Kathryn K. Trase, Rachel A. Barch, Ryan E. Chaney, Rachel A. Coulter, Hui Gao, David P. Huynh, 

Nicholas A. Iaconis, Todd S. MacMillan, Gregory M. Pitner, and Devin T. Schwab

Glenn Research Center, Cleveland, Ohio

SCience Hybrid Orbiter and Lunar Relay (SCHOLR) 

Architecture and Design

NASA/TM—2011-216894

July 2011

AIAA–2010–813610
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Kathryn K. Trase, Rachel A. Barch, Ryan E. Chaney, Rachel A. Coulter, Hui Gao, David P. Huynh, 

Nicholas A. Iaconis, Todd S. MacMillan, Gregory M. Pitner, and Devin T. Schwab

Glenn Research Center, Cleveland, Ohio

SCience Hybrid Orbiter and Lunar Relay (SCHOLR) 

Architecture and Design

NASA/TM—2011-216894

July 2011

AIAA–2010–813610

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

Space 2010 Conference and Exposition

sponsored by the American Institute of Aeronautics and Astronautics

Anaheim, California, August 30–September 2, 2010
Acknowledgments

The authors would like to thank Dr. Kul Bhasin for his mentorship and passionate support of the SCHOLR Project. The authors 

also thank members of the Collaborative Modeling for Parametric Assessment of Space Systems (COMPASS) Team at NASA 

Glenn Research Center who provided much of the technical assistance essential to SCHOLR’s design—specifically, Laura Burke, 

Kristen Bury, Anthony Colozza, James Fincannon, James Fittje, John Gyekenyesi, Jennifer Jordan, Geoffrey Landis, Michael 

Martini, Melissa McGuire, Steve Oleson, Thomas Packard, and Glenn Williams. The authors would like to recognize its SCaN 

mentors: Brian Barritt, Michael Fuentes, Bertsel Golden, and Eric Knoblock for their guidance throughout this project as well as 

Amber Gallihar and Katy Kafantaris for their editorial and graphic assistance. Lastly, the authors wish to thank NASA SCaN’s 

Office of Policy and Strategic Communications for its financial support of this project.

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Trade names and trademarks are used in this report for identification 

only. Their usage does not constitute an official endorsement, 

either expressed or implied, by the National Aeronautics and 

Space Administration.

Level of Review: This material has been technically reviewed by technical management. 
NASA/TM—2011-216894 1

SCience Hybrid Orbiter and Lunar Relay (SCHOLR) 

Architecture and Design

Kathryn K. Trase, Rachel A. Barch,*

National Aeronautics and Space Administration

Ryan E. Chaney,* Rachel A. Coulter,* Hui Gao,* David P. Huynh,*

Nicholas A. Iaconis,* Todd S. MacMillan,* Gregory M. Pitner,* and Devin T. Schwab*

Glenn Research Center

Cleveland, Ohio 44135

Abstract

Considered both a stepping-stone to deep space and a key to unlocking the mysteries of planetary 

formation, the Moon offers a unique opportunity for scientific study. Robotic precursor missions are 

being developed to improve technology and enable new approaches to exploration. Robots, lunar landers, 

and satellites play significant roles in advancing science and technologies, offering close range and in-situ 

observations. Science and exploration data gathered from these nodes and a lunar science satellite is 

intended to support future human expeditions and facilitate future utilization of lunar resources. To attain 

a global view of lunar science, the nodes will be distributed over the lunar surface, including locations on 

the far side of the Moon. Given that nodes on the lunar far side do not have direct line-of-sight for Earth 

communications, the planned presence of such nodes creates the need for a lunar communications relay 

satellite. Since the communications relay capability would only be required for a small portion of the 

satellite’s orbit, it may be possible to include communication relay components on a science spacecraft. 

Furthermore, an integrated satellite has the potential to reduce lunar surface mission costs. A SCience 

Hybrid Orbiter and Lunar Relay (SCHOLR) is proposed to accomplish scientific goals while also 

supporting the communications needs of landers on the far side of the Moon. User needs and design 

drivers for the system were derived from the anticipated needs of future robotic and lander missions. 

