

The System Evaluations and Estimation of Resources-Software Estimation Model (SEER-SEM), 

Version 7.7.21 was used to estimate the cost of software development. The SLOC estimate from the 

previous section was used as the major input for this model. SLOC includes all executable LOC, nonexecutable declarations and compiler directives, but excludes comments, banners, blank lines, and nonblank spacers. Additional SEER-SEM specific assumptions for this study are as follows: 

• Platform: Unmanned space

• Application: Shown in Table 42

• Development method: No knowledge

• Development standard: No knowledge

• Class: No knowledge

Using the inputs and assumptions listed above, the detailed breakdown of the software cost estimate is 

shown in Table 43. Monte Carlo risk analysis is performed on the estimate based on the SLOC range for 

each subprogram modeled in SEER-SEM. The risk analysis results for this estimate are as Table 44.
NASA/TM—2012-217139 76

TABLE 43.—SEER-SEM RESULTS: SOFTWARE COST BREAKDOWN FOR LNS
NASA/TM—2012-217139 77

TABLE 44.—SEER-SEM RESULTS: SOFTWARE COST RISK ANALYSIS FOR LNS

7.0 Cost, Risk and Reliability

7.1 Costing

The S/C cost and life cycle cost of a mission are important products developed by the COMPASS 

team. Just as each subsystem lead performs a risk assessment on their system, cost risk must also be 

addressed. In order to evaluate uncertainties in a mission that affect cost, the COMPASS team uses the 

following steps to develop the cost estimate and address cost risk for that mission:

1. Take the final MEL and Map each element of the MEL to a cost spreadsheet.
NASA/TM—2012-217139 78

2. Apply an appropriate Cost Estimating Relationship (CER) to each element and/or subsystem of 

the MEL. Currently, the cost estimation office has a repository of mass-based, parametric CERs 

that were developed using historical costs. Additional cost methodologies may be developed and 

applied to address and unique elements or subsystems for a given mission.

3. Discuss with the COMPASS team if any of the subsystems have inheritance from previous 

missions, determine the flight sparing philosophy, determine the development approach (prototype, proto-flight, etc.), and decide if there are any other factors that may affect cost.

4. Once each element or subsystem is appropriately modeled with a CER, quantitative risk analysis 

is performed on the S/C cost using Monte Carlo simulation based on mass and CER uncertainties.

5. Life Cycle Costs for the mission may also be developed by adding launch services, mission 

operations, project management, reserves and other relevant costs for that mission.

In order to provide a cost estimate for each of the four various trades in this report, the following 

assumptions apply:

• A proto-type development approach is assumed for each option, except for some payload 

development items, which include both an engineering model and a qualification article. For a 

proto-flight development, 50 percent of the hardware cost for each subsystem/component is 

added to the development cost for hardware refurbishment.

• The cost for a single flight spare is included where appropriate.

• The life cycle cost estimates are for two satellites and assume that both satellites are launched on 

a single launch vehicle.

• Mission operations costs and the cost for any ground system necessary to support these satellites 

are not included in the estimates provided in this section.

Table 45 is a life cycle cost comparison for each of the four cases. The cost for two satellites is 

included for each of the cases. All costs are in FY08$M.

As seen in Table 45, NASA Project Office/Technical Oversight includes 5 percent of all other costs.

Phase A costs are calculated at 5 percent of the S/C cost and fee. The satellite cost is for two satellites and

is the mean estimate base on the stochastic simulation results. For the second satellite, a learning curve of 

95 percent is assumed. S/C Prime Contractor Fee is 10 percent of the S/C cost. The launch services costs 

are provided by KSC and assume that both satellites are launched on a single platform. Reserves are 

calculate at 25 percent and are not applied to launch services.

TABLE 45.—LIFE CYCLE COST COMPARISON OF THE FOUR OPTIONS

Lunar Network Satellite

FY08$M Frozen RF Frozen Opt Lagrange RF Lagrange Opt

NASA Project Office/Technical Oversight 36 36 42 43

Phase A 27 27 33 33

Two Satellites 482 495 597 605 

Spacecraft Prime Contractor Fee (10%) 48 49 60 61

Launch Services 158 158 158 158 

Reserves (25%) 148 152 183 185 

Life Cycle Cost 899 918 1,073 1,085 

*Life cycle cost if for two satellites

Table 46 to Table 49 show breakdowns of costs at the subsystem level for each of the various trades.

In each case, the costs provided are in FY08$M and only include the cost of a single satellite.
NASA/TM—2012-217139 79

