Melissa L. McGuire and Steven R. Oleson

Glenn Research Center, Cleveland, Ohio

COMPASS Final Report:

Low Cost Robotic Lunar Lander

NASA/TM—2010-216218

December 2010

CD–2006–03
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Melissa L. McGuire and Steven R. Oleson

Glenn Research Center, Cleveland, Ohio

COMPASS Final Report:

Low Cost Robotic Lunar Lander

NASA/TM—2010-216218

December 2010

CD–2006–03

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135
Acknowledgments

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subject to revision as analysis proceeds.
NASA/TM—2010-216218 iii

Contents

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

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

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

2.2 Assumptions and Approach ................................................................................................3

2.2.1 Survey Starting Points ...........................................................................................3

2.2.2 Fault Tolerance ......................................................................................................5

2.3 Growth, Contingency and Margin Policy ...........................................................................5

2.4 Mission Description............................................................................................................6

2.4.1 Mission Analysis Assumptions..............................................................................6

2.4.2 Main Mission Trajectory Options..........................................................................6

2.4.3 Mission Analysis Event Timeline..........................................................................8

2.4.4 Mission Trajectory Details.....................................................................................8

2.5 Small Launch Vehicle Details ............................................................................................9

2.5.1 Minotaur ................................................................................................................9

2.5.2 SpaceX—Alternate Launch Vehicle Option..........................................................9

2.6 System Design Trade Space: Preliminary Analysis..........................................................10

2.7 Baseline System Design....................................................................................................10

3.0 Baseline Design ............................................................................................................................. 12

3.1 Top Level Design (MEL and PEL)...................................................................................12

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

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

3.2 System Level Summary ....................................................................................................13

3.3 Design Concept Drawing and Description........................................................................13

4.0 Subsystem Breakdown................................................................................................................... 14

4.1 Communications...............................................................................................................14

4.1.1 Communications Requirements...........................................................................14

4.1.2 Communications Assumptions............................................................................15

4.1.3 Communications Design and MEL......................................................................15

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

4.1.5 Communications Analytical Methods..................................................................16

4.1.6 Communications Risk Inputs...............................................................................16

4.1.7 Communications Recommendation.....................................................................16

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

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

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

4.2.3 Avionics Design and MEL ..................................................................................17

4.2.4 Avionics Trades...................................................................................................18

4.2.5 Avionics Analytical Methods ..............................................................................18

4.2.6 Avionics Risk Inputs............................................................................................19

4.2.7 Avionics Recommendation..................................................................................19

4.3 Electrical Power System...................................................................................................19

4.3.1 Electrical Power Requirements............................................................................19

4.3.2 Electrical Power Assumptions.............................................................................19

4.3.3 Electrical Power Design and MEL ......................................................................19

4.3.4 Electrical Power Trades.......................................................................................20

4.3.5 Electrical Power Analytical Methods..................................................................20

4.3.6 Electrical Power Risk Inputs................................................................................20

4.3.7 Electrical Power Recommendation......................................................................20

4.4 Structures and Mechanisms ..............................................................................................21
NASA/TM—2010-216218 iv

4.4.1 Structures and Mechanisms Requirements..........................................................21

4.4.2 Structures and Mechanisms Assumptions ...........................................................21

4.4.3 Structures and Mechanisms Design and MEL.....................................................21

4.4.4 Structures and Mechanisms Trades .....................................................................22

4.4.5 Structures and Mechanisms Analytical Methods.................................................22

4.4.6 Structures and Mechanisms Risk Inputs..............................................................25

4.4.7 Structures and Mechanisms Recommendation ....................................................25

4.5 Propulsion and Propellant Management ...........................................................................25

4.5.1 Propulsion and Propellant Management Requirements.......................................25

4.5.2 Propulsion and Propellant Management Assumptions........................................25

4.5.3 Propulsion and Propellant Management Analytical Methods..............................26

4.5.4 Propulsion and Propellant Management Design and MEL..................................29

4.5.5 Propulsion and Propellant Management Trades..................................................30

4.5.6 Propulsion and Propellant Management Risk Inputs...........................................31

4.5.7 Propulsion and Propellant Management Recommendation.................................31

4.6 Thermal Control................................................................................................................31

4.6.1 Thermal Requirements.........................................................................................31

4.6.2 Thermal Assumptions..........................................................................................31

4.6.3 Thermal Design and MEL ...................................................................................32

4.6.4 Thermal Trades....................................................................................................35

4.6.5 Thermal Analytical Methods...............................................................................35

4.6.6 Thermal Risk Inputs.............................................................................................37

4.6.7 Thermal Recommendation...................................................................................37

5.0 Cost, Risk and Reliability .............................................................................................................. 37

5.1 Costing: Baseline Chemical Lunar Lander.......................................................................37

5.2 Cost Modeling Assumptions.............................................................................................37

5.3 Cost Modeling Approach..................................................................................................37

6.0 Trade Space Iterations.................................................................................................................... 38

6.1 Case 1: Off-the-Shelf Chemical Propulsion .....................................................................38

6.2 Case 2: Off-the-Shelf Electric Propulsion ........................................................................38

6.3 Case 3: Advanced Direct Drive Electric Propulsion.........................................................39

Appendix A.—Acronyms and Abbreviations............................................................................................. 43

Appendix B.—Case 1 Rendered Design Drawings.................................................................................... 45

Appendix C.—Study Participants............................................................................................................... 47

Bibliography ............................................................................................................................................... 48
NASA/TM—2010-216218 1

COMPASS Final Report:

Low Cost Robotic Lunar Lander

Melissa L. McGuire and Steven R. Oleson

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

1.0 Executive Summary

In the Fall of 2006 the NASA Headquarters (HQ) Lunar Architecture Team/Lunar Precursor and 

Robotic Program (LAT/LPRP) (Program) leader Tony Lavoie was interested in how much payload can an 

inexpensive chemical or Electric Propulsion (EP) system deliver to the Moon’s surface. Subsequently 

management at the NASA Glenn Research Center (GRC) (Rob Jankovsky et al.), were contacted by the 

NASA Ames Research Center (ARC) and asked if the COMPASS team could model a small lunar 

lander mission given a low cost launch vehicle and its performance characteristics. The goal of this 

COllaborative Modeling for the Parametric Assessment of Space Systems (COMPASS) session was to 

use Total Low Cost as the objective function, and design a Robotic Lunar Lander to deliver an 

unspecified payload (greater than zero) to the lunar surface for the lowest cost.

The spacecraft designed as the baseline out of this study was a solar powered robotic lander, launched 

on a Minotaur V launch vehicle on a direct injection trajectory to the lunar surface. A Star 27 solid rocket 

motor does lunar capture and performs 88 percent of the descent burn. The Robotic Lunar Lander softlands using a hydrazine propulsion system to perform the last 10 percent of the landing maneuver, leaving 

the descent at a near zero, but not exactly zero, terminal velocity. This low-cost robotic lander delivers 

10 kg of science payload instruments to the lunar surface.

Collected in Table 1.1 is the top-level summary of each of the subsystems in the Robotic Lunar 

Lander design.

TABLE 1.1—MISSION AND SPACECRAFT SUMMARY

Subsystem area Details Total mass 

with growth

Top Level Robotic 

Lunar Lander

One week, Sunlit, global access, 25 W for payload on lunar surface 462 kg

Mission and 

Operations

Chemical direct injection mission (a.k.a. Surveyor), Minotaur V to TLI and direct lunar 

injection, solid rocket motor for descent, hydrazine for landing

Guidance, 

Navigation and 

Control (GN&C)

Avionics hardware located on top of lander for improved cooling: Star trackers, Sun

sensors, flight control electronics and radar altimeter for landing

24 kg

Launch Minotaur V Launch, 92 in. payload fairing

Science Science payload as yet undetermined. Used as a metric in the design. 15 kg

Power 125 W GaAs solar array (two-axis pointing) on landed vehicle, 100 W housekeeping, 

25 W to science instruments

31 kg

Propulsion Descent motor: Star 27 Solid Rocket (disposed before landing), hydrazine propulsion 

system for final landing

44 kg

Structures and 

Mechanisms

Mechanisms to separate Star 27, and point solar arrays.

Structural designed for 12 g axial launch, 5g axial landing loads, AL 2090 material, 

hexagonal bus with three Fixed leg landing system using Doppler radar (antennas on feet)

48 kg

Communications High gain antenna (HGA) on lander 8 kg

Command and 

Date Handling 

(C&DH)

General processor (high speed, 1394 B), harness, wiring 10 kg

Thermal Fixed radiator panels, heat sinks, heaters, multilayer insulation (MLI) blankets 22 kg
NASA/TM—2010-216218 2

Figure 1.1—Robotic Lunar Lander shown on top of Star 27 solid rocket motor.
NASA/TM—2010-216218 3

2.0 Study Background and Assumptions

2.1 Introduction 

The goal for this study was to design the architecture necessary to deliver a low cost Robotic Lander 

to the Moon’s surface using a small class launch vehicle. The overall study objective was to provide a low 

cost (~$100M) capability to place small payloads (10’s of kilograms of science or tech demos) on the 

Moon. The science and power goals were as follows:

 One week mission duration on sunlit portion of the Moon (i.e., approximately on half of an 

average lunar day from “Sun up” to “Sun down” at a location on the surface of the Moon. This is 

dependent on choice of latitude/longitude of the landing site)

 Provide global access to the science payload, i.e., deliver science anywhere on the lunar surface

2.2 Assumptions and Approach

Starting with similar missions, the following missions were looked at as historic references for the 

COMPASS design session:

 Start with previous missions

− Lunar

• Surveyor (http://www.lpi.usra.edu/expmoon/surveyor/surveyor.html) 

♦ Landed ~33 kg of payload with 1960s technology off of the first Atlas Centaur 

vehicle

♦ Lunar Prospector (http://lunar.arc.nasa.gov/)

♦ Clementine (http://www.cmf.nrl.navy.mil/clementine/)

• Mars

♦ Viking (http://www.nasa.gov/mission_pages/viking/)

♦ Pathfinder (http://marsprogram.jpl.nasa.gov/MPF/)

♦ Mars Polar Lander (http://mars.jpl.nasa.gov/msp98/)

♦ Mars Exploration Rover (MER) (http://marsrover.nasa.gov/home/index.html)

Starting points for the specific areas of the design are

 Power available for payload: 25 W

 Launch vehicle options to consider

− Minotaur I, IV, V (http://www.orbital.com/SpaceLaunch/Minotaur/index.html)

− Falcon 1

 Stages to design: TLI, Lunar capture and descent, Lander

 Propulsion Options

− Chemical: Solids, etc.

− Electric Propulsion

− Combination of both

2.2.1 Survey Starting Points

The Surveyor spacecraft were used as starting points in this design. From the historical JPL website, 

the description of the Surveyor program reads as follows:

The Surveyor program consisted of seven unmanned lunar missions that were launched between May 

1966 and January 1968. Five of these spacecrafts, Surveyor 1, 3, 5, 6, and 7 successfully soft-landed 

on the lunar surface. In addition to demonstrating the feasibility of lunar surface landings, the 
NASA/TM—2010-216218 4

Surveyor missions obtained lunar and cislunar photographs and both scientific and technological 

information needed for the Apollo manned landing program. Four spacecraft, Surveyor 1, 3, 5, and 6, 

returned data from selected mare sites from Apollo program support, and Surveyor 7 provided data 

from a contrasting rugged highland region.

For a sanity check in the bottoms up subsystem mass numbers, Table 2.1 calls out the main 

subsystems in the surveyor spacecraft and their relative magnitude of total dry mass.

TABLE 2.1—SURVEYOR SUBSYSTEM 

RELATIVE MASS DISTRIBUTION

Surveyor reference Mass 

(kg) 

Percent

of dry 

Flight control 32.8 8 

Communications 38.7 9 

Radar 15.4 4 

Propulsion 99.7 24 

Power 48.6 12 

Mechanical/structural 142.8 35 

Science 33.5 8 

Dry total 411.5 

From the Surveyor Web site, the surveyor spacecraft are described as:

Each spacecraft weighed 1000 kg at launch, was 3.3 m high, and had a 4.5-m diameter. The 

tripod structure of aluminum tubing provided mounting surfaces for scientific and engineering 

equipment. Onboard equipment consisted of a 3-m-square solar panel that provided 

approximately 85-W output, a main battery and 24-V nonrechargeable battery that together 

yielded a 4090-W total output, a planar array antenna, two omnidirectional antennas, and a 

radar altimeter. The soft landing was achieved by the spacecraft free falling to the lunar surface 

after the engines were turned off at a 3.5-m altitude. Operations began shortly after landing.