Based on these drivers and user requirements, accommodations for communications payload aboard a 

science spacecraft were developed. A team of interns identified and compared possible SCHOLR 

architectures. The final SCHOLR architecture was analyzed in terms of orbiter lifetime, lunar surface 

coverage, size, mass, power, and communications data rates. This paper presents the driving 

requirements, operational concept, and architecture views for SCHOLR within a lunar surface nodal 

network. Orbital and bidirectional link analysis, between lunar nodes, orbiter, and Earth, as well as a 

conceptual design for the spacecraft are also presented. 

1.0 Introduction

As Earth’s nearest celestial neighbor, the Moon is both a gateway to deep space exploration and a 

feasible platform from which to conduct scientific study of planetary formation. Although past lunar 

missions have provided great insight into the Moon’s many mysteries, much science has yet to be 

conducted. Samples returned from Apollo and Luna missions, in addition to orbital data gathered from 

such satellite missions as Clementine and Lunar Prospector, provided a wealth of information on the lunar 

system. More recently, international orbiters and NASA’s Lunar Reconnaissance Orbiter (LRO) are 

returning more complete and sophisticated data sets. However, even these missions leave questions to be 

answered and discoveries to be made about the Moon’s origin and processes as well as the presence of 

water (Ref. 1). 

Lunar surface missions are being proposed to help close the gaps in the current understanding of the 

Moon and planetary processes in general. These landers and robots will be deployed across the near and 

 *

Summer interns supported by the Lewis’ Educational and Research Collaborative Internship Program (LERCIP), a 

collaborative undertaking by the NASA Glenn Research Center and the Ohio Aerospace Institute
NASA/TM—2011-216894 2

far sides of the Moon to better obtain a global perspective of the Moon’s environment and processes. For 

assets on the far side, communication with Earth is impossible, and for many locations on the near side, 

communication can be difficult. Including a powerful communications system or planning for a partner 

lunar relay satellite would quickly increase mission costs. A SCience Hybrid Obiter and Lunar Relay 

(SCHOLR) is proposed to provide communications capabilities to these landers as they explore the lunar 

surface at a reduced net cost. SCHOLR aims to serve as a data relay from the lunar surface to Earth in 

order to enable exploration of the lunar far side while additionally capturing scientific data on the lunar 

North Pole.

SCHOLR would expand our knowledge of the Moon’s environment by characterizing water content, 

surface composition, variety and distribution of lunar rocks, and crust thickness at the lunar North Pole. 

Whether preparing for prolonged human stay, mining lunar resources, or using the Moon as a technology 

demonstration platform, a more complete understanding of the surface will aid mission planners in 

selecting the best approach to produce a hardy system and emplace it in the most scientifically 

advantageous location on the Moon. Study of the lunar crust would also provide insight into planetary 

evolution and advance our knowledge of how the Earth was formed (Ref. 1).

As a secondary goal, SCHOLR may become a tool for creating excitement for space exploration 

among students. Inspired by man’s presence on the Moon during the Apollo era, the previous generation 

of young scientists and engineers were motivated by a sense of exploration and discovery. Presently, the 

fervor and excitement for space exploration has faded in the more than 40 yr since the Apollo missions, 

leaving today’s young scientists and engineers searching for a renewed source of inspiration and purpose. 

Given the immense advancement of current technologies and capabilities since the 1960s, the space 

programs of this era are poised to provide the new generation with limitless possibilities. Small lunar 

landers and robots could soon be developed by commercial institutions or university students and 

operated by grade school science classes. 

This paper presents SCHOLR’s mission context, innovations, and science and communication 

objectives in Section 2.0. Section 3.0 discusses high level mission requirements, system architecture, and 

concept of operations. SCHOLR’s science and communication payloads and orbit are presented in 

Section 4.0. The subsystems integration process is discussed in Section 5.0, and a conceptual discussion 

of the spacecraft subsystems follows: Structures; Environmental Control; Guidance, Navigation and 

Control (GN&C); Power; Command and Data Handling (C&DH); and Communications. 

2.0 SCHOLR Mission Context and Objectives

2.1 Mission Context

Both SCHOLR and surface assets would collectively contribute to the advancement of lunar science 

as well as general space technology. Lunar science data gathered from these sources would be used to 

demonstrate new technologies and serve to maintain scientific momentum as new discoveries are made 

that may be applied throughout NASA and the global community. The Object-Process Methodology

(Ref. 2) was used to develop the mission context in terms of the users, mission objectives, and the role of 

SCHOLR. The main context diagram illustrating the ultimate users of the data returned by SCHOLR and 

lunar surface assets are illustrated in Figure 1.