TABLE 46.—SUBSYSTEM COST BREAKDOWN FOR A

SINGLE SATELLITE FOR THE FROZEN RF OPTION IN FY08$M

WBS Description DDT&E total by 

$M

Flight hardware by 

$M

Mfg/DDT&E total 

by $M

6.0.1 RF Communications Package $53.1 $23.3 $76.4 

6.0.3 C&DH $42.0 $13.2 $55.2 

Flight software $19.8 $0.0 $19.8 

6.0.4 GN&C $15.2 $10.5 $25.7 

6.0.5 Structures and Mechanical Systems $20.4 $9.7 $30.1 

6.0.6 Power System $12.5 $3.0 $15.5 

6.0.10 Propulsion (Chemical) $1.4 $0.6 $2.0 

6.0.11 Propellant Management (Chemical) $9.6 $4.0 $13.6 

6.0.12 Propellant (Chemical) $0.0 $0.0 $0.0 

6.0.13 Thermal Control (Non-Propellant) $0.0 $1.6 $15.3 

Contingency Mass $6.8 $2.7 $9.4 

Subtotal $174.6 $66.0 $243.3 

Systems Integration $95.6 $42.7 $138.3 

1.6.2.2 IACO $23.4 $32.0 $55.4 

1.6.2.2 STO $23.4 $0.0 $23.4 

1.6.2.2 GSE Hardware $8.9 $0.0 $8.9 

1.6.2.2 SE&I $19.9 $5.4 $25.3 

1.6.2.1 PM $19.9 $5.4 $25.3 

1.6.2.2 LOOS $0.0 $0.0 $0.0 

Total Prime Cost $270.2 $108.7 $378.9 

TABLE 47.—SUBSYSTEM COST BREAKDOWN FOR A SINGLE 

SATELLITE FOR THE FROZEN OPTICAL OPTION IN FY08$M

WBS Description DDT&E total by 

$M

Flight hardware by 

$M

Mfg/DDT&E total 

by $M

6.0.1 RF Communications Package $32.4 $16.8 $49.2 

6.0.2 Optical Communications Package $38.1 $10.1 $48.3 

6.0.3 C&DH $36.2 $6.5 $42.8 

Flight software $19.8 $0.0 $19.8 

6.0.4 GN&C $17.2 $10.5 $27.7 

6.0.5 Structures and Mechanical Systems $20.2 $9.6 $29.8 

6.0.6 Power System $14.9 $3.0 $18.0 

6.0.10 Propulsion (Chemical) $1.3 $0.6 $2.0 

6.0.11 Propellant Management (Chemical) $12.4 $4.6 $17.0 

6.0.12 Propellant (Chemical) $0.0 $0.0 $0.0 

6.0.13 Thermal Control (Non-Propellant) $0.0 $1.2 $10.8 

Contingency Mass $6.7 $2.3 $8.9 

Subtotal $189.1 $63.0 $254.4 

Systems Integration $102.5 $41.2 $143.7 

1.6.2.2 IACO $25.0 $30.9 $55.9 

1.6.2.2 STO $25.0 $0.0 $25.0 

1.6.2.2 GSE Hardware $9.8 $0.0 $9.8 

1.6.2.2 SE&I $21.4 $5.2 $26.6 

1.6.2.1 PM $21.4 $5.2 $26.6 

1.6.2.2 LOOS $0.0 $0.0 $0.0 

Total Prime Cost $291.6 $104.2 $395.8 
NASA/TM—2012-217139 80

TABLE 48—SUBSYSTEM COST BREAKDOWN FOR A SINGLE SATELLITE

FOR THE LAGRANGE RF OPTION IN FY08$M

WBS Description DDT&E total by 

$M

Flight hardware by 

$M

Mfg/DDT&E total 

by $M

6.0.1 RF Communications Package $73.3 $29.0 $102.3 

6.0.3 C&DH $34.4 $13.0 $47.4 

Flight software $19.8 $0.0 $19.8 

6.0.4 GN&C $16.3 $12.7 $29.0 

6.0.5 Structures and Mechanical Systems $34.3 $16.1 $50.4 

6.0.6 Power System $15.4 $5.7 $21.1 

6.0.10 Propulsion (Chemical) $2.7 $2.4 $5.1 

6.0.11 Propellant Management (Chemical) $20.1 $6.2 $26.3 

6.0.12 Propellant (Chemical) $0.0 $0.0 $0.0 

6.0.13 Thermal Control (Non-Propellant) $0.0 $1.7 $16.6 

Contingency Mass $6.6 $2.7 $9.3 

Subtotal $211.4 $86.7 $307.5 

Systems Integration $113.8 $52.8 $166.6 

1.6.2.2 IACO $27.5 $39.2 $66.8 

1.6.2.2 STO $27.5 $0.0 $27.5 

1.6.2.2 GSE Hardware $11.2 $0.0 $11.2 

1.6.2.2 SE&I $23.7 $6.8 $30.5 

1.6.2.1 PM $23.7 $6.8 $30.5 

1.6.2.2 LOOS $0.0 $0.0 $0.0 

Total Prime Cost $325.2 $139.5 $464.8 

TABLE 49—SUBSYSTEM COST BREAKDOWN FOR A SINGLE SATELLITE

FOR THE LAGRANGE OPTICAL OPTION IN FY08$M

WBS Description DDT&E total by 

$M

Flight hardware by 

$M

Mfg/DDT&E total 

by $M

6.0.1 RF Communications Package $40.7 $27.2 $67.9 

6.0.2 Optical Communications Package $38.1 $10.1 $48.3 

6.0.3 C&DH $36.2 $6.5 $42.8 

Flight software $19.8 $0.0 $19.8 

6.0.4 GN&C $19.6 $12.7 $32.3 

6.0.5 Structures and Mechanical Systems $30.4 $14.1 $44.5 

6.0.6 Power System $18.2 $5.7 $23.9 

6.0.10 Propulsion (Chemical) $1.3 $0.6 $2.0 

6.0.11 Propellant Management (Chemical) $21.4 $6.2 $27.6 

6.0.12 Propellant (Chemical) $0.0 $0.0 $0.0 

6.0.13 Thermal Control (Non-Propellant) $0.0 $1.3 $13.1 

Contingency Mass $5.9 $2.3 $8.2 

Subtotal $223.7 $84.4 $310.3 

Systems Integration $116.2 $51.7 $167.8 

1.6.2.2 IACO $28.1 $38.4 $66.5 

1.6.2.2 STO $28.1 $0.0 $28.1 