Figure 2.1 is the surveyor spacecraft with major subsystem components identified.

Figure 2.1—Surveyor Lander.
NASA/TM—2010-216218 5

2.2.2 Fault Tolerance

In the interests of minimizing costs this design is assumed to use a zero-fault tolerant approach.

2.3 Growth, Contingency and Margin Policy

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

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

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

TABLE 2.2—PERCENT MASS GROWTH ALLOWANCE (MGA)

Code

Design Maturity

(Basis for Mass 

Determination)

Percent Mass Growth Allowance

Electrical/Electronic 

Components

StructureThermal ControlPropulsionBatteriesWire Harnesses

MechanismsInstrumentation

0 to 5 

kg

5 to 

15 kg >15 kg

E Estimated

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

L

Layout

(or major modification 

of existing hardware)

25 20 15 12 12 12 15 30 12 30

P

Pre-Release Drawings

(or minor modification 

of existing hardware)

20 15 10 8 8 8 10 25 8 25

C Released Drawings

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

X

Existing Hardware

(actual mass from 

another program)

3 3 3 2 2 2 3 3 2 3

A

Actual Mass

(measured flight 

hardware)

0 0 0 0 0 0 0 0 0 0

CFE Customer Furnished 

Equipment 0 0 0 0 0 0 0 0 0 0

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

design is calculated. The COMPASS system designed to total growth of 30% 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% 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.

From the Discovery AO (Announcement of Opportunity): Definitions of Contingency and Mass 

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

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

contingency.

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

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

available margin divided by its maximum expected value.

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

modeling the power system. See Section 3.1.2 for the power system assumptions.
NASA/TM—2010-216218 6

2.4 Mission Description

2.4.1 Mission Analysis Assumptions

2.4.1.1 Landing Options

The following set of bullets describes the options for landing a Robotic Lunar Lander that were 

discussed during the brain storming portion of the COMPASS design sessions. The powered landing was 

chosen in the baseline chemical configuration.

 Powered hover or ‘drop’

 Terminal Systems

− Legs (crushable)

− Vented Airbag (no bounce)

• The 100 kg rover/instrument example shows feasibility of landing from drop height of 

100 m. At this height, the landing velocity is 18 m/s.

• It is assumed that a payload will be oriented and released so that the bag side remains 

down (without tumble) and the package will need to land with a single bag stroke. Vent 

patches can be designed so as to attenuate most of the kinetic energy in the impact and 

keep the package from contacting the surface and rebounding. 

• With a 1/6 G environment, a rebounding bag similar to the Mars rover landing could 

bounce for a significant amount of time and could settle in an undesirable location. 

Deceleration rates are highly variable depending on the design chosen. However, the 

need for the bag appears feasible to keep the landed package below 30 G’s to avoid 

special shock hardened technology for the instruments. 

(http://marsrovers.nasa.gov/mission/spacecraft_edl_airbags.html)

• BEAGLE—The British-led BEAGLE2 project is the probe for the European Space 

Agency’s (ESA) Mars Express mission. This airbag design makes use of a multisectioned 

bag assembled to form a single sphere. The system went through extensive and successful 

testing at NASA’s Plum Brook and JSC facilities. The airbags were integrated to the 

Lander in January 2003, and launched the following June. Unfortunately the airbags did not 

have an opportunity to function, due to an undetermined overall mission failure 

− Survivable penetrator

− Crushable engines/structure (dual use structure)

2.4.1.2 Main Propulsion Options

Table 2.3 shows the comparison between case 1 (baseline—chemical) and Cases 2 and 3 using EP for 

lunar transfer.

2.4.2 Main Mission Trajectory Options

2.4.2.1 Ballistic Earth-Moon Transfer (Direct Hohmann)

The baseline case chosen used a direct chemical injection Hohmann transfer to the Moon. The launch 

vehicle injects the Robotic Lunar Lander/Star motor stack directly to Trans-Lunar Injection (TLI). The 

Star Motor performs approximately 88% of the descent burn, and a hydrazine propulsion system on the 

lander performs the rest for a soft landing. 

Table 2.4 breaks out the three stages in this transfer. Because of the generic design of the system 

tools, there is a line in the table in the Lunar Orbit Insertion (LOI) stage for useable propellant. In the 

baseline Case 1 detailed in this report, the Minotaur V rocket is used to perform a TLI burn direct 

injection to the Moon, therefore there is no useable propellant carried on the Robotic Lander portion of 

the vehicle. Cases 2 and 3 would have a useable propellant number in this cell where the EP system’s 

Xenon (Xe) propellant would be book-kept.
NASA/TM—2010-216218 7

TABLE 2.3—MISSION ANALYSIS DELTA V COMPARISON

Case 1—Chemical Propulsion Case 2—Electric Propulsion 

Phase name Main DV main 

propulsion item 

RCS DV RCS 

propulsion item 

Main DV main 

propulsion item 

RCS DV RCS 

propulsion item 

Launch from Earth

Checkout 

Loiter to TLI window opening 

TLI opening to ignition 

TLI burn 

TLI stage disposal 10 m/s Lander 10 m/s Lander 

Trans-Earth mid-course corrections

Trans-lunar coast 

20 m/s Lander 10 m/s Lander 20 m/s Lander 10 m/s Lander 

Lunar orbit capture burn coast

LOI plane change burn 

Coast 

Lunar orbit circularization burn 

4545 m/s EP Stage 

Lunar descent burn 2191 m/s Descent stage 11 m/s Lander 1499 m/s Descent stage 11 m/s Lander 

Lunar landing burn 299 m/s Lander 224 m/s Lander 

Launch: Minotaur V launch to TLI

Delivers 550 kg 

TLI: Performed by launch vehicle 

LOI: None, direct landing

Landing: Direct from TLI

Star 27 performs 88% of the descent 

Vernier system used for remainder 

Launch: Minotaur V launch to GTO

Delivers 710 kg 

TLI: Performed by EP stage 

LOI: Performed by EP stage 

Landing: Descend from LLO

Star 17A performs 88% of the descent

Vernier system used for remainder 

TABLE 2.4—CHEMICAL DIRECT TRANSFER STAGING

LOI stage

Useable propellant, kg..................................................... 0.0 

Total wet mass, kg ........................................................ 29.5 

Inert mass, kg .................................................................. 0.0 

EDS adaptor mass + 30%, kg ....................................... 29.5 

Descent stage

Useable propellant, kg................................................. 238.3 

Total wet mass, kg....................................................... 266.8 

Inert mass, kg ................................................................ 28.5

Lander stage

Useable propellant, kg................................................... 30.3 

Total wet mass, kg....................................................... 165.7

Inert mass, kg .............................................................. 135.5

2.4.2.2 Earth-Moon Transfer Weak Stability Boundary (WSB) Analysis

Belbruno (AIAA–2000–4142) and Parker (AAS 06–132) and others (Lo, Koon, etc.) have 

demonstrated ballistic lunar transfers via the “weak stability boundary” (WSB). WSB transfers require 

more ∆V for TLI than a direct Hohmann-like (Apollo) lunar transfer, but the spacecraft is captured by 

solar/lunar gravity into a lunar libration orbit without any propulsive ∆V. Apogee is ~1.5×106 km (235 

Earth radii), where the Sun’s gravity perturbs the trajectory and raises perigee to lunar-centered orbit of 

radius (384,000 km). Typically, the spacecraft is captured into a “halo” orbit about the lunar libration 

point L2 (far side of the Moon on Earth-Moon line). The spacecraft can leave the unstable halo orbit (for 

zero ∆V) and follow the invariant manifold (surface of trajectories) to the vicinity of the Moon. WBS 

transfers also require much longer trip time (>90 days to the lunar libration orbit).
NASA/TM—2010-216218 8

Figure 2.2—Baseline mission event timeline illustration.

2.4.3 Mission Analysis Event Timeline

Figure 2.1 shows the major events in the baseline all-chemical propulsion option (case 1). The 

Minotaur V Launch Vehicle lifts the Robotic Lunar Lander to TLI using five stages. The Star 27 motor 

performs the direct lunar descent maneuver and then is jettisoned. A small hydrazine propulsion system 

on the Robotic Lander performs the last landing maneuver to provide controlled touchdown and safe 

landing of the science instruments.

2.4.4 Mission Trajectory Details

For comparison, Table 2.5 presents the Delta V at different burns along the trans-lunar mission. The 

direct Hohmann type transfer was used in the baseline mission design. The WSB trajectory analysis was 

left for another mission if an Electric Propulsion option were found that delivered positive payload. Note 

that (TOF) is time of flight, or the total trip time to the Moon. A standard direct, Hohmann-type transfer 

takes approximately 4 days.

For the comparison from low Earth orbit (LEO) to low lunar orbit (LLO) the initial LEO was 

assumed to be 300-km altitude circular orbit and the target LLO was a 100-km altitude circular polar 

orbit.

TABLE 2.5—TRAJECTORY DELTA V TOTALS LEO TO LLO

Trajectory Method DVTLI

(km/s)

DVLOI

(km/s)

DVtotal

(km/s)

TOF

(days)

WSB 3.176 0.664 3.840 100.5

Direct (Hohmann) 3.107 0.819 3.926 4.1

Table 2.6 shows the Delta V’s for comparing the mission phases of a WSB trajectory to a direct lunar 

transfer assume impulsive Delta V for landing (no gravity losses and no LLO parking orbit). For the 

direct transfer trajectory, assume perilune of incoming hyperbola grazes lunar surface. The trajectories 

were run starting from Initial LEO altitude of 300 km (circular) and target was lunar surface.

TABLE 2.6—TRAJECTORY DELTA V TOTALS LEO TO LUNAR SURFACE

Trajectory Method DVTLI

(km/s)

DVLAND

(km/s)

DVtotal

(km/s)

TOF

(days)

WSB 3.176 2.365* 5.541 100.5

Direct (Hohmann) 3.107 2.490 5.597 4.1
NASA/TM—2010-216218 9

2.5 Small Launch Vehicle Details

A number of small class launch vehicles were examined for use in this mission. The Minotaur and 

Falcon class vehicles were the two that were most appropriate to the launch mass required for the Robotic 

Lunar Lander.

2.5.1 Minotaur

Minotaur V is a five-stage evolutionary version of the Minotaur IV space launch vehicle (SLV) to 

provide an extremely cost-effective capability to launch small spacecraft into high-energy trajectories, 

including geosynchronous transfer orbits (GTO) and trans-lunar missions. 

The first three stages of the Minotaur V are the unmodified Peacekeeper solid rocket motors. The 

stage four motor is nominally a Star-48GV. The fifth stage can be either attitude controlled or spinning. 

The attitude-controlled version nominally uses the same Orion-38 motor that has been extensively flight 

demonstrated on multiple Orbital launch vehicles, including Pegasus, Taurus, and Minotaur I. For a 

spinning configuration, a Star-37FM is used to provide maximum performance.

The Minotaur V avionics, structures, and fairing are common with the Minotaur IV SLV, with 

relatively minor changes to create the five-stage configuration. Moreover, the avionics and flight software 

are highly common across all Minotaur family vehicles (Orbital Sciences Minotaur V Fact Sheet pdf).

Figure 2.3 shows the “projected performance to C3” values of the Minotaur V launch vehicle taken 

from the Orbital Sciences Minotaur V Fact Sheet.

2.5.2 SpaceX—Alternate Launch Vehicle Option

For comparison with the Minotaur launch vehicle family, the SpaceX Falcon launch vehicles were 

looked at as a possible launcher for this payload. Figure 2.4 is a graphic taken from the SpaceX site listing 

the relative characteristics of the family of Falcon launch vehicles. (http://www.spacex.com/falcon1.php).

Figure 2.3—Minotaur V launch vehicle performance curve.
NASA/TM—2010-216218 10

Figure 2.4—Space X Falcon launch vehicle family.

2.6 System Design Trade Space: Preliminary Analysis

The preliminary analysis on landed mass using the small launch vehicle performance shows that a 

150 to 200 kg Landed Mass is feasible. For EP the trade is: Minotaur I with EP (150 kg LM) versus 

Minotaur IV with Chemical (170 kg LM). 