2.2 SCHOLR Innovations and State-of-the-Art

2.2.1 Hybrid Satellites

Concepts of “hybrid” satellites in a general sense are not new. Currently, the Canadian Space Agency 

has plans to launch a hybrid Earth satellite in 2011 (Ref. 3) with a primary science mission and secondary 

communications demonstration (Ref. 4). NASA’s Mars Reconnaissance Orbiter, launched in 2005, has 

taken scientific data of the Martian surface while also serving as a communications link for rovers back to 
NASA/TM—2011-216894 3

Earth (Ref. 5). Design teams have investigated constellations of lunar satellites to provide continuous 

communications coverage for the lunar surface (Refs. 6 and 7), but these designs have not incorporated 

the capability to collect scientific data on the spacecrafts. 

SCHOLR would be the first lunar science satellite to provide communications relay to assets on the 

lunar surface. The hybrid nature of SCHOLR would permit scientific observation of the North Pole when 

communications relaying is infeasible. Conversely, SCHOLR would serve as a data relay when practical 

science cannot be conducted. 

2.2.2 Science Instruments

The Science Instrument Payload (SIP) includes an Active Infrared Spectrometer (AIRS) that has 

never flown in space. This instrument would provide new insights into the water content of the Moon that 

previous spectrometers, passive or otherwise, have not been able to produce due to inherent limitations in 

the method of observation.

The optical communications technology demonstration would be a follow-on to the laser 

communication demonstration on the Lunar Atmosphere and Dust Environment Explorer (LADEE), 

expected to launch in 2013 (Ref. 8).

2.2.3 Mission Cost

By effectively providing the capability of two satellites within one orbiter, SCHOLR seeks to reduce 

mission costs. Although a noble quest by itself, scientific exploration is costly and careful considerations 

must be made to ensure a maximum return on investment. SCHOLR’s lifecycle cost is estimated to be 

between $600 to $700 million, including the cost of launch and ground operations for the nominal 3 yr

mission. SCHOLR’s relay capabilities may serve to reduce mission costs for future landers and rovers it 

would support by reducing the surface assets’ communications constraints and landing site restrictions.

2.3 Science Objectives

SCHOLR’s science goals and objectives were adapted from the priorities outlined in the 2007 study 

by the National Research Council, entitled “Scientific Context for Exploration of the Moon (SCEM)” 

(Ref. 1). SCEM sought to provide scientific guidance for future lunar activities by prioritizing lunar 

science goals. These science objectives include gathering additional data to answer questions of: lunar 
NASA/TM—2011-216894 4

bombardment history, polar environment, and volcanism; structure and composition; planetary, impact, 

and regolith processes; and permanently shadowed craters. SCEM gave consideration to the intended 

measurements by recent lunar orbiter missions, Chang’e-1, SELENE, Chandrayaan-1, and LRO, and new 

objectives were developed under the assumption that each mission would be successfully completed. 

Given the overall success of these programs, the goals outlined in the study were selected to become 

SCHOLR’s mission goals. SCHOLR’s science goals and objectives, as adapted from SCEM, are listed in 

order of importance in Table 1.

TABLE 1.—SCHOLR SCIENCE GOALS AND MEASUREMENT OBJECTIVES

Science goal Science measurement objective

Determine the extent of water on the lunar surface and in the 

subsurface.

Provide high resolution measurements of hydroxyl groups 

and identify water ice deposits.

Determine the age and chronology of lunar craters. Provide higher-resolution images of lunar terrains.

Determine the compositional state and compositional 

distribution of volatile components in lunar polar regions.

Provide high-spatial resolution distribution of volatiles on 

and in the regolith.

Determine the extent and composition of crust, KREEP layer. Provide higher-resolution geophysical measurements of 

representative regions.

Determine the thickness of lunar crust and its variability. Provide high-resolution gravity measurements.

Determine the variety, age, and distribution of lunar rock. Provide higher-resolution global and regional mineralogic 

and geochemical maps.