1.6.2.2 GSE Hardware $11.5 $0.0 $11.5 

1.6.2.2 SE&I $24.2 $6.6 $30.9 

1.6.2.1 PM $24.2 $6.6 $30.9 

1.6.2.2 LOOS $0.0 $0.0 $0.0 

Total Prime Cost $339.8 $136.1 $475.9 
NASA/TM—2012-217139 81

7.2 Risk Analysis and Reduction

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

technology. Risk management is used to innovate and shape the future. Risks are a change to do better 

than planned. 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.

7.2.1 Assumptions

Risk attributes are based on CEV risk values. Risk List is not based on trends or criticality. Some 

mitigation plans are offered as suggestions 

7.2.2 Risk List and Summary

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

in the development life cycle. The risks gathered by the design team are captured in Table 50.

TABLE 50.—RISK SUMMARY TABLE

No. L×C Team Risk title

1 4×5 Communications Multiple data channel feed through

2 2×5 Structures Potential structural failure

3 3×3 Propulsion RCS thruster plume impingement

4 3×3 Communications LNS antenna pointing

5 2×4 Operations Lunar insertion burn accuracy

6 2×4 Avionics Failure of avionics to recover from switch to backup

7 2×2 Power Science data memory overflow

8 1×4 Avionics Major solar burst radiation damage

9 1×4 Operations Dual launch impacts on Constellation

10 1×3 GN&C Navigation sensor view blockage

11 1×2 Power Freezing of liquid propellant in distribution system Likelihood

5

4 1 Criticality

3 3,4 High

2 7 5,6 2 Med

1 11 10 8,9 Low

1 2 3 4 5

Consequences
NASA/TM—2012-217139 82

7.2.3 Trade Space Iterations

Two different orbital constellation options, Frozen orbit and Lagrange orbit, of the available trade 

space of orbit types were investigated in order to weigh the cost and benefits of the communications 

coverage available to a lunar South Pole Constellation human lunar landing site. The summary masses by 

subsystem are shown in Figure 59. For each of these orbital options, two communications technology 

options were measured against one another: an RF communications system and an Optical 

communications system as summarized in Table 51.

Figure 60 shows the comparison of the four LNS spacecraft designs by total wet mass. Each 

subsystem is called out in a separate color. It can be seen that the Lagrange RF spacecraft is the greatest 

wet mass of the four designs.

Table 52 is the top level MEL and the system summary rack up of the case in order to show top level 

masses, total propellant, additional growth carried at the system level for each case. In the optical cases, 

the optical comm. only provided the DTE communications link while RF communications were still used 

to the lunar surface.

Figure 59.—LNS trade space total dry mass comparison by subsystem.

Figure 60.—LNS trade space wet mass comparison by subsystem.
NASA/TM—2012-217139 83

TABLE 51.—LNS-HR TRADE SPACE SUMMARY

WBS

no.

Description

Case—LNS-HR

Frozen 

RF

Frozen

optical

Lagrange

RF

Lagrange

optical

Mass Type Predicted

mass, 

kg

Predicted

mass, 

kg

Predicted

mass, 

kg

Predicted

mass, 

kg

6.0 Lunar Network Satellite-High Rate 1430.9 1406.1 2213.1 2121.7

6.0.1 RF Communications Package 138.0 102.8 248.2 184.1

6.0.2 Optical Communications Package 0.0 23.8 0.0 23.8

6.0.3 C&DH 97.4 97.4 97.4 97.4

6.0.5 GN&C 51.0 51.0 94.2 94.2

6.0.5 Structures & Mechanical Systems 288.4 277.5 362.8 360.6

6.0.6 Power System 107.9 107.9 124.8 124.8

6.0.10 Propulsion (Chemical) 11.7 11.7 20.5 20.5

6.0.11 Propellant Management (Chemical) 49.5 59.2 111.2 111.2

6.0.12 Propellant (Chemical) 617.6 605.4 1084.6 1035.6

6.0.13 Thermal Control (Non-Propellant) 69.5 69.5 69.5 69.5

TABLE 52.—LNS TRADE SPACE MASS RACK-UP

O rbit Type

Comm unica tions Type RF Optical RF Optical

WBS Main Subsystems

Predicted 

(Total) Mass 

(kg)

Predicted 

(Total) Mass 

(kg)

Predicted 

(Total) Mass 

(kg)

Predicted 

(Total) Mass 

(kg)