TABLE 2.7—SYSTEM DESIGN SPACE PRELIMINARY MISSION SIZING

Chemical missions Type Estimate

landed mass Assumptions

Minotaur I LEO-Direct 50 kg 12% dry stage fraction assumed

Minotaur IV LEO-Direct 170 kg 12% dry stage fraction assumed

Minotaur V TLI-Direct 215 kg Solid (Isp = 285 sec)

Minotaur V TLI-Direct 150 kg Monoprop (Isp = 225 sec)

SEP missions Type Estimate

landed mass Assumptions

Minotaur I LEO-LLO 148 kg 470 day total trip time

Minotaur IV LEO-LLO 500 kg 471 day total trip time

Secondary P/L GTO-LLO 170 kg 500 kg assumed initial mass

2.7 Baseline System Design

The basic design consists of launch on a Minotaur V to TLI on a direct trajectory (similar to 

Surveyor) to the Moon. The propulsion system is a combination of solid stage for descent and a 

Monoprop Hydrazine propulsion system on the Lander for landing. The TLI is a 4-day transfer, landing 

approximately 15 kg useable payload on lunar surface. 

Figure 2.5 shows the Robotic Lunar Lander from two angles, with all major components labeled. The 

Star 27 motor is shown attached along with the spacecraft adaptor. The landing legs are extended and 
NASA/TM—2010-216218 11

both the solar array and the antenna are deployed. Figure shows a simple drawing of the Robotic Lander 

on the Star 27 motor with the Minotaur V adaptor at the bottom. The landing legs, antenna, and solar 

array panel are deployed. Figure 2.7 shows the rendered lander portion of the baseline design after the star 

motor and associated adaptor have been jettisoned.

Figure 2.5—Baseline Robotic Lunar Lander/Star 27 stack with component labels.

Figure 2.6—Baseline Robotic Lunar Lander/Star 27 stack.

Figure 2.7—Baseline Robotic Lunar Lander: 

Lander portion.
NASA/TM—2010-216218 12

3.0 Baseline Design 

3.1 Top Level Design (MEL and PEL)

For the purpose of modeling in the Master Equipment List (MEL), this baseline design is broken into 

three stages: the LOI stage, the descent stage, and the lander stage. The LOI stage is the structure 

necessary to attach the Star 27 rocket to the Minotaur V rocket. The decent stage is the Star 27 Rocket, 

and the Lander stage is the Robotic Lunar Lander (structure, power, thermal, etc.) and science payload 

package along with the hydrazine system used for landing.

3.1.1 Master Equipment List (MEL)

Table 3.1 MEL is the top level MEL tool developed in the COMPASS design session. Note that the 

Growth (%) column displays the total top level growth percentage for the wet portions of the vehicle. This 

is not the way growth is used in the design session, but is still calculated in the tool. The only growth 

percentages to refer to are on the dry masses, not the total wet mass (i.e., with propellant). The Growth 

column in kg is the summation of the system level growths of the dry masses. These calculations are a 

function of the tool. For more clear growth calculations, see the system summary Table 3.3. The total 

growth percentage of the wet mass calculation need to be refined in future versions of the MEL tool. The 

total growth percentage using the wet mass number in the percentage calculation gives a deceptive 

percentage number. Growth percentage is calculated growth divided by total current base estimate (CBE) 

mass. When the wet mass is used instead of the dry in the CBE mass, the growth percentage is skewed.

TABLE 3.1—MEL

WBS Description CBE mass

(kg)

Growth

(%)

Growth

(kg)

Total mass 

(kg)

01 Robotic Lunar Lander 429 7 33.3 462

01.01 Lunar Orbit Insertion Stage 25 15 4.5 29.5 

01.02 Lunar Descent Stage 266 0 1.1 266.8 

01.02.02 

01.02.03

01.02.04

Propulsion (Chemical) 

Propellant Management (Chemical) 

Propellant Hardware (Chemical) 

27 

0 

238 

4

0

0 

1.1 

0.0 

0.0 

28.5 

0.0 

238.3 

01.03 Lunar Lander Stage 138 16.7 27.7 165.7 

01.03.01 Avionics, Instrumentation, GN&C 35 16 6.8 42.0 

01.03.02 Propulsion Hardware (Chemical) 5 5 0.3 5.6 

01.03.03 

01.03.04

01.03.05

Propellant Management (Chemical) 

Propellant (Chemical) 

Propulsion Hardware (EP) 

9 

30 

0 

6

20

0

0.5 

7.6 

0.0 

9.6 

37.8

0.0 

01.03.06 Propellant (EP) 0 0 0.0 0.0 

01.03.07 

01.03.08 

01.03.09 

01.03.10 

01.03.11 

Propellant Management (EP) Power 

Processing (EP) 

Power 

Thermal Control (Non-Propellant) 

Structures & Mechanical Systems 

0 

0 

24 

18 

15

0

0

21

15

15

0.0 

0.0 

6.5 

3.3 

2.8

0.0 

0.0 

30.9 

21.7 

18.2 

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

item by item from the subsystem designer, Table 3.3, wraps up those total masses, CBE and total mass 

after applied growth percentage. In order to meet the total of 30% at the system level, an allocation is 

necessary for system level growth. 

3.1.2 Power Equipment List (PEL)

The PEL for nominal loads is show in Table 3.2.
NASA/TM—2010-216218 13

TABLE 3.2—PEL

Subsystem Peak power 

(W)

Keep alive power 

(W)

Science payload 25 10

Thermal control 20 --

Communications 34 --

Avionics 15 --

Total to size solar arrays 94 --

3.2 System Level Summary

The system level summary will track the total MEL and add on the margins and growths per the 

ANSI/AIAA R-020A-1999 growth schedule as described in Section 2.4. The COMPASS team used 10% 

of Launch Mass as the launch Margin set aside. Margin is not flown, nor can it be used as performance by 

the ELV. The hydrazine propellant growth was assumed at 25% (per the definitions in Space Mission 

Analysis and Design (SMAD)).

The total mass of the robotic lander is calculated by summing the total masses of the LOI stage 

(30 kg), the descent stage (267 kg) and the Lander stage (166 kg). This sum was 462 kg. Subtracting the 

propellant (238 kg on the descent stage 38 kg on the lander stage) from this number leaves a total dry 

mass of 186 kg. The total inert mass is calculated by subtracting the total useable propellant (268 kg 

summed from the stages in Table 2.3) from the ideal rocket equation from the total lander wet mass. This 

leaves 8 kg (total propellant—useable propellant) of total residuals trapped in the tanks and carried as 

inert in terms of the system mass that the propulsion system must push through the application of the 

Delta V. These calculations are shown in Table 3.3 in the system summary chart.

The system summary chart shows the top level total masses with subsystem growth margin applied for each of 

the three stages in the Robotic Lunar Lander: LOI stage, descent Stage, and lander stage.

TABLE 3.3—CASE 1 (BASELINE) SYSTEM SUMMARY CHART

Robotic Lander Stage Masses 

(kg)

Current values with growth margin 

(kg)

Totals Percent of dry 

mass 

Subsystems LOI 

stage 

Descent 

stage 

Lander

stage 

Total Robotic 

Lander 

Total Robotic 

Lander 

Avionics 0 0 42 42 23 

Propulsion (chemical) 0 29 6 34 18 

Prop Management (chemical) 0 0 10 10 5 

Propellant (chemical) 0 238 38 276 N/A

Power 0 0 31 31 17 

Thermal Control 0 0 22 22 12 

Structure 30 0 18.15 48 26 

Total Mass (Wet) 30 267 166 462 N/A

Lander Total Wet Mass 462 

Lander Total Dry mass (kg) 186 

Total Residuals 8 

Lander Total Inert Mass (kg) 194 

ELV Performance to TLI 477 

Gross Payload Capability 14.9 

At the time of this analysis, there was no additional growth carried at the system level for this design. 

The growth is entered in by the subsystem lead per instructions in Table 3.1 MGA Schedule.

3.3 Design Concept Drawing and Description

Figure 3.1 shows the side view of the Robotic Lunar Lander on top of the Star 27 motor, with the 

payload adaptor truss structure. Figure 3.2 shows the top down view of the Robotic Lunar Lander with 

dimensions.
NASA/TM—2010-216218 14

Figure 3.1—Baseline Robotic Lunar Lander: Total stack with dimensions.

Figure 3.2—Baseline Robotic Lunar Lander: Top down view with dimensions.

4.0 Subsystem Breakdown 

4.1 Communications

4.1.1 Communications Requirements

The Spacecraft RF Power level is driven by data transmission needs of the science payload. Since 

there is no defined science payload at this step in the design, the following two items are used as the 

requirements for data levied on the communications system.
NASA/TM—2010-216218 15

 Communications Data Rate

− Telemetry: 6 kbps

− Command: up to 2 kbps

4.1.2 Communications Assumptions

 Deep Space Network (DSN) or Tracking and Data Relay Satellite System (TDRSS) for Earth 

antenna tracking capability tone ranging 

 Communications system incorporates radio frequency (RF) transponder and RF antenna functions

 Transponder handles RF/digital signals between the antenna and digital interface

 RF amplification/attenuation, diplexing, mod/demod, synchronization, ranging signal turn-around 

(navigation (nav) transponder)

 Four conical spiral omnidirectional antennas have no active pointing mechanism

 Best antenna selected based on strongest received signal strength (antenna pointing at source)

 Signal processor performs encoding, decryption, and frame synchronization

 Space serial data system through IEEE 1394 bus (or 1553B bus) 

 State vector requirements: (where am I, what direction, how fast?)

− Doppler/tone ranging for state vector generation command data 

− Uplink—provides state vector, and timing updates, and real time commands

− Downlink—vehicle telemetry data

4.1.3 Communications Design and MEL

The only communications hardware was on the Lunar Lander stage. This is the stage with the 

structure, power and science instruments ultimately landed on the lunar surface. There is no 

communications hardware on the LOI or Descent stages. Figure 4.1 shows the schematic of the design of 

the Lunar Lander communications system. Table 4.1 is the communications system bottoms up MEL. The 

Growth percentage is only important on the line items. Refer to the growth (kg) totals for details.

Figure 4.1—Lunar Lander communications system.
NASA/TM—2010-216218 16

TABLE 4.1—COMMUNICATIONS SYSTEM MEL

WBS Description

Qty Unit mass 

(kg) 

CBE mass 

(kg) 

Growth 

(%) 

Growth 

(kg) 

Total mass 

(kg) 

01 Robotic Lunar Lander - --- 429 7 33.3 462

01.01 LOI Stage - --- 25 15 4.5 29.5 

01.02 Lunar Descent Stage - --- 266 0 1.1 266.8 

01.03 Lunar Lander Stage - --- 138 16.7 27.7 165.7 

01.03.01 Avionics, Instrumentation, GN&C - --- 35 16 6.8 42.0 

01.03.01.01 C&DH - --- 8.3 16 1.6 9.9 

01.03.01.02 Communication - --- 6.9 11 0.9 7.7 

01.03.01.02.01 S-Band Comm System - --- 6.9 11 0.9 7.7 

01.03.01.02.02 Transponder 1 3.4 3.4 3 0.1 3.5 

01.03.01.02.03 RF Assembly 0 5.9 0.0 20 0.0 0.0 

01.03.01.02.04 Processing Module 0 0.0 0.0 30 0.0 0.0 

01.03.01.02.05 Antenna 4 0.8 3.0 20 0.6 3.6 

01.03.01.02.06 Coaxial Cable 1 0.5 0.5 30 0.2 0.7 

01.03.01.02.07 Installation—Mounting & Circuitry 0 1.9 0.0 30 0.0 0.0 

01.03.01.03 GN&C - --- 20.1 18 4.4 24.4 

4.1.4 Communications Trades

None.

4.1.5 Communications Analytical Methods

None.

4.1.6 Communications Risk Inputs

None.

4.1.7 Communications Recommendation

None.

4.2 Avionics

For the purpose of this design and report, Avionics consisted of the C&DH system as well as the 

GN&C system.

4.2.1 Avionics Requirements

The avionics system needs to provide all processing support for the Earth-Moon transit, landing and 

surface operations. Storage requirements were estimated to be minimal since direct communications will 

be available during landed science operations. The processing however will have to be quick enough to 

handle landing control. However, in the interest of cost, a smaller computer will be sought with an 

estimated MIPS performance at 1 MIPS or less and memory in the range of 8 to 16 MB.