2.4 Communications Relay Objectives

The Lunar Communications Payload (LCP) in conjunction with the SIP would allow SCHOLR to 

fulfill its innovative role as a hybrid science and communications satellite. Since the lunar poles and lunar 

far side have minimal to no direct access to Earth, it is impossible to effectively communicate with 

surface assets in these areas of the Moon without a lunar relay. The LCP would provide surface assets at 

the lunar South Pole and far side with communications access to Earth, thus enabling surface missions to 

take place in low access, high scientific gain areas. To provide an effective means of communicating with 

surface assets in these areas, the LCP must support the large volume of science data generated by the 

surface assets, as well as low data rates for real-time command and monitoring of the assets. The LCP 

would also provide these capabilities to sites that do have direct to Earth (DTE) access to ease the 

communications burden (in terms of size, weight, and power) on those assets. By providing relay access 

to Earth at high rates, scientists could collect surface measurements from sites that have been impractical 

in the past due to limited DTE access. This data, when combined with in-orbit measurements, would give 

scientists a more accurate and global picture of the Moon. 

3.0 High Level Mission Requirements and Architecture

High-level mission requirements were derived from NASA’s lunar exploration objectives and are 

discussed in Section 3.1. SCHOLR’s system architecture was developed following Department of 

Defense Architecture Framework and can be found in Section 3.2. From these requirements and 

architecture, SCHOLR’s concept of operations was created and is presented in Section 3.3. 

3.1 High-Level Requirements

Considering both the scientific and communications objectives of SCHOLR, high-level mission 

requirements were developed to ensure those goals would be realized. The high-level mission 

requirements include:

• SCHOLR shall collect lunar scientific data to fill knowledge gaps of the lunar North Pole at both 

the surface and subsurface levels.
NASA/TM—2011-216894 5

• SCHOLR shall provide the capability to relay science data and real time commands between the 

Mission Operations Center (MOC) and the lunar surface mission.

• SCHOLR shall characterize exploration sites and transportation routes for future lunar surface 

missions.

• SCHOLR shall provide the capability to perform communications technology demonstrations.

• SCHOLR shall demonstrate the communications relay capability on the lunar far side below the 

lunar equator.

From these high-level requirements, the Science Instrument Payload (SIP) was selected and mission 

orbit and concept of operations were determined. 

3.2 System Architecture

SCHOLR’s system architecture was derived based on the high-level requirements with the 

Department of Defense Architecture Framework (Ref. 9). In the system view, SCHOLR would perform 

scientific observation of the lunar surface 100 km above the North Pole when communication with 

surface assets would be impractical. While orbiting 5798 km over the South Pole, SCHOLR would 

function as a data relay for surface assets within view when scientific measurement would be infeasible. 

SCHOLR’s system views can be seen in Figure 2. 

In the operational view, science data gathered by surface assets would be received by SCHOLR and 

transmitted to Earth through future 18-m dish ground stations at the Deep Space Network (DSN) sites, a 

component of the Space Communications and Navigation (SCaN) network. Science data gathered by 

SCHOLR would also be sent to Earth through the same network. SCaN networks would send raw science 

data to the Science Data Processing Center, where the data could be prepared for release to the public. 

Both the science community and international entities would make use of the data generated by the 

surface assets and SCHOLR. Telemetry data would also be received by SCaN ground stations and 

forwarded to SCHOLR’s MOC. The MOC would then uplink new commands and telemetry data through 

SCaN networks to SCHOLR to continue the mission. Figure 3 depicts the operational view of SCHOLR. 
NASA/TM—2011-216894 6

3.3 Mission Concept of Operations

SCHOLR would launch in 2016 from NASA’s Kennedy Space Center aboard a Falcon 9 launch 

vehicle. Following launch, SCHOLR would begin a trans-lunar injection after separation from the 

Falcon 9 vehicle. A series of course correction burns would be performed to ensure SCHOLR would 

reach the Moon’s gravitational field. Arriving at the Moon four days after launch, SCHOLR would 

perform several capture burns to enter a highly elliptical polar orbit with an 8 hr period, requiring 

approximately 500 m/s delta-V. SCHOLR would remain in this orbit for a nominal 3 yr mission. 