6.0 Lunar Network Satellie-High Rate 1431 1406 2213 2122

6.0.1 RF Communications Package 138 103 248 184

6.0.2 Optical Communications Package 0 24 0 24

6.0.3 Command & Data Handling (C&DH) 97 97 97 97

6.0.4 Guidance, Navigation and Control System (GN&C) 51 51 94 94

6.0.5 Structures & Mechanical Systems 288 277 363 361

6.0.6 Power System 108 108 125 125

6.0.7 Propulsion (Electric) 0 0 0 0

6.0.8 Propellant Management (EP) 0 0 0 0

6.0.9 Propellant (EP) 0 0 0 0

6.0.10 Propulsion (Chemical) 12 12 21 21

6.0.11 Propellant Management (Chemical) 50 59 111 111

6.0.12 Propellant (Chemical) 618 605 1085 1036

6.0.13 Thermal Control (Non-Propellant) 69 69 69 69

Estimated Spacecraft Dry Mass 813 801 1129 1086

Estimated Spacecraft Wet Mass 1431 1406 2213 2122

Estimated Spacecraft Inert Mass (for traj.) 890 871 1228 1174

Dry Mass Desired System Level Growth 872 854 1197 1144

Mass Margin (carried at system level) 58 53 68 58

Total Wet Mass with Growth 1489 1459 2281 2179

Launch Vehicle Atlas V 511 Atlas V 511 Atlas V 521 Atlas V 521

Available Launch Performance to C3 (kg) 3915 3915 4710 4710

Total Number of LNS-HR Satellites 2 2 2 2

Total mass of Two LNS-HR satellites 2978 2918 4563 4359

Launch margin available (kg) 937 997 147 351

Frozen Orbit Lagrange Orbit

Spacecraft Master Equipment List Rack-up (Mass)
NASA/TM—2012-217139 84

7.3 Case 1—Frozen RF

Case 1 designed an RF communications driven S/C in frozen orbit (Table 53). Launching two LNSHR satellites on an Atlas 511 with a performance to C3 = –2 km2

/sec2 (lunar) of 3915 kg allows for a 

937 kg margin. An additional 58 kg are carried at the system level in order to reach a dry mass growth of 

30 percent.

TABLE 53.—LNS FROZEN RF SYSTEM SUMMARY

COMPASS study: Near Earth Asteroid Sampler Mission Study Date 9/19/08

GLIDE 

container: LunarRelaySat: LNS_Frozen_RF

Spacecraft Master Equipment List Rack-up (Mass)

COMPA SS S/C 

Design 

WBS Main Subsystems

Basc (CBE) 

Mass (kg)

Growth 

(kg)

Predicted (Total) 

Mass (kg)

Aggregate 

Growth (%)

6.0 Lunar Network Satellie-High Rate 1288.1 142.8 1430.9

6.0.1 RF Communications Package 123.1 14.9 138.0 12%

6.0.2 Optical Communications Package 0.0 0.0 0.0 TBD

6.0.3 Command & Data Handling (C&DH) 77.2 42.5 97.4 55%

6.0.4 Guidance, Navigation and Control System (GN&C)42.5 8.5 51.0 20%

6.0.5 Structures & Mechanical Systems 233.7 54.7 288.4 23%

6.0.6 Power System 83.0 24.9 107.9 30%

6.0.7 Propulsion (Electric) 0.0 0.0 0.0 TBD

6.0.8 Propellant Management (EP) 0.0 0.0 0.0 TBD

6.0.9 Propellant (EP) 0.0

6.0.10 Propulsion (Chemical) 10.6 1.1 11.7 10%

6.0.11 Propellant Management (Chemical) 43.9 5.7 49.5 13%

6.0.12 Propellant (Chemical) 617.6

6.0.13 Thermal Control (Non-Propellant) 56.5 13.0 69.5 23%

Estimated Spacecraft Dry Mass 670 143 813 21%

Estimated Spacecraft Wet Mass 1288 165 1431

Estimated Spacecraft Inert Mass (for traj.) 689 201 889.7

System LeveL Growth Calculations Total Growth

Dry Mass Desired System Level Growth 670 201 871.6 30%

Mass Margin (carried a t system level) 58 9%

Total We t Mass with Growth 1288 201 1489.2

Available Launch Perform ance to C3 (kg) 3915.0 kg

Total Num ber o f LNS-HR Sa tellites 2

Total m ass o f Two LNS-HR sa tellites 2978.5

Launch m argin available (kg) 936.5 656
NASA/TM—2012-217139 85

7.4 Case 2—Frozen Optical

Case 2 designed an Optical communications driven S/C with RF communications, where necessary, 

in frozen orbit. Launching two LNS-HR satellites on an Atlas 511 with a performance to C3 = –2 km2

/sec2

(lunar) of 3915 kg allows for a 997 kg margin (Table 54). An additional 53 kg are carried at the system 

level in order to reach a dry mass growth of 30 percent.