4.2.2 Avionics Assumptions

4.2.2.1 Guidance, Navigation and Control (GN&C) Assumptions

The goal of the GN&C subsystem design was to provide for a successful lunar landing at minimum 

cost, and design a system that may be zero fault tolerant.

Assumptions

 Successful lunar landing requires attitude control

− For solar panel

− For propulsion
NASA/TM—2010-216218 17

− For communications

 Successful lunar landing requires knowledge of position

− For propulsion

4.2.3 Avionics Design and MEL

The GN&C system consisted of

 Hardware suite based on Surveyor

− Surveyor was a minimalist design

− Technological advances may make our implementation more than minimalist

 Catalog items

− Sun sensor: Sodern Digital Sun Sensor (http://www.sodern.eu/site/FO/scripts/index.php)

• French

− Star tracker: Sodern SED 16

• French

− Inertial measurement unit: Honeywell HG9900 

• (http://www.honeywell.com/sites/aero/Military-Aircraft3_CF9E0EA52-E7BD-F572-

03CA-068C338328C1_HB0A141BB-2FD0-7339-851E-22466484B4C6.htm)

 Augmented catalog item

− Radar altimeter: Honeywell HG9550 augmented

• http://www.honeywell.com/sites/aero/Military-Aircraft3_CC97F1FA1-ED1D-D7D4-

15D8-1272F2E6C702_HFC5482F9-3DD3-2A4F-DE15-0A1BCB4CA9C6.htm

Figure 4.2 shows the schematic design of the avionics system—this is primarily the C&DH system 

but also includes the linkage to the GN&C system.

Figure 4.2—Lunar Lander avionics system.
NASA/TM—2010-216218 18

Table 4.2 is the combined avionics MEL. The Growth percentage is only important on the line items. 

Refer to the growth (kg) totals for details. The total growth percentage of the wet mass calculation need to 

be refined in future versions of the MEL tool.

TABLE 4.2—AVIONICS SYSTEM MEL

WBS Description

Qty Unit 

mass 

(kg) 

CBE 

mass 

(kg) 

Growth 

(%) 

Growth 

(kg) 

Total 

mass 

(kg) 

01 Robotic Lunar Lander - ---- 429 7 33.3 462

01.01 LOI Stage - ---- 25 15 4.5 29.5 

01.02 Lunar Descent Stage - ---- 266 0 1.1 266.8 

01.03 Lunar Lander Stage - ---- 138 16.7 27.7 165.7 

01.03.01 Avionics, Instrumentation, GN&C - ---- 35 16 6.8 42.0 

01.03.01.01 C&DH - ---- 8.3 16 1.6 9.9 

01.03.01.01.01 General Avionics Processor 1 5.3 5.3 15 0.8 6.0 

01.03.01.01.02 Time Generation Unit 0 0.3 0.0 30 0.0 0.0 

01.03.01.01.03 Command and Control Harness (data) 1 1.2 1.2 25 0.3 1.5 

01.03.01.01.04 Instrumentation & Wiring 1 1.9 1.9 25 0.5 2.3 

01.03.01.02 Communication - ---- 6.9 11 0.9 7.7 

01.03.01.03 GN&C - ---- 20.1 18 4.4 24.4 

01.03.01.03.01 Inertial Measurement Units 1 3.0 3.0 25 0.8 3.8 

01.03.01.03.02 Star Sensor 1 3.0 3.0 25 0.8 3.8 

01.03.01.03.03 Sun Sensors 1 0.3 0.3 25 0.1 0.3 

01.03.01.03.04 Star Tracker 0 0.0 0.0 30 0.0 0.0 

01.03.01.03.05 Attitude control system 0 0.0 0.0 30 0.0 0.0 

01.03.01.03.06 Flight Control Electronics 2 2.5 5.0 3 0.2 5.2 

01.03.01.03.07 Radar System 1 8.8 8.8 30 2.6 11.4 

01.03.01.03.08 Visualization System 0 12.0 0.0 30 0.0 0.0 

4.2.4 Avionics Trades

None.

4.2.5 Avionics Analytical Methods

The avionics in this design were made on a modern serial bus design. The reasons from the designers 

are as follows:

 Crew Exploration Vehicle (CEV) is using it: Development (costs) is addressed by the Human 

Exploration Program

 Modular and light weight connections: Reduces vehicle harness weight and provides an compact 

interface, increased flexibility to add or remove systems.

 Widespread availability: Low cost commercial off the shelf (COTS) microcontrollers are 

available with built-in 1394B (or USB) interfaces.

 Single string systems do not require the expensive fly by wire 1553B bus

 Embedded microcontrollers are small, compact, power efficient and include instrumentation 

Interfaces and Network Connections. 

 CEV is expected to employ a “hardened” version 1394B 

 LXI LAN Extensions for Instrumentation—(IEEE 1558) typically applied to Ethernet also applies 

to virtually any high-speed serial bus that will support the protocols for time synchronization, 

Network Control and Hardware Triggering.

 Real Time or Deterministic operations are achieved by exploiting bus high speed.

 Execution of controllers accurate to tens of nanoseconds.
NASA/TM—2010-216218 19

4.2.6 Avionics Risk Inputs

None.

4.2.7 Avionics Recommendation

None.

4.3 Electrical Power System

4.3.1 Electrical Power Requirements

Provide power to spacecraft and payloads. Distribute and control power to each electrical load on 

spacecraft.

4.3.2 Electrical Power Assumptions

 The spacecraft will only operate on the surface of the Moon during sunlit periods.

 Assumed 100 W housekeeping power (70 W with 40% growth) and 25 W payload power for 

nominal sizing (if housekeeping loads are lower than 100 W, then excess power may be used for 

payload)

 Peak power loads and periods from available hardware used to confirm energy storage size. 

Driving energy storage factor is the pad-fairing jettison period.

 Solar array sizing based on efficiencies/knockdown factors/solar array lay-down

 Battery sizing based on specific energy metric derived from accepted average values.

 Electronics and electrical based on specific power metric derived from historical data.

 Solar array boom/hinge sizing based on fraction of supported mass.

 Gimbals sizing based on off the shelf (OTS) the shelf mass (Moog).

 Support functions supplied by the other subsystems

− GN&C provided pointing knowledge and orients the spacecraft in orbit/transit to maximize 

solar array visibility to the Sun (within 2°). 

− Structures provides housing boxes for electronics and batteries.

− Thermal regulates the battery and electronics temperature to a safe operating range.

4.3.3 Electrical Power Design and MEL

Components

 Solar array: 3.6 kg (with 25% growth = 4.5 kg), 0.55 m2 (0.74 by 0.74 m)

 Solar array gimbal/electronics: 2.8 kg (with 18% growth = 3.3 kg), 8 W/axis

 Solar array hinge/boom: 4.8 kg (with 18% growth = 5.7 kg)

 Battery assembly: 2.4 kg (with 20% growth = 2.9 kg)

 Power management/control electronics: 6.8 kg (with 25% growth = 8.5 kg)

 Power distribution harness and sensor/control wiring: 4.0 kg (with 50% growth = 6.0 kg)

Table 4.3 is the power system MEL. Note that there is no power system on the LOI stage (the payload 

adaptor) or the Lunar Descent stage (the Star 27 motor), only on the Lander stage. This was the power to 

the landing propulsion system and for use on the surface. The Growth percentage is only important on the 

line items. Refer to the growth (kg) totals for details. The total growth percentage of the wet mass 

calculation need to be refined in future versions of the MEL tool.
NASA/TM—2010-216218 20

TABLE 4.3—ELECTRICAL POWER SYSTEM MEL

WBS Description

Qty Unit 

mass 

(kg) 

CBE 

mass 

(kg) 

Growth 

(%) 

Growth 

(kg) 

Total 

mass 

(kg)

01 Robotic Lunar Lander - --- 429 7 33.3 462

01.01 LOI Stage - --- 25 15 4.5 29.5 

01.02 Lunar Descent Stage - --- 266 0 1.1 266.8 

01.03 Lunar Lander Stage - --- 138 16.7 27.7 165.7 

01.03.09 Power - --- 24 21 6.5 30.9 

01.03.09.01 Battery System - --- 2 17 0.5 2.9 

01.03.09.01.01 Battery Assembly-Primary 0 0.0 0 20 0.0 0.0 

01.03.09.01.02 Battery Assembly-Secondary 1 2.4 2 20 0.5 2.9 

01.03.09.02 Solar Array - --- 11 17 2.3 13.5 

01.03.09.02.01 Solar Array Hinge/Boom 1 4.8 5 18 0.9 5.7 

01.03.09.02.02 Solar Array Drive Assemblies/Motors 1 2.8 3 18 0.5 3.3 

01.03.09.02.03 Solar Array Mass 1 3.6 4 25 0.9 4.5 

01.03.09.03 Power Management & Distribution - --- 10.8 26 3.7 14.5 

01.03.09.03.01 Power management/control electronics 1 6.8 6.8 25 1.7 8.5 

01.03.09.03.02 Power distribution/monitoring wiring harness 1 4.0 4.0 50 2.0 6.0 

4.3.4 Electrical Power Trades

Options considered

 Flight proven, standard practice components.

 Triple junction solar cells on a rigid panel were assumed because this was the best trade of mass 

and cost.

 OTS lithium ion batteries used for low cycle geostationary communication satellites assumed for 

high energy density (relatively low number of cycles needed). Primary (nonrechargeable) lithium 

batteries were considered initially due to high “pad-fairing jettison” energy requirements, but 

when this time period was reduced, rechargeable batteries were optimum.

 OTS dual axis solar array drive and electronics.

 Other electronics are fairly standard but some modifications are needed.

4.3.5 Electrical Power Analytical Methods

None.

4.3.6 Electrical Power Risk Inputs

None.

4.3.7 Electrical Power Recommendation

 Lessons learned

− Reduce the energy requirements on the pad as much as possible since that drives the battery 

size.

 Issues for further study

− Need to assess solar array structures to handle various launch/thruster loads.

− Peak loads (e.g., communications and radar) must be adjusted to have realistic (<15 min/hr) 

durations
NASA/TM—2010-216218 21

4.4 Structures and Mechanisms

4.4.1 Structures and Mechanisms Requirements

The Robotic Lunar Lander and Star 27 Solid Rocket motor must fit inside the Minotaur V Launch 

Vehicle 92-in. payload fairing (see Figure 4.3). Mechanisms are required which separate the various 

portions of the spacecraft just after launch and after Star 27 rocket burnout. An additional mechanism is 

required to point the solar array. Payload specific mechanisms are not considered here.

4.4.2 Structures and Mechanisms Assumptions

Spacecraft must survive the following assumed loads

 Launch axial (12gs), no lateral loads used in modeling, from the Minotaur V

 Landing axial (5gs), no lateral loads used in modeling, from the Star 27 motor firing

Figure 4.3 shows the payload fairing of the Minotaur V Launch Vehicle used as the dimensions for 

limits in the payload packaging of the spacecraft during the design. This is the assumption used for the 

dimensions into which the structure and stack must fit. Note that the study used metric units in its design. 

However, some of the graphics from third parties use English units.

4.4.3 Structures and Mechanisms Design and MEL

Figure 4.4 shows the conceptual Robotic Lunar Lander as packaged inside the Minotaur 92-in. 

payload fairing. Note that the legs are deployed while in the fairing and that there is plenty of volume 

room for the spacecraft, antenna, legs, and solar arrays.

Figure 4.3—Minotaur V Launch Vehicle 92-in. 

payload fairing.

Figure 4.4—Robotic Lunar Lander with Star 27 inside 

Minotaur V Launch Vehicle 92-in. payload fairing.
NASA/TM—2010-216218 22

Table 4.4 lists the Structures and Mechanical System MEL for the three stages making up the Robotic 

Lunar Lander. The Growth percentage is only important on the line items. Refer to the growth (kg) totals 

for details. The total growth percentage of the wet mass calculation need to be refined in future versions 

of the MEL tool. The total growth percentage using the wet mass number in the percentage calculation 

gives a deceptive percentage number. Typically growth percentage is calculated growth divided by total 

CBE mass.