Upon reaching lunar orbit, SCHOLR would deploy the radar sounder boom and verify the subsystems 

are functioning properly. SCHOLR would gather scientific data as it passes over relevant sites at 100 km 

above the lunar North Pole, while additionally continuing to communicate with the Earth. During the pass 

5798 km above the lunar South Pole, SCHOLR would relay its own data, as well as data collected by 

surface assets in the southern hemisphere, back to Earth at 200 Mbps. SCHOLR’s mission architecture is 

represented in the system views seen in Figures 2(a) and (b). SCHOLR would periodically perform orbital 

maintenance burns of approximately 400 m/s delta-V to counteract the influence of the Moon’s uneven 

gravitational field.

When in the shadow of the Moon, SCHOLR would enter into a power-save mode, where 

functionality of the SIP and other spacecraft components are limited to conserve the battery. These 

periods of shadow are expected to occur several times throughout the mission. Upon re-entry into the 

sunlight, SCHOLR would produce power using the solar array.

Upon successful completion of the nominal mission, SCHOLR could perform an orbit maneuver to 

reach a frozen elliptical orbit for a 4 yr extended mission. SCHOLR would remain in this low 

maintenance orbit through the end of the mission. SCHOLR would primarily function as a 

communications relay for surface assets during this extended mission.
NASA/TM—2011-216894 7

4.0 SCHOLR Payloads and Orbit

Following development of SCHOLR’s mission requirements and architecture, the science instrument 

and communications relay payloads were determined. The instrument requirements and payload selection 

is presented in Section 4.1. Section 4.2 discusses the communications relay payload. To best support the 

needs of both payloads, SCHOLR’s orbit was optimized. The results of this analysis are presented in 

Section 4.3.

4.1 Science Instrument Payload 

High-level science requirements were developed to accomplish the mission goals and objectives 

recommended in SCEM while additionally supporting the high-level mission requirements. SCHOLR’s 

SIP was selected to best support these science mission requirements at the lowest cost and weight. The 

SIP includes a stereo camera, Active Infrared Spectrometer (AIRS), Gamma Ray Spectrometer (GRS), 

and a radar sounder. An optical communications technology demonstration would also be flown with the 

SIP. The first priority of SCHOLR is to investigate the presence of water ice at the lunar North Pole. By 

flying the SIP, our understanding of the Moon would be advanced through SCHOLR’s investigation of 

the presence of the hydroxyl functional group, hydrogen, and water ice. Table 2 summarizes SCHOLR’s 

science requirements and denotes the instruments that contribute to satisfying those requirements.

TABLE 2.—SCHOLR SCIENCE REQUIREMENTS AND INSTRUMENTS

Science measurement requirement Stereo 

camera

Active IR 

spectrometer

Gamma ray 

spectrometer

Radar 

sounder

Identify hydroxyl functional groups at the lunar North Pole •

Identify water ice deposits at the lunar North Pole • • •

Obtain stereo imaging of North Pole lunar surfaces at multiple 

spatial resolutions and at multiple illumination angles

•

Characterize the distribution of volatiles on and in the regolith at 

the lunar North Pole

• •

Map lunar mineralogy and geochemistry in the North Pole • • •

Provide high-resolution gravity measurements • • •

4.1.1 Stereo Camera

The SCHOLR High Resolution Stereo Camera’s measurement objective is to provide targeted crater 

counts as well as reference photography for data from the other sensors in the SIP. Borrowing 

significantly from the heritage of the Lunar Reconnaissance Orbiter Camera (LROC), this camera 

package would provide detailed images in three dimensions of areas of interest to the science community. 

Two narrow angle cameras taking overlapping pictures as the spacecraft moves through orbit would allow 

for three dimensional terrain maps to be constructed. A wide-angle camera would provide contextual 

images for the science instruments.

4.1.2 Active Infrared Spectrometer (AIRS)

The SCHOLR AIRS would use infrared lasers to characterize the presence of the hydroxyl functional 

group in areas without illumination on the Moon, such as permanently shadowed craters. AIRS aims to 

fill in where the Moon Mineralogy Mapper flown on Chandrayaan-1 left off (Ref. 11). In much the same 

way as the Lunar Orbiter Laser Altimeter (LOLA) measured surface roughness; the AIRS would use three 

lasers in the mid-infrared spectrum, at 2.7, 2.8, and 3.1 μm, to determine the presence of OH via 

absorption: hydroxyl concentration would be determined by the intensity of the laser pulses reflected by 

the lunar surface.