TABLE 54.—LNS FROZEN OPTICAL SYSTEM SUMMARY

COMPASS study: Near Earth Asteroid Sampler Mission Study Date 9/22/08

GLIDE 

container: LunarRelaySat: LNS_Frozen_Optical

Spacecraft Master Equipment List Rack-up (Mass)

COMPA SS S/C 

Design 

WBS Main Subsystems

Basc (CBE) 

Mass (kg)

Growth 

(kg)

Predicted (Total) 

Mass (kg)

Aggregate 

Growth (%)

6.0 Lunar Network Satellie-High Rate 1261.9 144.2 1406.1

6.0.1 RF Communications Package 92.3 10.5 102.8 11%

6.0.2 Optical Communications Package 18.1 5.8 23.8 32%

6.0.3 Command & Data Handling (C&DH) 77.2 42.5 97.4 55%

6.0.4 Guidance, Navigation and Control System (GN&C)42.5 8.5 51.0 20%

6.0.5 Structures & Mechanical Systems 225.0 52.5 277.5 23%

6.0.6 Power System 83.0 24.9 107.9 30%

6.0.7 Propulsion (Electric) 0.0 0.0 0.0 TBD

6.0.8 Propellant Management (EP) 0.0 0.0 0.0 TBD

6.0.9 Propellant (EP) 0.0

6.0.10 Propulsion (Chemical) 10.6 1.1 11.7 10%

6.0.11 Propellant Management (Chemical) 51.5 7.8 59.2 15%

6.0.12 Propellant (Chemical) 605.4

6.0.13 Thermal Control (Non-Propellant) 56.5 13.0 69.5 23%

Estimated Spacecraft Dry Mass 657 144 801 22%

Estimated Spacecraft Wet Mass 1262 167 1406

Estimated Spacecraft Inert Mass (for traj.) 674 197 871.3

System LeveL Growth Calculations Total Growth

Dry Mass Desired System Level Growth 657 197 853.5 30%

Mass Margin (carried a t system level) 53 8%

Total We t Mass with Growth 1262 197 1458.9

Available Launch Perform ance to C3 (kg) 3915.0 kg

Total Num ber o f LNS-HR Sa tellites 2

Total m ass o f Two LNS-HR sa tellites 2917.8

Launch m argin available (kg) 997.2 698
NASA/TM—2012-217139 86

7.5 Case 3—Lagrange RF

Case 3 designed an RF communications driven S/C in a Lagrange orbit. Launching two LNS-HR 

satellites on an Atlas 521 with a performance to C3 = –2 km2

/sec2 (lunar) of 4710 kg allows for a 147 kg 

margin (Table 55). An additional 68 kg is carried at the system level in order to reach a dry mass growth 

of 30 percent.

TABLE 55.—LNS LAGRANGE ORBIT RF SYSTEM SUMMARY

COMPASS study: Near Earth Asteroid Sampler Mission Study Date 9/29/08

GLIDE 

container: LunarRelaySat: LNS_Lagrange_RF

Spacecraft Master Equipment List Rack-up (Mass)

COMPA SS S/C 

Design 

WBS Main Subsystems

Basc (CBE) 

Mass (kg)

Growth 

(kg)

Predicted (Total) 

Mass (kg)

Aggregate 

Growth (%)

6.0 Lunar Network Satellie-High Rate 2005.2 207.9 2213.1

6.0.1 RF Communications Package 204.1 44.1 248.2 22%

6.0.2 Optical Communications Package 0.0 0.0 0.0 TBD

6.0.3 Command & Data Handling (C&DH) 77.2 20.2 97.4 26%

6.0.4 Guidance, Navigation and Control System (GN&C)78.5 15.7 94.2 20%

6.0.5 Structures & Mechanical Systems 291.0 71.8 362.8 25%

6.0.6 Power System 96.0 28.8 124.8 30%

6.0.7 Propulsion (Electric) 0.0 0.0 0.0 TBD

6.0.8 Propellant Management (EP) 0.0 0.0 0.0 TBD

6.0.9 Propellant (EP) 0.0

6.0.10 Propulsion (Chemical) 18.7 1.9 20.5 10%

6.0.11 Propellant Management (Chemical) 98.6 12.6 111.2 13%

6.0.12 Propellant (Chemical) 1084.6

6.0.13 Thermal Control (Non-Propellant) 56.5 13.0 69.5 23%

Estimated Spacecraft Dry Mass 921 208 1129 23%

Estimated Spacecraft Wet Mass 2005 208 2213

Estimated Spacecraft Inert Mass (for traj.) 952 276 1228.1

System LeveL Growth Calculations Total Growth

Dry Mass De sired System Level Growth 921 276 1196.8 30%

Mass Margin (carried a t system level) 68 7%

Total We t Mass with Growth 2005 276 2281.4

Available Launch Perform ance to C3 (kg) 4710.0 kg

Total Num ber o f LNS-HR Sa tellites 2

Total m ass o f Two LNS-HR sa tellites 4562.7

Launch m argin available (kg) 147.3 103
NASA/TM—2012-217139 87

7.6 Case 4—Lagrange Optical

Case 4 designed an Optical communications driven S/C with RF communications, where necessary, 

in a Lagrange orbit (Table 56). Launching two LNS-HR satellites on an Atlas 521 with a performance to 

C3 = –2 km2

/sec2 (lunar) of 4710 kg allows for a 351 kg margin. An additional 58 kg is carried at the 

system level in order to reach a dry mass growth of 30 percent.