TABLE 4.4—STRUCTURE AND MECHANICAL SYSTEM MEL

WBS Description

Qty Unit 

mass 

(kg)

CBE 

mass 

(kg)

Growth 

(%)

Growth 

(kg)

Total 

mass

(kg)

01 Robotic Lunar Lander - ---- 429 7 33.3 462

01.01 LOI Stage - ---- 25 15 4.5 29.5

01.01.11 Structures & Mechanical Systems - ---- 25 15 4.5 29.5

01.01.11.01 Primary Structures - ---- 25 15 4.5 29.5

01.01.11.01.01 Landing Gear (includes mechanisms) 0 5.0 0.0 30 0.0 0.0

01.01.11.01.02 EDS adaptor interface 1 25.0 25.0 18 4.5 29.5

01.01.11.01.03 Main Structure/Fuselage 0 25.0 0.0 30 0.0 0.0

01.01.11.02 Installation - ---- 0.0 0 0.0 0.0

01.01.11.03 Mechanisms - ---- 0.0 0 0.0 0.0

01.02 Lunar Descent Stage - ---- 266 0 1.1 266.8

01.03 Lunar Lander Stage - ---- 138 16.7 27.7 165.7

01.03.11 Structures & Mechanical Systems - ---- 15 15 2.8 18.2

01.03.11.01 Primary Structures - ---- 6 15 1.1 7.3

01.03.11.01.01 Landing Gear (includes mechanisms) 3 0.4 1.2 18 0.2 1.4

01.03.11.01.02 EDS adaptor interface 0 0.0 0.0 30 0.0 0.0

01.03.11.01.03 Main Structure/Fuselage 1 5.0 5.0 18 0.9 5.9

01.03.11.02 Installation - ---- 7.7 15 1.4 9.1

01.03.11.02.01 C&DH Installation 1 0.7 0.7 18 0.1 0.8

01.03.11.02.02 Communications Installation 1 0.5 0.5 18 0.1 0.6

01.03.11.02.03 GN&C Installation 1 1.7 1.7 18 0.3 2.0

01.03.11.02.04 Power Installation 1 2.2 2.2 18 0.4 2.6

01.03.11.02.05 Propulsion Installation 1 0.4 0.4 18 0.1 0.5

01.03.11.02.06 Propellant Storage Installation 1 0.7 0.7 18 0.1 0.8

01.03.11.02.07 Thermal Installation 1 1.5 1.5 18 0.3 1.8

01.03.11.03 Mechanisms - ---- 1.5 15 0.3 1.8

01.03.11.03.01 Solar array deployment mechanism 1 0.0 0.0 30 0.0 0.0

01.03.11.03.02 Radiator deployment mechanism (if applicable) 0 0.0 0.0 30 0.0 0.0

01.03.11.03.03 Separation mechanism (pyros) 1 1.5 1.5 18 0.3 1.8

4.4.4 Structures and Mechanisms Trades

Only metallic (aluminum-lithium, AL 2090) structure was considered in this design. No other trades 

on structural material were conducted.

4.4.5 Structures and Mechanisms Analytical Methods

The effect of the weight of the rest of the subsystems was considered in the structure.

Level and Approach of Analyses

 Geometry

− Hexagonal bus with three legs. See Figure 4.6, Figure 4.7, and Figure 4.8.

− Support structure made of 1.5 in.2 and 1.5-in.-diameter round tubes.

 Material

− Aluminum lithium—Al 2090

 Loads:

− 12-g acceleration in the axial direction (from Minotaur I & IV Users Guide). 
NASA/TM—2010-216218 23

− Mass effect of other subsystems were added

 Mass calculation

− Run finite element analysis of current design for linear static failure and buckling. If positive 

margins were obtained use model mass, otherwise modify and repeat. Finite element analysis 

was done only for the main truss structure, the rest was added as a lumped mass.

− Add other subsystem installation mass (10% of subsystem mass)

− Separation mechanisms in MEL were scaled from Surveyor spacecraft databook.

Figure 4.5 is output from the finite element model (FEM) structural analysis program and represents 

the solid model of the Case 1 (baseline) chemical lander used in the FEM analysis.

For the FEM analysis, the following structural assumptions were used

 Maximum displacement = 1.1 in.

 Maximum stress = 3520 psi

 Minimum stress = 4500 psi

 Minimum margin of safety = 7.4 (ultimate allowable 38000 psi)

Figure 4.6, Figure 4.7, and Figure 4.8 are the results of the FEM analysis using the above assumptions 

on displacement and stress.

Figure 4.5—Case 1—Finite Element Model of Robotic Lunar Lander stack.
NASA/TM—2010-216218 24

Figure 4.6—Case 1 minimum combined stress.

Figure 4.7—Case 1 maximum combined stress.
NASA/TM—2010-216218 25

Figure 4.8—Case 1 displacement.

4.4.6 Structures and Mechanisms Risk Inputs

The risk identified for the structures was the potential impact with foreign object or due to nearby 

operations.

4.4.7 Structures and Mechanisms Recommendation

Mass optimization needs to be done in the structure such that sizing is done for a minimum ultimate 

safety factor of 2 and a minimum buckling load factor of 1.4. In addition, it will be necessary to perform 

frequency/normal modes analysis especially on the rigid solar array panel.

4.5 Propulsion and Propellant Management

4.5.1 Propulsion and Propellant Management Requirements

The spacecraft propulsion subsystem was required for up to three propulsion operations

1. Orbit insertion

2. Descent to lunar surface

3. Surface landing

Depending on the study case, these operations were performed by different propulsion systems. 

Table 2.3 shows the performance requirements for the chemical and electric propulsion options. 

4.5.2 Propulsion and Propellant Management Assumptions

The propulsion subsystems designed for these studies used only COTS components to mitigate 

development costs and time requirements. All solid rocket motors and chemical thrusters used in the 

design as well as the propellant management components and propellant tanks were from operating 

manufactures. The electric thrusters were assumed to come from known vendors or obtained directly from 
NASA/TM—2010-216218 26

the NASA development teams. Electric thruster performance is based on demonstrated operation and does 

not require any new qualification testing.

4.5.3 Propulsion and Propellant Management Analytical Methods

Because the propulsion subsystems were assembled from existing components, the analysis required 

consisted primarily of maintaining a mass roll up for the various subassemblies. The first two propulsion 

operations were performed with solid rocket motors that were selected from the Star motor catalog. The 

selection was based primarily on total spacecraft mass and the propellant load fraction of the motor.

The primary analysis that was actively performed was to determine the propellant tank sizes based on 

propellant conditions over the mission duration. The tank requirements were determined using propellant 

density and storage pressure through Hoop Stress Analysis. These requirements were then used to select 

the best match from the PSI and Arde Inc. storage tank catalogs.

Once the storage tank(s) were selected, the helium pressurization requirements were determined. A 

conventional He pressurization system configuration was used, based on our experience with previous 

lander and Orion Service Module studies.

4.5.3.1 LOI and Descent Propulsion System Modeling

For the study of the all-chemical propulsion case, the orbit insertion and descent burns were 

performed with solid rocket motors:

 Solid rocket motors were selected from existing catalog of flight-proven options to minimize 

development requirements. The motors were selected based on:

− Thrust level and required propellant load, that were provided by Mission Analyses

− Best match for solid motor performance was selected and motor weight was adjusted to 

account for off-loaded propellant (as required)

 Use adjusted propellant mass fraction to estimate propulsion subsystem final mass for spacecraft 

mass determination in order to perform mission analyses

 Known configuration of existing motor used by Structures and computer aided design (CAD) 

teams to establish spacecraft configuration

Figure 4.9 shows a pictorial illustration of the issues associated with choosing the appropriate solid 

rocket motor for a mission. The trade between the Star 27 and Star 24 motors is noted at the bottom of the 

figure. The Star 27 motor was chosen for the Case 1 (baseline) design. (Star 27 online data sheet 

http://www.atk.com/customer_solutions_missionsystems/documents/atk_catalog_may_2008.pdf.)

The summary of the propulsion technology choices made for the two cases that were investigated in 

this study are listed in Table 4.5 for the All-Chemical option and Table 4.6 for the Electric Propulsion 

option. 

Figure 4.9—Solid motor sizing dependencies.
NASA/TM—2010-216218 27

TABLE 4.5—ALL-CHEMICAL PROPULSION SUBSYSTEM BREAKDOWN

Stage Engine type Thrust level, 

N

Specific impulse, 

sec

Size, 

kg

Engine 

ID

Orbit insertion 

stage

Main: Solid rocket motor ---- ---- ---- ----

RCS: cold gas jets ---- ---- ---- ----

Descent stage Main: Solid rocket motor 26996 287.8 265.5 STAR 27

RCS: cold gas jets ---- ---- ---- ----

Lander stage Main: Liquid monoprop engine 220 222 0.63 MR–107K

RCS: cold gas jets 2 229 0.35 MR–111

TABLE 4.6—ELECTRIC PROPULSION OPTION ORBIT INSERTION SUBSYSTEM BREAKDOWN

Element Type Number Mass,

kg

Size,

m

Operating 

power, 

W

Power duty 

cycle/

duration

Engines HiVHAC Hall Effect thruster 2 3.6 ---- 1350 50%

Thruster gimbal 2 4.6 TBD ------ ----

Thruster gimbal drive 2 2 TBD TBD TBD

Power management Power processing unit 2 6 TBD 80 50%

Propellant storage Xe high pressure storage tank—ATK–PS–

80412–1

2 7 0.33 diam by 

0.70

------ ----

Propellant 

management

Feed system, fixed 1 4 TBD TBD TBD

Feed system, per thruster 2 1 TBD TBD TBD

For the study of the electric propulsion case, the orbit insertion and descent burns were performed 

with Hall Effect thrusters:

 The High Voltage Hall ACcelerator (HiVHAC) Hall Thrusters are under active development at 

GRC through the In Space Propulsion program. These were selected based on:

− Optimum specific impulse capability allowed within available power budget

− High efficiency operation of this thruster enabled this option to be competitive with the solid 

motor option while enabling increasing mission flexibility

 Xe propellant stored supercritically in high-pressure and light-weight Carbon-Overwrapped 

Pressure Vessels (COPV) 

The liquid propellant storage tank was sized according to propellant amount and storage conditions 

(temperature and pressure):

 Used diaphragm tanks for blow down operation

 Selected tank from ATK-PSI catalog with necessary capacity after determining size based on 

propellant volume

The approach for determining the size and mass of the xenon propellant storage tank for the electric 

propulsion option case was similar however the xenon is stored at substantially higher pressures than the 

hydrazine.

Propellant management system (PMS) mass model derived from Aerojet’s Hydrazine Auxiliary 

Propellant System (HAPS) configuration, illustrated in Figure 4.10.

 Used controlled database for component masses (valves, sensors, filters, etc)

 Assumed fixed feed line lengths

The PMS masses were composed of the component masses that were relatively static for each 

propellant option. Thermal control via insulation in addition to line and tank heaters was assumed to be 

included in the PMS mass estimate. The thermal requirements for Xe storage were significantly less than 

those of the hydrazine and were limited to insulation overwrap on the storage tank.

Direct trajectory
NASA/TM—2010-216218 28

Figure 4.10—Aerojet’s HAPS propulsion system schematic.

The Xe PMS was similarly modeled after historical flight systems and likewise was represented by a 

mass roll up of the individual components. Additionally, the power processing unit developed for the 

HiVHAC hall thruster was used for this model.

4.5.3.2 Lander Propulsion System Modeling

The propulsion subsystem on the Lander portion of the spacecraft performed two functions: first, 

controlled (“soft”) landing which had traditionally required variable thrust engines, and second, reaction 

control propulsion for vehicle attitude maintenance and specialized maneuvers. The reaction control 

system (RCS) on the lander vehicle provided reaction and attitude control for the entire spacecraft over 

the mission.

For the main propulsion system (MPS), a liquid monopropellant engine was selected based on the 

performance requirements determined by trajectory analysis discussed earlier in this report that could be 

evolved to obtain the necessary throttling capability. 

 Three MR-107K monoprop rockets were baselined for landing

− Operating at: 257 N (57 lbf) thrust; 236 sec specific impulse

− Operated with Hydrazine to reduce propellant management system complexity

For reaction/attitude control, small monoprop rockets were selected. This propulsion subsystem was 

derived from similar historical systems and was a conventional configuration:

 Three Aerojet MR-111E rockets were baselined

− Operating at: 4.4 N (0.99 lbf) thrust, 229 sec specific impulse

− Used Hydrazine propellant for reduced propellant management system complexity; common 

propellant for all-chemical option

The study scope was limited to Delta-V burn matching and did not require detailed analysis of 

reaction and attitude control performance. Hence, a simple mass model for the RCS was sufficient. All 

performance information for the rockets used in the propulsion subsystem model were obtained from 

either Aerojet product data sheets or from entries on the Astronautix website. 
NASA/TM—2010-216218 29

TABLE 4.7—LANDER PROPULSION SUBSYSTEM BREAKDOWN

Element Type Number Mass, 

kg

Size, 

m

Operating 

power, 

W

Power duty 

cycle/ 

duration

Engines Main: MR-107 MMH monoprop 3 0.63 0.66 diam by 2.2 60 1 min. 