4.1.3 Gamma Ray Spectrometer (GRS)

The SCHOLR GRS would create high-resolution maps of the chemical composition of the lunar 

surface. It would detect characteristic gamma rays from lunar aluminum, calcium, hydrogen, iron, 
NASA/TM—2011-216894 8

magnesium, oxygen, potassium, silicon, thorium, titanium, and uranium. These measurements would 

contribute to our knowledge of the lunar crust composition and planetary formation processes. 

SCHOLR’s GRS aims to improve upon the energy resolution and precision of the germanium gamma-ray 

spectrometer aboard Japan’s SELENE spacecraft (Ref. 12).

4.1.4 Radar Sounder

The SCHOLR radar sounder would investigate lunar subsurface ice with two 15-m dipole antennas. 

The radar sounder would emit radio waves at 5 MHz, a frequency able to penetrate the lunar surface as far 

as 25 km. A circularly polarized radar reflection would indicate the presence of water ice. Additionally, 

data from this instrument could be combined with data previously acquired from the LOLA (Ref. 13) to 

determine the thickness and subsurface structure of the Moon’s crust.

4.1.5 Optical Communications Technology Demonstration

A portion of SCHOLR’s mass and power budgets has been allotted to accommodate an optical 

communications technology demonstration. Compared to traditional RF communication, optical can 

deliver higher data rates, up to 1 Gbps, at a lower mass and power (Ref. 14). Optical communication 

technology is capable of communicating at high bandwidth and at distances up to 40 astronomical units. 

These communication links could enable the next generation of exploration missions. The Lunar Laser 

Communication Demonstration aboard LADEE, to be launched in 2013, is being developed to 

demonstrate an optical 622 Mbps downlink, and up to 20 Mbps uplink (Ref. 15). The NASA Jet 

Propulsion Laboratory (JPL) is also developing optical communication technology for missions to the 

Moon and outer planets (Ref. 14). Given that JPL’s technology is not being developed to fly a particular 

mission, SCHOLR could be available to demonstrate this system and help further this revolutionary 

communication technology.

4.2 Communications Payload

4.2.1 Link Requirements

In order to provide relay capabilities as functional as DTE access, it was determined that the LCP 

should provide three different links to the surface. One link should receive scientific data at a high rate. 

The other two links would provide real-time Tracking, Telemetry, and Command (TT&C) at low data 

rates. All three of these links must minimize the number of errors in the transmission in order to ensure 

the data is of high quality and that commands can be processed correctly. To satisfy these requirements, it 

was determined that all three links must exhibit a bit error rate (BER) of 10–8 or less, with a link margin of 

3 dB or more. The low BER reduces the probability that an error will occur during the transmission, and 

the link margin allows for unexpected and uncalculated losses in the communications system (Ref. 6).

4.2.2 Surface Asset Assumptions

It is important to have reasonable assumptions about the capabilities of the surface assets’ 

communications systems in order to appropriately design the LCP and the parameters of each link. To 

make these assumptions, various other lunar communication packages, both conceptual and flown, were 

researched (Refs. 16 to 18). The assumptions based on that research are summarized in Table 3.

TABLE 3.—SURFACE ASSET ASSUMPTIONS

Property Assumed value

Antenna type .........................................................High gain parabolic

Antenna diameter..........................................................0.3 m or larger

Transmit power ................................................................ 5 W or more

Receiver system noise temperature ................................. 290 K or less

Modulation..................................................... Any SCHOLR supports

Coding............................................................ Any SCHOLR supports
NASA/TM—2011-216894 9

Figure 4.—LCP communications architecture. Yellow represents 

the command link; Red, telemetry and tracking link; Green, high 

data rate science link.

The data rates for each link were determined based on the expected types of instruments on the lunar 

surface assets and the amount of data these instruments would generate. It was determined that the high 

data rate link should support rates up to 100 Mbps, accommodating both instruments that generate large 

volumes of data quickly and instruments that generate small volumes of data but store a large amount. 