TABLE 56.—LNS LAGRANGE ORBIT OPTICAL SYSTEM SUMMARY

COMPASS study: Near Earth Asteroid Sampler Mission Study Date 9/30/08

GLIDE 

container: LunarRelaySat: LNS_Lagrange_Optical

Spacecraft Master Equipment List Rack-up (Mass)

COMPA SS S/C 

Design 

WBS Main Subsystems

Basc (CBE) 

Mass (kg)

Growth 

(kg)

Predicted (Total) 

Mass (kg)

Aggregate 

Growth (%)

6.0 Lunar Network Satellie-High Rate 1915.4 206.3 2121.7

6.0.1 RF Communications Package 147.0 37.1 184.1 25%

6.0.2 Optical Communications Package 18.1 5.8 23.8 32%

6.0.3 Command & Data Handling (C&DH) 77.2 20.2 97.4 26%

6.0.4 Guidance, Navigation and Control System (GN&C)78.5 15.7 94.2 20%

6.0.5 Structures & Mechanical Systems 289.3 71.3 360.6 25%

6.0.6 Power System 96.0 28.8 124.8 30%

6.0.7 Propulsion (Electric) 0.0 0.0 0.0 TBD

6.0.8 Propellant Management (EP) 0.0 0.0 0.0 TBD

6.0.9 Propellant (EP) 0.0

6.0.10 Propulsion (Chemical) 18.7 1.9 20.5 10%

6.0.11 Propellant Management (Chemical) 98.6 12.6 111.2 13%

6.0.12 Propellant (Chemical) 1035.6

6.0.13 Thermal Control (Non-Propellant) 56.5 13.0 69.5 23%

Estimated Spacecraft Dry Mass 880 206 1086 23%

Estimated Spacecraft Wet Mass 1915 206 2122

Estimated Spacecraft Inert Mass (for traj.) 910 264 1173.8

System LeveL Growth Calculations Total Growth

Dry Mass De sired System Level Growth 880 264 1143.7 30%

Mass Margin (carried a t system level) 58 7%

Total We t Mass with Growth 1915 264 2179.3

Available Launch Perform ance to C3 (kg) 4710.0 kg

Total Num ber o f LNS-HR Sa tellites 2

Total m ass o f Two LNS-HR sa tellites 4358.7

Launch m argin available (kg) 351.3 246

NASA/TM—2012-217139 89

Appendix A.—Acronyms and Abbreviations

ACS Attitude Control System

AIAA American Institute for Aeronautics and Astronautics

Al aluminum

ANSI American National Standards Institute

ATCS Active Thermal Control System

BER bit error ratio

BOL beginning of life

C&DH Command and Data Handing

C&T Command and Telemetry 

CAD computer aided design

CBE current best estimate

CCAFS Cape Canaveral Air Force Station

CER Cost Estimating Relationship

Comm Communications

COTS commercial off-the-shelf

dLQR design linear-quadratic regulator

DMR design for minimum risk

DPAF Dual Payload Attach Fitting

DPSK-RZ differential phase-shift keying-return-to-zero

DSN Deep Space Network

DTE direct to Earth

EELV Evolved Expendable Launch Vehicles

EIRP equivalent isotropically radiated power

EOL end of life

FEC forward error correction

FOM figure of merit

GaAs gallium arsenide

GEO geosynchronous Earth orbit

GN&C Guidance, Navigation and Control

GRC NASA Glenn Research Center

GSFC NASA Goddard Space Flight Center

hab habitat

He helium

HQ NASA Headquarters

IMDC I M Design Center

IMU inertial measurement unit

ISRU in situ resource utilization

ISS International Space Station

JPL NASA Jet Propulsion Laboratory

KSC NASA Kennedy Space Center
NASA/TM—2012-217139 90

LAN local area network

LCT Lunar Communications Terminal

Li lithium

LNS Lunar Network Satellite

LOI lunar orbit insertion

LPDC low-density parity check

LRO Lunar Reconnaissance Orbiter 

LRS Lunar Relay Station

LSP Launch Service Program

LSTO Launch Service Task Order

LV launch vehicle

LVA launch vehicle adaptor

MGA mass growth allowance

MLI multilayer insulation

MMOD Micrometeoroid and Orbital Debris

MPU Makeup Power Unit

Nav navigation

NLS NASA Launch Services

OMS Orbital Maneuvering System 

OTS off-the-shelf

PMAD Power Management and Distribution

PPM Pulse Position Modulation 

QPSK quadrature phase-shift keying

RAAN right ascension of the ascending node

RF radio frequency

RFI Request for Information

RHU radioisotope heater unit

RTOS real-time operating system

S/C spacecraft

SA solar array

SCAN NASA Space Communications and Navigation

SLOC source lines of code

SMA semimajor axis

SNAP Spacecraft N-Body Analysis Program

SOA state-of-the-art

SOAP Satellite Orbit Analysis Program

SOHO Solar and Heliospheric Observatory

SPA South Pole-Aitken

SPU Solar Power Unit

TBD to be discussed

TBR to be resolved
NASA/TM—2012-217139 91

TDRSS Tracking and Data Relay Satellite System 

Ti titanium

TLI translunar injection

TRL technology readiness level

TT&C Telemetry, Tracking and Command

TWTA traveling wave tube amplifier

USAF U.S. Air Force

USO ultra-stable oscillator

WAN wide area network

WMAP Wilkinson Microwave Anisotropy Probe

NASA/TM—2012-217139 93

Appendix B.—Rendered Images

Figure B.1—LNS-HR satellite, frozen orbit, 

deployed configuration—thruster view.