RCS: MR-111 N2 Cold gas jets 6 0.35 0.36 diam by 0.17 12 5% 

Main Engine Gimbal 1 1 TBD 10 1 min. 

Propellant 

storage 

MMH (hydrazine) storage tank 

ATK-PSI 80259-1 

2 6 0.48 Dia. 10 0.01 min. 

Valves, Service 3 0.28 TBD TBD 7% 

Valve, Pyro 1 0.15 TBD TBD >1% 

Propellant Filter 7 0.05 TBD Š Š

Management 

Transducer 2 0.1 TBD TBD TBD

Valve, RCS Control 6 0.1 TBD TBD TBD

Feed Lines 9 0.1 0.01 Š Š

The liquid propellant storage tank was sized according to propellant amount and storage conditions 

(temperature and pressure):

 Used diaphragm tanks for blow down operation

 Selected tank from ATK-PSI catalog with necessary capacity after determining size based on 

propellant volume

PMS mass model derived from Aerojet’s Hydrazine Auxiliary Propellant System (HAPS) 

configuration

 Used controlled database for component masses (valves, sensors, filters, etc.)

 Assumed fixed feed line lengths

The PMS masses were composed of the component masses that were relatively static for each 

propellant option. Thermal control via insulation in addition to line and tank heaters was assumed to be 

included in the PMS mass estimate. 

4.5.4 Propulsion and Propellant Management Design and MEL

Table 4.8 lists the line items in the Propulsion and Propellant Management MEL including the 

propellant. The MEL was created in a generic fashion for the all-chemical propulsion systems, but did 

include line items for the EP subsystems. Those systems are hidden in the table view in Table 4.6. The 

growth percentage (growth mass/total mass) is only tracked and calculated on the line items. Refer to the 

growth (kg) totals for details. The total growth percentage of the wet mass calculations needs to be 

removed to prevent confusion in future versions of the MEL tool. The total growth percentage using the 

wet mass number in the percentage calculation gives a deceptive percentage number. Typically growth 

percentage is calculated growth divided by total current best estimate (CBE) mass.
NASA/TM—2010-216218 30

TABLE 4.8—PROPULSION AND PROPELLANT MANAGEMENT MEL

WBS Description Qty Unit

mass

(kg)

CBE

mass

(kg)

Growth

(%)

Growth

(kg)

Total

mass

(kg)

01 Robotic Lunar Lander - ---- 429 7 33.3 462

01.01 Lunar Orbit Insertion Stage - ---- 25 15 4.5 29.5 

01.02 Lunar Descent Stage - ---- 266 0 1.1 266.8 

01.02.02 Propulsion (Chemical) - ---- 27 4 1.1 28.5 

01.02.02.01 Main Engine - ---- 27.5 4 1.1 28.5 

01.02.02.01.01 Main Engine 1 27.5 27.5 4 1.1 28.5 

01.02.02.01.02 Main Engine Gimbal 0 0.0 0.0 4 0.0 0.0 

01.02.02.02 Reaction Control System - ---- 0.0 0 0.0 0.0 

01.02.02.02.01 RCS Engine 0 0.0 0.0 4 0.0 0.0 

01.02.03 Propellant Management (Chemical) - ---- 0 0 0.0 0.0 

01.02.04 Propellant Hardware (Chemical) - ---- 238 0 0.0 238.3 

01.02.04.01 Main Engine Propellant - ---- 238 0 0.0 238.3 

01.02.04.01.01 Fuel - ---- 238.3 0 0.0 238.3 

01.02.04.01.02 Oxidizer - ---- 0.0 0 0.0 0.0 

01.02.04.01.03 Pressurant 0 0.0 0.0 0 0.0 0.0 

01.02.04.02 RCS Propellant - ---- 0.0 0 0.0 0.0 

01.03 Lunar Lander Stage - ---- 138 16.7 27.7 165.7 

01.03.02 Propulsion Hardware (Chemical) - ---- 5 5 0.3 5.6 

01.03.02.01 Main Engine - ---- 3.3 4 0.1 3.5 

01.03.02.01.01 Main Engine 3 0.8 2.3 4 0.1 2.4 

01.03.02.01.02 Main Engine Gimbal 1 1.0 1.0 4 0.0 1.0 

01.03.02.02 Reaction Control System - ---- 2.0 7 0.2 2.1 

01.03.02.02.01 RCS Engine 6 0.3 2.0 8 0.2 2.1 

01.03.03 Propellant Management (Chemical) - ---- 9 6 0.5 9.6 

01.03.03.01 OMS Propellant Management - ---- 8 5 0.4 8.3 

01.03.03.01.01 Fuel Tanks 1 6.0 6.0 4 0.2 6.3 

01.03.03.01.02 Fuel Lines 1 0.3 0.3 18 0.1 0.3 

01.03.03.01.03 Oxidizer Tanks 0 0.0 0.0 0 0.0 0.0 

01.03.03.01.04 Oxidizer Lines 0 0.0 0.0 0 0.0 0.0 

01.03.03.01.05 Pressurization System - tanks, panels, lines 0 0.0 0.0 0 0.0 0.0 

01.03.03.01.06 Feed System - regulators, valves, etc 1 1.6 1.6 8 0.1 1.7 

01.03.03.02 RCS Propellant Management - ---- 1.2 9 0.1 1.3 

01.03.03.02.01 Fuel Tanks 0 0.0 0.0 0 0.0 0.0 

01.03.03.02.02 Fuel Lines 1 0.3 0.3 18 0.1 0.3 

01.03.03.02.03 Pressurization System - tanks, panels, lines 1 0.0 0.0 0 0.0 0.0 

01.03.03.02.04 Feed System - regulators, valves, etc 1 0.9 0.9 8 0.1 1.0 

01.03.04 Propellant (Chemical) - ---- 30 20 7.6 37.8 

01.03.04.01 Main Engine Propellant - ---- 26 20 6.5 32.3 

01.03.04.01.01 Fuel - ---- 25.8 20 6.5 32.3 

01.03.04.01.01.01 Fuel Usable 1 25.8 25.8 25 6.5 32.3 

01.03.04.01.01.02 Fuel Boiloff 0 0.0 0.0 0 0.0 0.0 

01.03.04.01.01.03 Fuel Residuals (Unused) 1 0.0 0.0 0 0.0 0.0 

01.03.04.01.02 Oxidizer - ---- 0.0 0 0.0 0.0 

01.03.04.01.03 Pressurant 0 0.0 0.0 0 0.0 0.0 

01.03.04.02 RCS Propellant - ---- 4.4 20 1.1 5.5 

01.03.04.02.01 Fuel - ---- 4.4 20 1.1 5.5 

01.03.04.02.01.01 Fuel Usable 1 4.4 4.4 25 1.1 5.5 

01.03.04.02.01.02 Fuel Boiloff 0 0.0 0.0 0 0.0 0.0 

01.03.04.02.01.03 Fuel Residuals (Unused) 1 0.0 0.0 0 0.0 0.0 

01.03.04.02.02 Oxidizer - ---- 0.0 0 0.0 0.0 

01.03.04.02.03 Pressurant 0 0.0 0.0 0 0.0 0.0 

4.5.5 Propulsion and Propellant Management Trades

Three design trades for the propulsion subsystem were performed over the course of this study.
NASA/TM—2010-216218 31

 Trade 1: Solid Rocket Motor Selection

− Motor choices limited to selecting OTS motors

• Advanced technology readiness and demonstrated heritage of these systems were 

required to ensure relatively low cost as well as availability

• Mono- and bipropellant liquid rocket engine alternatives were examined but rejected 

because:

♦ Propulsion subsystems heavier, much more complicated

♦ Technology availability in the near-term

 Trade 2: Lander Rocket Engine Selection

− Hydrazine-based engines were selected because

• Technology mature and has extensive demonstrated heritage for similar applications

• Engine class considered OTS (i.e., HAPS derivative)

• Propellant management system less complex

− Liquid propellant engine necessary for throttling performance required for landing on lunar 

surface

• Cost of modification needs to be done

 Trade 3: Use of Electric Propulsion Insertion Stage (where applicable)

− Electric thruster selection driven by lifetime requirements

• HiVHAC thruster capability > 150 kg Xe throughput

• SP-100 Hall thruster also considered because of match with performance requirements

♦ Limited benefits because lifetime capability ~80 kg throughput so multiple thrusters 

will be needed

4.5.6 Propulsion and Propellant Management Risk Inputs

The only risk identified for this study was the availability of the electric propulsion technology and its 

associated costs.

4.5.7 Propulsion and Propellant Management Recommendation

The chemical propulsion system option for trans-lunar injection was chosen. Electric propulsion will 

add capability but at the cost of a baseline vehicle mass that is at the limit of the Minotaur class launch 

vehicle capability. Additionally, the primary benefits of the electric propulsion technology are only 

realized when hardware reuse is required to enable or enhance mission return. One example of such a 

benefit is that the EP stage could remain in orbit as communications relay satellite, thereby extending the 

operational capability of the landing vehicle.

4.6 Thermal Control

4.6.1 Thermal Requirements

The thermal system must preserve acceptable temperature ranges on all systems from pad to launch to 

transit to landing to end of mission on the lunar surface. Each of these environments drives different parts 

of the thermal design, however the lunar surface environment was chosen as the driving case.

4.6.2 Thermal Assumptions

 Radiator panels are fixed (nondeployable) and located on the ascent vehicle.

 Radiators: Radiate heat to space. Low absorptive (a) and high emissivity (e) coatings are 

desirable to minimize solar input and maximize heat rejection to space.

 Various absorptivity (a) and emissivity (e) coatings are desirable for different locations within 

Lander.
NASA/TM—2010-216218 32

 Heat Sinks: Placed in thermal contact with high dissipation components. Combined with radiators to 

dissipate the heat, which they conduct. Could use the structure of the lander itself as a heat sink.

 Heaters: Provide heating for temperature sensitive components (fuel, battery, etc.)

 MLI: blankets with low absorptivity and high emissivity. Reduce the heat flow rate of the system 

while preventing large heat flux. They can be used to wrap around sensors and payloads for 

thermal insulation and to reduce thermal requirements.

 Temperature Sensors: Resistive type thermal sensors. Monitor temperature of sensitive 

components. (Fuel, fuel line, batteries, etc.)

4.6.3 Thermal Design and MEL

The thermal design, shown in the Figure 4.11 schematic, for dissipating heat in the following major 

subsystem areas consisted of the following assumptions and design details: 

Electronics:

 15 W dissipation

 270 K lunar sink

 40 °C mounting plate maximum temperature

 Requires 0.085 m2 radiator (45- by 19-cm) on the top cover. Box cover is 30- by 10-cm.

 Painted with Z-93 (e = 0.9, a = 0.2)

Figure 4.11—Thermal control schematic.
NASA/TM—2010-216218 33

Figure 4.12—Thermal control for electronics.

Figure 4.13—Thermal control for batteries.

Figure 4.12 shows a notional diagram of the thermal system designed to handle the heat of the 

electronics.

Batteries:

 3 W dissipation

 270 K lunar sink

 20 °C mounting plate maximum temperature

 Requires 0.035 m2 radiator (15- by 24-cm) on the top cover. Box cover is 15- by 15-cm.

 Painted with Z-93 (e = 0.9, a = 0.2)

 Heaters and MLI

Figure 4.13 shows a notional diagram of the thermal system designed to handle the heat of the 

batteries. Avionics (total of three boxes):

 270 K lunar sink

 40 °C mounting plate maximum temperature

 Radiators painted with Z-93 (e = 0.9, a = 0.2)

 RF assembly

− 0.95 W dissipated

− Requires 0.005 m2 radiator. 10- by 8-cm top cover is sufficient.

 Transponder

− 0.95 W dissipated

− Requires 0.005 m2 radiator. 20- by 16-cm top cover is sufficient.