Low data rate TT&C links were determined to require 2 Mbps to support real-time monitoring and 

control. S-band was found to be the best option for TT&C links and Ka-band would support the high data 

rate link. S-band is well developed and would allow maximum usage of commercial off the shelf (COTS) 

packages, minimizing cost and facilitating greater interoperability between the LCP and the lunar surface 

assets. Ka-band was chosen for the high data rate link because it has been proven to support data rates up 

to 100 Mbps by the LRO16. Figure 4 illustrates the communications architecture on the lunar surface. A 

summary of the parameters for each link is shown in Figure 5.
NASA/TM—2011-216894 10

4.2.3 Link Design

Using these parameters, a model of the scenario was created using an analysis software tool called the 

Satellite Tool Kit (STK) (Ref. 19). Using STK, link properties were explored and data was collected for 

the different options available in the LCP’s design. Quadrature phase shift keying (QPSK) modulation 

with low-density parity-check (LDPC) coding was determined to be the best combination of modulation 

and coding because it had the lowest required signal-to-noise ratio (Eb/N0) necessary to meet the BER 

requirement (Ref. 20).

The Equivalent Isotropic Radiated Power (EIRP) necessary to close the command link was first 

determined in order to minimize the number of unknown factors while designing the LCP. To determine 

the necessary EIRP, access time was used as a figure of merit because it would indicate the amount of 

time the link was closed. It was assumed that more access time was better because it would allow for 

more surface assets to be serviced and each asset to be serviced longer. The transmit power and antenna 

diameter were sized using the determined EIRP of 25 dBW. A plot of the relationship between transmit 

power and antenna diameter, assuming 55 percent antenna efficiency, is shown in Figure 6 (Ref. 21).

Based on this relationship it was determined that a 0.5 m antenna with 4.8 W of transmit power would be 

able to meet the requirements for the command link while remaining within constraints for power, size, 

and mass.

It was then verified that a 0.5 m antenna required an acceptable system noise temperature to meet the 

required gain over system noise temperature, G/T. A plot of system noise temperature vs. antenna 

diameter for both links is shown in Figure 7. With a 0.5 m antenna, the Ka-band receiver requires a 

system noise temperature no greater than 322 K and the S-band receiver requires no greater than 290 K. 

Because 290 K is the standard system noise temperature (Ref. 21), it was determined that a 0.5 m antenna 

would allow the LCP to meet the requirements for all three links.

4.2.4 Final LCP Design

To validate that the selected parameters for each link provided adequate coverage, STK was used to 

generate coverage maps for each link. The three coverage maps are shown in Figures 8(a), (b), and (c).

The coverage maps show the relative amounts of access time for different areas of the lunar surface. 

The colors on each map show the total percentage of time that the area has communication access during 

the 3 yr mission. A darker color indicates more access time. The maps reveal that both the far side and 

near side of the southern hemisphere can be provided relay communications and that the access time in

this region increases the closer a surface asset is placed to the South Pole. Both of these characteristics are 
NASA/TM—2011-216894 11

desirable, as the majority of the science performed by surface assets would be done close to the poles, 

while the capability to support surface assets anywhere in the southern hemisphere also exists.

The finalized parameters are shown in a block diagram describing the entire LCP, Figure 9. 
NASA/TM—2011-216894 12

By utilizing different frequencies and an RF matrix switch for all three of the links, each 

communication link can use the same 0.5 m high gain parabolic antenna. The S-band transmitter uses a 

solid-state amplifier to produce the transmit power needed to close the links. The two receivers each use a 

low noise amplifier in order to pick up the signals from the lunar surface asset. The data entering and 

leaving the LCP is sent by the C&DH system to a router that decides which path the information should 

take to reach the correct place. The final design parameters are summarized in Table 4.

TABLE 4.—LCP LINK PARAMETERS

Parameters S-band receive S-band transmit Ka-band receive

Frequency, GHz 2.25 2.1 26 

Transmit power, W 5 4.8 5 

Polarization RHCP RHCP RHCP

Transmit antenna gain, dBi 14.4 18.2 35.6 

EIRP, dBW 21.4 25 42.6 

Receive antenna gain, dBi 18.8 13.8 40.1 

System noise temperature 290 290 322

G/T, dB/K –5.8 –10.8 15 

Data rate, Mbps 2 2 100 

Required Eb/N0, dB 2.6 2.6 2.6 

Required Eb/N0 with link margin, dB 5.6 5.6 5.6 

The LCP is able to provide the three links necessary to allow lunar surface assets to communicate with 

SCHOLR. Simulation of the scenario in STK confirms the SCHOLR LCP could provide service to lunar 

surface elements while meeting all of the link requirements.

4.3 Mission Orbit Determination

SCHOLR’s orbit is an integral contributor to meeting stakeholder requirements as it defines what 

functions SCHOLR will be capable of performing. It affects the operation of both the SIP and the LCP. 