Figure B.2—LNS-HR satellite, frozen orbit, 

stowed configuration.
NASA/TM—2012-217139 94

Figure B.3—LNS-HR satellite, frozen orbit, 

deployed configuration—antenna view.
NASA/TM—2012-217139 95

Figure B.4—LNS-HR Satellite, frozen orbit, stowed 

configuration—tank face.

Figure B.5—LNS-HR satellite, frozen orbit, stowed 

configuration—top surface.

NASA/TM—2012-217139 97

Appendix C.—LNS Lagrange Orbit 

Rendered Design Drawings

Figure C.1—LNS-HR satellite, Lagrange orbit, deployed configuration.

Figure C.2—LNS-HR satellite, Lagrange orbit, stowed configuration.

NASA/TM—2012-217139 99

Appendix D.—SNAP Input File for Frozen Orbit

Trajectory Propagation Analysis

The following text is the input file for the SNAP trajectory propagation code used to model the frozen 

lunar orbit in Section 2.5.2. This input file should work with snap version xx.

---------begin input deck here, do not include this line------------------ &input

@ There is only 1 phase in this example

nphase=1

precision='double'

@ Input state

tin(1)=20180101.

tin(2)=000001.0

tintyp = 'UTC'

dut = 65.184d0

xintyp(1)='kepler'

xintyp(2)='tod'

xin(1) = 9751.0d0

xin(2) = 1.d-7

xin(3) = 39.2315d0

xin(4) = 50.0d0

xin(5) = 1.707547293e-6

xin(6) = 44.9999996d0

xin(7) = 10000.d0

@----Print Options--------------------------------------

printflg(1)='on'

printflg(2)='on'

printflg(3)='on'

printflg(4)='on'

printflg(5)='on'

printflg(6)='on'

printflg(7)='on'

printflg(26)='on'

printflg(27)='on'

printflg(28)='on'

@----End Print Options----------------------------------

scrflg='on' 

subsurface_stop = 'on'

Moon_equinox_method = 1

@=====Begin Phase 1===================================

title(1)='Propagate in LEO'

strflg(1)='lvlh'

cbody(1) = 'Moon'

obody(1,1) = 'Earth'

obody(2,1) = 'Sun'

@ obody(3,1) = 'jupiter barycenter'

@--- Stop Condition

@-- Propagate for 1 day
NASA/TM—2012-217139 100

stopphs(0,1)='dtime'

stopval(1,0,1)=2557.0d0

stopval(2,0,1)=86400.0d0

@ stopval(2,0,1)=0.0d0

stoptol(0,1)=1.0d-9

@--- Integration Method

inttyp(1)='cart'

relerr(1)=1.d-10

abserr(1)=1.d-10

@-- Output Conditions

stopphs(1,1)='dtime'

stopval(1,1,1)=1.d0

stopval(2,1,1)=86400.d0

stopflg(1,1)='on'

@ lharm(1)=70

@ mharm(1)=70

@ nrmlzd(1)='yes'

@=====End Phase 1=====================================

@-----Load PCK Files-------------------------------------------------

data_fnum=2

data_file(1)='pck_files/planets-Moon.tpc'

data_file(2)='pck_files/Moon_cs.tpc'

@-----End loading PCK files------------------------------------------

@-----Load Ephemeris Files------------------------------

eph_fnum=1

ephfil(1)='bsp_files/de405.bsp'

@-----End loading Ephemeris Files-----------------------

&end

---------end input deck here, do not include this line------------------
NASA/TM—2012-217139 101

Appendix E.—Study Participants

Lunar Network Satellite–High Rate (LNS–HR)—Design Session

Subsystem Name Center Email

Study Lead Kul Bhasin GRC kul.b.bhasin@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

Final Documentation and Editing Les Balkanyi GRC Leslie.R.Balkanyi@nasa.gov

Requirements Tracking Eric Knoblock GRC Eric.J.Knoblock@nasa.gov

Mission Design, ACS, GN&C and 

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

Orbital Analysis and Simulation Mike Martini GRC Michael.C.Martini@nasa.gov

Orbital Analysis and Simulation Carl Sandifer GRC Carl.E.Sandifer@nasa.gov

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

Orbital Analysis and Simulation Laura Burke GRC Laura.M.Burke@nasa.gov

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

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

Structures and Mechanisms John Gyekenyesi GRC john.z.gyekenyesi@nasa.gov

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

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

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

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

Command and Data Handling

(C&DH) Glenn L. Williams GRC glenn.l.williams@nasa.gov

Communications Joe Warner GRC joseph.d.warner@nasa.gov

Communications Abhijit Biswas JPL Abhijit.Biswas@jpl.nasa.gov

Launch Vehicle and Test Jeff Woytach GRC Jeffrey.M.Woytach@nasa.gov

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

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

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

Reliability Joe Hemminger GRC joseph.a.hemminger@nasa.gov
NASA/TM—2012-217139 102

References

1. ANSI/AIAA R-020A-1999, Recommended Practice for Mass Properties Control for Satellites, 

Missiles, and Launch Vehicles

2. Space Mission Analysis and Design, 3rd edition (Space Technology Library), Wiley J. Larson and 

James R. Wertz, eds., October 1999.