 General avionics

− 0.95 W dissipated

− Requires 0.005 m2 radiator. 13.35- by 11.4-cm top cover is sufficient.
NASA/TM—2010-216218 34

Figure 4.14—Thermal control for avionics.

Figure 4.15—Thermal control for propulsion system.

Figure 4.14 shows a notional diagram of the thermal system designed to handle the heat of the 

avionics. Solid rocket motor, hydrazine fuel tank, and hydrazine fuel lines:

 Allowable temperature range: 7 to 35 °C

 270 K lunar sink

 Heaters and MLI

 Hydrazine fuel: ~45 kg, tank is 53.3 cm (21 in.) diameter made of titanium.

 Solid Rocket: STAR 27, 238 kg total mass, roughly a 61-cm-diameter ball made of titanium.

Figure 4.15 shows a notional diagram of the thermal system designed to handle the heat of the 

propulsion system.

Table 4.9 lists the line items in the thermal control system MEL. These numbers are sent to the main 

system MEL for summary.
NASA/TM—2010-216218 35

TABLE 4.9—THERMAL CONTROL SYSTEM MEL

WBS Description Qty Unit

Mass

(kg)

CBE

Mass

(kg)

Growth

(%)

Growth

(kg)

Total

Mass

(kg)

01 Robotic Lunar Lander --- -------- 429 7 33.3 462

01.01 LOI Stage --- -------- 25 15 4.5 29.5 

01.02 Lunar Descent Stage --- -------- 266 0 1.1 266.8 

01.03 Lunar Lander Stage --- -------- 138 16.7 27.7 165.7 

01.03.10 Thermal Control (Non-Propellant) --- -------- 18 15 3.3 21.7 

01.03.10.01 Active Thermal Control --- -------- 9.5 15 1.7 11.2 

01.03.10.01.01 Heaters 31 0.3 9.0 18 1.6 10.6 

01.03.10.01.02 Thermal Control/Heaters Circuit 0 0.0 0.0 18 0.0 0.0 

01.03.10.01.03 Data Acquisition 0 0.0 0.0 18 0.0 0.0 

01.03.10.01.04 Thermocouples 50 0.0100 0.5 18 0.1 0.6 

01.03.10.02 Passive Thermal Control --- -------- 8.9 15 1.6 10.5 

01.03.10.02.01 Heat Sinks 1 0.3 0.3 18 0.1 0.3 

01.03.10.02.02 Heat Pipes 0 0.0 0.0 18 0.0 0.0 

01.03.10.02.03 Radiators 1 0.8 0.8 18 0.2 1.0 

01.03.10.02.04 MLI 4 1.7 6.7 18 1.2 8.0 

01.03.10.02.05 Temperature sensors 0 0.0 0.0 18 0.0 0.0 

01.03.10.02.06 Phase Change Devices 0 0.0 0.0 18 0.0 0.0 

01.03.10.02.07 Thermal Coatings/Paint 1 1.0 1.0 18 0.2 1.2 

01.03.10.03 Semi-Passive Thermal Control --- -------- 0.0 0 0.0 0.0 

4.6.4 Thermal Trades

No top-level trades were performed during this study. Additional ones on the value of using different 

thermal schemes are to be done in the future.

4.6.5 Thermal Analytical Methods

Microsoft Excel (Microsoft Corporation) based internal tools and C&R Technologies “Thermal Desktop” 

software were used in the analysis. The Thermal Desktop model used the simple model of the internal 

components shown in Figure 4.16. Figure 4.17 shows the results of applying the assumptions above in 

Section 4.6.5 to the Thermal Desktop simplified model. 

Table 4.10 shows the results of the thermal desktop modeling through a snapshot of the Thermal 

Desktop modeling tool. 

Figure 4.16—Thermal desktop simplified model. Figure 4.17—Thermal desktop simplified model analysis results.
NASA/TM—2010-216218 36

TABLE 4.10—THERMAL CONTROL SYSTEM THERMAL DESKTOP MODELING

Radiators Input Input Input Input Input Input Output Input Output Input Input Output

Q,

(W)

e,

(ε)

Fin,

(eff.)

σ,

(W/m2K)

Ts,

(K)

Tb,

(C)

Box X,

(cm)

Box Y,

(cm)

Box Z,

(cm)

Tb,

(K)

(Tb4–Ts4) Area,

(m2)

Radiator 

X,

(m)

Radiator 

Y,

(m)

Radiator 

Z,

(m)

Radiator 

density,

(kg/m3

)

Radiator 

mass,

(kg)

Batteries 3 0.9 80% 5.67×10–8 270 20 15 15 15 293.15 2.07×109 0.035 0.150 0.237 0.003 2770 0.295

Electronics 15 0.9 80% 5.67×10–8 270 40 30 10 10 313.15 4.30×109 0.085 0.450 0.190 0.003 2770 0.710

RF assembly 0.95 0.9 80% 5.67×10–8 270 40 10 8 12 313.15 4.30×109 0.005 0.100 0.054 0.003 2770 0.045

Transponder 0.95 0.9 80% 5.67×10–8 270 40 20 16 11 313.15 4.30×109 0.005 0.200 0.027 0.003 2770 0.045

General avionics 0.95 0.9 80% 5.67×10–8 270 40 13.35 11.4 11.4 313.15 4.30×109 0.005 0.134 0.041 0.003 2770 0.045

1.140

Heaters Qty Unit 

mass, 

(kg)

Total 

mass,

(kg)

Notes

Fuel tank 1 5.0 5 Common Lunar Lander used 42 kg for 900 kg of fuel. For our 50 kg of fuel we need ~42/900*50~2.4*FS of 2~5 kg

Fuel line 20 0.1 2 JIMO estimate

Battery 10 0.2 2

Totals 31 0.3 9

MLI Qty Mass 

(kg)

Ti density 

(kg/m3)

Ti Cp

(J/kg K)

Volume

(m3)

Diam

(cm)

Length

(m)

Wall 

thickness

(m)

Surface 

area

(m2

)

MLI 

layers

MLI 

specific 

weight

(kg/m2

)

MLI

unit mass

(kg)

Total MLI 

mass

(kg)

Hydrazine fuel tank (sphere) 1 45 4500 520 0.0100 53.34 0.89383 100 2.856 2.55278

Solid rocket: Star 24C 

(sphere)

1 238 4500 520 0.0529 63.5 1.26677 50 1.428 1.8 1.80895

Hydrazine fuel lines 1 3.53 4500 520 0.0008 1 5 0.005 0.15708 30 0.8568 0.1 0.13459

Structure 1 5 10 1.5708 50 1.428 2.2 2.24310

Totals 4 1.7 6.7

MLI specific weight (kg/m2), JPL 10- by 10-cm, 15 layer sample measured, weight = 4.283 grams.

Heater power Input Input Input Input Input Input Output Output Output Input Input Output

Q,

(W)

e,

(ε)

n,

(eff.)

σ,

(W/m2K)

Ts,

(K)

Tb,

(C)

Tb,

(K)

(Tb4–Ts4) Area,

(m2)

Radiator 

X,

(m)

Radiator 

Y,

(m)

Radiator 

Z,

(m)

Radiator 

density,

(kg/m3

)

Radiator 

mass,

(kg)

Batteries 3 0.9 80% 5.67×10–8 4 –60 213.150 2.06×109 0.036 0.150 0.237 0.003 2770 0.295

Batteries + heaters 8.66 0.9 80% 5.67×10–8 4 5 278.150 5.99×109 0.035 0.188 0.188 0.003 2770 0.295

Battery heater power 5.66

Electronics 15 0.9 80% 5.67×10–8 4 –17 256.150 4.31×109 0.085 0.292 0.292 0.003 2770 0.709

Electronics +heaters 16.7 0.9 80% 5.67×10–8 4 –10 263.150 4.80×109 0.085 0.292 0.292 0.003 2770 0.709

Electronics heater power 1.7

Avionics 3.8 0.9 80% 5.67×10–8 4 –17 256.150 4.31×109 0.022 0.147 0.147 0.003 2770 0.180

Avionics + heaters 4.3 0.9 80% 5.67×10–8 4 –10 263.150 4.80×109 0.022 0.148 0.148 0.003 2770 0.183

Avionics heater power 0.5

T initial 

(C)

Time 

(s)

Q req for 

heat up

(W) a

Hydrazine fuel tank 15.01 0.1 80% 5.67×10–8 270 35 308.150 3.70×109 0.8938 20 36000 9.75

Hydrazine fuel tank + heaters 24.97 0.1 80% 5.67×10–8 4 7 280.150 6.16×109 0.8938

Hydrazine fuel tank heater 

power

9.96

a

Fuel lines 2.64 0.1 80% 5.67×10–8 270 36 308.150 3.70×109 0.1571 7 36000 1.43

Fuel lines + heaters 4.39 0.1 80% 5.67×10–8 4 7 280.150 6.16×109 0.1571

Fuel lines heater power 1.75 a

Assume allowable temperature ranges of 7 to 35 °C

Q = power dissipated

A = radiator area

e = emissivity of box top surface

n = fin effectiveness of box top surface

σ = Stefan-Boltz constant

Tb = temperature of box top surface

Ts = Moon sink temperature for horizontal surface (radiator at equator and angle of the sun above the horizon is 90°)

A = Q/(e*n*σ*(Tb4 – Ts4

))

Assume 40 °C Tb for electronics

Assume 20 °C Tb for batteries

Assume horizontal radiators at lunar equator
NASA/TM—2010-216218 37

Figure 4.18—Thermal desktop simplified model.

Figure 4.18 shows a graphical representation of the thermal environment modeled over a lunar 

day/night cycle.

4.6.6 Thermal Risk Inputs

To be completed in the future.

4.6.7 Thermal Recommendation

None.

5.0 Cost, Risk and Reliability

5.1 Costing: Baseline Chemical Lunar Lander

5.2 Cost Modeling Assumptions

Cost estimates were based on the Case 1 (baseline) MEL received on September 29, 2006. The scope 

of the estimates is prime contractor costs for DDT&E and first-unit production. Costs for contractor fee, 

government support and operations are not included. Launch vehicle costs are not included. Estimates are 

based on a mixture of subsystem level estimates (Avionics, Thermal Control, and Structure) and estimates 

developed at the component level (Propulsion and Power). The estimates provided represent the mean of 

the probability distribution and are in FY07 dollars. A protoflight approach was assumed for development 

and testing.

5.3 Cost Modeling Approach

The costs were estimated in Microsoft Excel (Microsoft Corporation) using an approach similar to 

NAFCOM, i.e., the elements of a product-oriented WBS were estimated and the sum of these product 
NASA/TM—2010-216218 38

costs were used to estimate system integration costs (“wraps”). Estimates generated account for all prime 

contractor costs; however they do not break out labor and material costs. @Risk v. 4.5.5, a Monte Carlo 

simulation application that integrates with Excel, was used for probabilistic risk assessment. Wherever 

possible, the Cost Estimating Relationships (CER) were developed using as many relevant data points as 

available so the standard errors used to develop the risk model would have a strong statistical basis. All 

in-house CERs were developed using DataFit v. 8.1.69 with the dual priorities of minimizing standard 

error and maximizing the probability that the selected independent variables have significant correlation 

with regards to cost. Table 5.1 shows the Phase C/D Cost Estimates.

TABLE 5.1—PHASE C/D COST ESTIMATES

FY07 $M: D&D Flight HW Total

LOI Stage (incl. with structures 

and mechanisms)

Lunar Descent Stage 0.7 0.5 1.2

Avionics, Instrumentation, GN&C 24.4 12.0 34.0

Propulsion & Propellant Management (Chemical) 3.5 1.3 4.9

Power 1.4 2.9 4.3

Thermal Control 3.4 1.5 4.9

Structures and Mechanisms 10.5 5.6 16.1

Subsystem Subtotal 44.0 23.9 65.4

Systems Integration

Integration, Assembly & Checkout 5.1 1.9 7.0

System Test Operations 2.7 2.7

Ground Support Equipment 3.5 3.5

Systems Engineering & Integration 2.9 1.4 4.2

Project Management 6.1 2.9 9.0

Launch Operations & Orbital Support 4.6 4.6

Systems Integration Subtotal 24.9 6.2 31.1

Total Prime Contractor 68.9 30.0 99.0

6.0 Trade Space Iterations

6.1 Case 1: Off-the-Shelf Chemical Propulsion

The baseline case (case 1) used all chemical propulsion systems to perform burns during the mission. 

The two trades done on the baseline involved the application of electric propulsion thrusters and low 

thrust mission analysis to the Lunar Lander problem. See Table 3.3 for the Case 1 system summary chart.

6.2 Case 2: Off-the-Shelf Electric Propulsion

This design used OTS Electric propulsion as the lunar propulsion system. The mission is launched on 

a Minotaur V to TLI, uses SEP to spiral to LLO, and a Solid/Monoprop propulsion system on the Lander. 

Unfortunately, this trade yielded no useable payload to the lunar surface. See the summary in Table 6.1.
NASA/TM—2010-216218 39

TABLE 6.1—CASE 2 SYSTEM SUMMARY CHART

Case 2—Robotic Lander Stage masses 

(kg)

Current Values with growth margin 

(kg) 

Totals Percent of 

dry mass 

Subsystems EDS 

stage 

Descent 

stage 

(solid) 

Lander 

stage 

(mono-p) 

Total 

Robotic 

Lander 

Total 

Robotic 

Lander 

Avionics 4.29 0.00 36.79 41 11 

Propulsion (chemical) 0 14.37 5.61 20 5 

Prop Management (chemical) 0 0.00 9.20 9 3 

Propellant (chemical) 0 105.72 29.79 136 N/A 

Propulsion (EP) 37.51 0.00 0 38 10 

Prop Management (EP) 21.60 0 0 22 6 

Propellant (EP) 142.49 0 0 142 N/A 

Power Processing (EP) 0 0 0 0 0 

Power 88.77 0 30.85 120 33 

Thermal Control 26 0 21.69 48 13 

Structure 52.38 0 17.69 70 19 

Total Mass (Wet) 373 120 152 645 N/A 

Lander Total Wet Mass 645 

Lander Total Dry mass (kg) 367 110 

Lander Total Inert Mass (kg) 391 

Payload (left over) –6 

6.3 Case 3: Advanced Direct Drive Electric Propulsion

The summary of the Case 3 design components is as follows: “Advanced” Direct-Drive Electric 

Propulsion, Minotaur V to TLI, and spiral to LLO, solid/monoprop lander. The case yielded a 1-yr 

transfer to the Moon and 27 kg useable payload on lunar surface.

In an attempt to make EP work, mass needs to be removed from the EP stage. Some options to reduce 

mass are as follows:

 Utilize ‘direct-drive’ technology

− 500 V solar array directly fed to Hall thruster(s)

− Thruster ancillaries fed by lander 28 V power

− Eliminates power converters, and most of Power Processing Units (PPUs), most of radiators, 

(>60 kg savings)

 Start in Geostationary Transfer Orbit (GTO) off an Minotaur V (case transportable to secondary 

payload consideration)

− More starting mass (710 kg)

− Faster trip time

− Less Radiation on spacecraft and array

− Similar to SMART-1 trajectory

Table 6.2 is the system summary results of the Case 3 design using direct drive EP.
NASA/TM—2010-216218 40

TABLE 6.2—CASE 3 SYSTEM SUMMARY CHART

Case 3—Robotic Lander Stage Masses 

(kg)

Current values with growth margin

(kg) 

Totals Percent of 

dry mass

Subsystems EDS 

stage

Descent 

stage 

(solid)

Lander 

stage 

(mono-p)

Total 

Robotic 

Lander

Total 

Robotic 

Lander

Avionics 0 0.00 44.45 44 15 

Propulsion (chemical) 0 19.06 5.61 25 9 

Prop Management (chemical) 0 0.00 9.20 9 3 

Propellant (chemical) 0 147.81 36.65 184 N/A

Propulsion (EP) 15.40 0.00 0 15 5 

Prop Management (EP) 21.60 0 0 22 7 

Propellant (EP) 139.57 0 0 140 N/A

Power Processing (EP) 0 0 0 0 0 

Power 47.71 0 33.73 81 28 

Thermal Control 6 0 21.69 28 10 

Structure 45.13 0 18.56 64 22 

Total Mass (Wet) 275 167 170 612 N/A

Lander Total Wet Mass 612 

Lander Total Dry mass (kg) 288 86 

Lander Total Inert Mass (kg) 314 

Payload (left over) 26.9 

The mission timeline for Case 3 is as follows:

 Minotaur V to GTO

 EP Spiral from GTO to LLO (a.k.a. SMART-1)

 Direct drive Hall thruster system to minimize mass

 Solid rocket motor for descent

 Hydrazine propulsion system for final landing

This provides for the delivery of a science payload onto the lunar surface capable of 1 week 

operation, landing in a sunlit area on the lunar surface, Global Access, 25 W for payload on lunar surface.

The Case 3 vehicle consisted of three stages:

Stage 1: Chemical Lander Stage (Similar to the Case 1 Chemical Lander)

 High Gain Antenna

 125 W GaAs Solar Array (two-axis pointing)

 Blow-down monoprop system for final landing and pointing

 Avionics on top for improved cooling

 Fixed leg landing system using Doppler radar (antennas on feet)

Stage 2: Descent Stage (Similar to the Case 1 Chemical Lander)

 Star 24 Solid Rocket (30% offload) (different motor than case 1 but same class, disposed before 

landing)

Stage 3: EP Stage (This Replaces the TLI Performed by the Minotaur V in Case1)

 3 kW Ultraflex Solar Array

 Xe Tanks (130 kg Xe)

 HiVHAC Hall Thrusters

 Launch vehicle adapter re-used for EP stage structure

Figure 6.1 shows the stacked version of the Case 3 vehicle with the Ultraflex solar array in stowed 

position. Figure 6.2 shows the Case 3 stack rendered with the Ultraflex solar array deployed.
NASA/TM—2010-216218 41

Figure 6.1—Case 3 Robotic Lunar Lander design.

Figure 6.2—Case 3 Robotic Lander, solar panel for EP system deployed.

EP system better choice when system can be reused after lunar arrival as a Relay/com/nav station, for 

additional power or as a station keeping propulsion system. Very low lunar orbit observer stationkeeping. Further collaboration with Relay Sat study and ARC planned. 

NASA/TM—2010-216218 43

Appendix A.—Acronyms and Abbreviations

ARC Ames Research Center

C&DH Command and Data Handing

CAD computer aided design

CBE current best estimate

CER Cost Estimating Relationship

CEV Crew Exploration Vehicle

Comm Communications

COMPASS COllaborative Modeling and 

Parametric Assessment of Space 

Systems

COPV Carbon-Overwrapped Pressure 

Vessels

COTS Commercial off the Shelf

DMR Design for Minimum Risk

DSN Deep Space Network

DTE direct to Earth

DDT&E Design Development Testing and 

Engineering

EELV Evolved Expendable Launch 

Vehicle

EP Electric Propulsion

FEA finite element analysis 

FEM Finite Element Model

FOM figure of merit

GLIDE GLobal Integrated Design 

Environment

GN&C Guidance, Navigation and Control

GRC NASA Glenn Research Center

GTO Geostationary Transfer Orbit

HAPS Hydrazine Auxiliary Propellant 

System 

HQ NASA Headquarters

HiVHAC High Voltage Hall ACcelerator)

IP internet protocol

LAT Lunar Architecture Team

LEO low Earth orbit

LLO low lunar orbit 

LOI Lunar Orbit Insertion

LPRP Lunar Precursor and Robotic 

Program

LSP Launch Service Program

LSTO Launch Service Task Order

MAC Media Access Control

MEL Master Equipment List

MLI Multilayer Insulation

MPU Makeup Power Unit

NASA National Aeronautics and Space 

Administration

Nav navigation

NLS NASA Launch Services 

OTS Off the Shelf

PEL Power Equipment List

PMS Propellant Management System

PN pseudo-noise

PPU Power Processing Units

RF radio frequency

RCS Reaction Control System

S/C spacecraft

SADA Solar Array Drive Assembly

SEP Solar Electric Propulsion

SN signal-to-noise

SPU Solar Power Unit

TDRSS Tracking and Data Relay Satellite 

System

TLI trans-lunar injection

TOF time of flight

TWTA Traveling Wave Tube Amplifier

WSB weak stability boundary

NASA/TM—2010-216218 45

Appendix B.—Case 1 Rendered Design Drawings

Figure B.1 shows the total Case 1 design with Star 47 motor, adaptor, and lander stack. Figure B.2

shows the bottom view of the total Case 1 design with Star 47 motor, adaptor, and lander stack as looking 

up the Star 47 Rocket motor.

Figure B.1—Case 1 Robotic Lander, stack rendered.

Figure B.2—Case 1 Robotic Lander, bottom view.
NASA/TM—2010-216218 46

Figure B.3 is the landed portion of the Robotic Lunar Lander as it would appear on the Moon’s 

surface with the solar array and antenna deployed.

Figure B.3—Case 1 landed portion of Robotic Lunar Lander.
NASA/TM—2010-216218 47

Appendix C.—Study Participants

Low Cost Robotic Lunar Lander Design Session

Subsystem Name Center Email

Customer ARC

Customer Lunar Architecture Team (LAT) HQ

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

Mission Design Doug Fiehler GRC Douglas.I.Fiehler@nasa.gov

Mission Design Craig Kluever UMo KlueverC@missouri.edu

Concept Design Integration Lead, 

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

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

System Integration Dave Grantier GRC David.T.Grantier@nasa.gov

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

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

Thermal John Siamidis GRC John.Siamidis-1@nasa.gov

Power James Fincannon GRC Homer.J.Fincannon@nasa.gov

Communications and Avionics Rich Oeftering GRC Richard.C.Oeftering@nasa.gov

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

Guidance, Navigation and Control Jeffrey Lee GRC Jeffrey.R.Lee@nasa.gov

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

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

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

Risk/Reliability Bill Strack GRC bstrack@wowway.com
NASA/TM—2010-216218 48

Bibliography

Aerojet engines and propellant systems (http://www.aerojet.com/capabilities/spacecraft.php)

Arde Inc tank Catalogs: http://www.ardeinc.com

Astronautix Web site: http://www.astronautix.com

ATK PSI tank catalogs: http://www.psi-pci.com

Beagle 2 official website: http://beagle2.open.ac.uk/index.htm

Belbruno, E. A., and J. P. Carrico (2000), Calculation of Weak Stability Boundary Ballistic Lunar 

Transfer Trajectories, AIAA/AAS Astrodynamics Specialist Conference, American Institute of 

Aeronautics and Astronautics, Denver, CO, abstract AIAA 2000-4142, August 14-17.

Orbital Science Corporation, Minotuar V Fact Sheet: 

http://www.orbital.com/NewsInfo/Publications/Minotaur_V_Fact.pdf.

Parker, J.S. “Families of Low Energy Lunar Halo Transfers,” AAS/AIAA Spaceflight Dynamics 

Conference, Paper No. AAS 06-132, Tampa, FL, January 22-26, 2006.

Recommended Practice for Mass Properties Control for Satellites, Missiles, and Launch Vehicles, 

ANSI/AIAA R-020A-1999, 1999.

Space Mission Analysis and Design, 3rd edition (Space Technology Library), Wiley J. Larson and James 

R. Wertz (eds.), October 1999.

Star Motor reference pages from ATK. (http://www.atk.com/starmotors/starmotors_star48v.asp)

Surveyor Web site: http://www.lpi.usra.edu/expmoon/surveyor/surveyor.html.
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13. SUPPLEMENTARY NOTES

14. ABSTRACT

The COllaborative Modeling for the Parametric Assessment of Space Systems (COMPASS) team designed a robotic lunar Lander to deliver 

an unspecified payload (greater than zero) to the lunar surface for the lowest cost in this 2006 design study. The purpose of the low cost 

lunar lander design was to investigate how much payload can an inexpensive chemical or Electric Propulsion (EP) system deliver to the 

Moon’s surface. The spacecraft designed as the baseline out of this study was a solar powered robotic lander, launched on a Minotaur V 

launch vehicle on a direct injection trajectory to the lunar surface. A Star 27 solid rocket motor does lunar capture and performs 88 percent 

of the descent burn. The Robotic Lunar Lander soft-lands using a hydrazine propulsion system to perform the last 10% of the landing 

maneuver, leaving the descent at a near zero, but not exactly zero, terminal velocity. This low-cost robotic lander delivers 10 kg of science 

payload instruments to the lunar surface. 

15. SUBJECT TERMS

Low cost; Robotics; Electric propulsion; Chemical propulsion; Lunar surface; Launch vehicles; Solid propellant 

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