3. Sauer, C., “A Users Guide to VARITOP: A General Purpose Low-Thrust Trajectory Optimization 

Program,” Advanced Projects Group, Jet Propulsion Laboratory, November 1991.

4. The Boeing Company, “Delta IV Payload Planner’s Guide,” MDC 00H0043, October 2000. 

5. The Boeing Company, “Delta IV Payload Planner’s Guide Update – April 2002,” MDC 00H0043, 

April 2002. 

6. International Launch Systems (ILS), “Atlas Launch System Planner’s Guide,” Rev. 9, September 

2001. 

7. Morton Thiokol, Inc., “Space Rocket Motors,” June 1987. 

8. http://en.wikipedia.org/wiki/Lagrangian_point

9. http://map.gsfc.nasa.gov/mission/observatory_l2.html

10. Keric Hill, Jeffrey Parker, George H. Born, and Nicole Demandante, “A Lunar L2 Navigation, 

Communication, and Gravity Mission,” AIAA/AAS Astrodynamics Specialist Conference and 

Exhibit 21 - 24 August 2006, Keystone, Colorado AIAA 2006-6662 

11. Todd A. Ely, and Erica Lieb, “Constellations of Elliptical Inclined Lunar Orbits Providing Polar and 

Global Coverage,” AAS/AIAA Astrodynamics Specialists Conference Lake Tahoe, CA, August 7-

11, 2005, AAS 05-343.

12. Daniel J. Grebow, Martin T. Ozimek, and Kathleen C. Howell, David C. Folta, “Multibody Orbit 

Architectures for Lunar South Pole Coverage”, JOURNAL OF S/C AND ROCKETS Vol. 45, No. 2, 

March–April 2008

13. NASA ELV performance estimation website: http://elvperf.ksc.nasa.gov/elvMap/index.html

14. Space Communications References

15. ETS-VI Experimental Optical Inter-Satellite Communication System, Mot okazu SHl KATANl , 

Tadashi ARUGA, Toshi kazu ITABE, Mi t uo ISHIZU, Masao TAKABE, Nor i hi sa HIROMOTO 

and Tadashi SHl OMl, 1999 IEEE

16. Jeffrey L. Kalb, John M. Eldridge, David Heine, David S. Lee, and Jason S. Wertz, “A Study of 

Node-based Architectures for Satellite Systems”, 26th International Communications Satellite 

Systems Conference (ICSSC) 10 - 12 June 2008, San Diego, CA, AIAA 2008-5435 

17. Paul R. Anderson and Lucy Bartamian, “Growth Trends in Communication Satellites and the Impact 

on Satellite System Architecture”, 26th International Communications Satellite Systems Conference 

(ICSSC) 10 - 12 June 2008, San Diego, CA, AIAA 2008-5440 

18. James Schier, “NASA’s Lunar Space Communication and Navigation Architecture”, 26th 

International Communications Satellite Systems Conference (ICSSC) 10 - 12 June 2008, San Diego, 

CA, AIAA 2008-5476.

19. Kul B. Bhasin, Joseph D. Warner., and Lynn M. Anderson, “Lunar Communication Terminals for 

NASA Exploration Missions: Needs, Operations Concepts and Architectures”, 26th International 

Communications Satellite Systems Conference (ICSSC) 10 - 12 June 2008, San Diego, CA, AIAA 

2008-5479.

20. Space Mission Analysis and Design, 3rd edition (Space Technology Library) (Space Technology 

Library) , Wiley J. Larson (Editor), James R. Wertz (Editor), October 1999.

21. ATK/PSI Propellant Storage Tank Specifications, www.psi-pci.com

22. Arde, Inc. Propellant Storage Tank Specifications, www.ardeinc.com

23. Star Motor reference pages from ATK. (http://www.atk.com/starmotors/starmotors_star48v.asp)

24. Aerojet engines and propellant systems (http://www.aerojet.com/capabilities/S/C.php)

25. Thermal References
NASA/TM—2012-217139 103

26. Gilmore, David G. Editor, S/C Thermal Control Handbook: Volume 1 Fundamental Technologies, 

AIAA, 2002.

27. Hyder, A.J., Wiley, R.L., Halpert, G., Flood, D.J. And Sabripour, S. S/C Power Technologies, 

Imperial College Press, 2000.

28. Larson, W.J. And Wertz, J.R., Space Mission Analysis and Design, Third Edition, Space 

Technology Library, Microcosm Press, 1999.

29. Bejan, A. and Kraus, A.D., Heat Transfer Handbook, John Wiley & Sons, 2003.

30. Incopera, F.P. And DeWitt, D.P., Fundamentals of Heat and Mass Transfer, John Wiley and Sons, 

1990.
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Two design options were explored to address the requirement to provide lunar piloted missions with continuous communications for outpost 

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15. SUBJECT TERMS

Moon; Stationkeeping; Parabolic reflectors; Constellations; Propellant; Optical communication; Data links; Communication networks; 

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