For both mission objectives, scientific data gathering and communications relay, SCHOLR requires an 

orbit which passes over the lunar poles. The resolution of the sensors in the SIP is directly related to the 

altitude at which the measurements are taken. Each sensor chosen to fly aboard SCHOLR requires an 

orbit altitude near 100 km. A communications relay needs a relatively long access time (i.e. time during 

which SCHOLR is in sight of the surface elements) to be effective. This implies that SCHOLR must be at 

a high altitude within the surface elements’ line-of-sight to the LCP.
NASA/TM—2011-216894 13

There is a major challenge inherent in designing an orbit for a hybrid science and communication 

satellite like SCHOLR. The conflicting needs of the two payloads for different altitudes make global 

hybridization practically impossible with current technology. Therefore, a solution is needed that satisfies 

SCHOLR’s requirement to be a hybrid orbiter. That solution is temporally isolated hybridization, in 

which SCHOLR performs both data gathering and communications relaying, but does so within separate 

portions of the orbit.

4.3.1 Orbit Trades

Over the course of the orbit determination process, three main trades were investigated. The trade 

selected for SCHOLR, a highly elliptical polar orbit (Trade 3, cyan orbit in Fig. 10), was determined to be 

the option best suited for the mission. It combines a low periapsis altitude at the lunar North Pole and a 

high apoapsis altitude at the lunar South Pole. The low periapsis altitude, approximately 100 km, is in 

accord with the range limitations of the SIP. The high apoapsis altitude, approximately 5798 km, enables 

a communication link to be established from SCHOLR to any surface node in the southern hemisphere for 

over 5 hr, a period long enough to forward stored data and facilitate teleoperation of surface elements 

from Earth.

An elliptical lunar frozen orbit (Trade 1 white orbit in Fig. 10) was not feasible because an orbit with 

an inclination of 90° must have an eccentricity of 1.0 to be a frozen orbit (Ref. 22). In that case, the orbit 

would be parabolic: an escape trajectory. Alternatively, if the 90° inclination is sacrificed for a practical 

eccentricity, the orbit no longer passes above the lunar North Pole.

Although the circular path and low 100 km altitude of the Trade 2 orbit, seen in orange in Figure 10, 

allowed for the best range and pass duration for the SIP, it also severely limited communication 

applications. SCHOLR could initially fly this 100 km orbit and boost to a higher orbit after the 3 yr of its 

primary mission. This operation could provide better relay capabilities during SCHOLR’s extended 

mission. However, the delta-V required to capture into low lunar orbit and later move to a higher orbit, 

nearly 1.5 km/s in total (compared to only 500 m/s for capture into the Trade 3 orbit), made the option 

undesirable.
NASA/TM—2011-216894 14

The orbit finally selected for the SCHOLR primary mission, summarized in Table 5, was an 8-hr

elliptical polar orbit. The 8 hr period was selected to enable the fulfillment of both science and 

communication goals and to complement the work schedules of ground crews supporting the SCHOLR 

mission. 

TABLE 5.—NOMINAL SCHOLR MISSION ORBIT ELEMENTS

Semimajor axis ............................................................... 4687 km

Eccentricity ........................................................................ 0.6078

Inclination................................................................................90°

Arg. of periapsis.......................................................................90°

Periapsis altitude ............................................................... 100 km

Apoapsis altitude............................................................. 5798 km

Period......................................................................................8 hr

4.3.2 Orbit Decay Analysis

In order to study mission lifetime and provide data for propellant usage analysis for SCHOLR, the 

orbit was modeled using STK (Ref. 19) and Satellite Orbit Analysis Program (Ref. 23) with third body 

perturbations to simulate orbit degradation due to forces not accounted for in traditional two body 

mechanics. The simulated orbit was permitted to decay until it intersected the lunar surface, implying 

SCHOLR would impact the Moon, as seen in Figure 11. Working back from that point, it was 

determined, based on standard practices and the simulation results, that a station-keeping maneuver would 

be performed every time the periapsis altitude reached 50 km. Based on this assumption, a burn of 5.6 m/s 

delta-V would be required approximately every fourteen days. Therefore, over the nominal 3 yr SCHOLR 

mission, a delta-V totaling 437 m/s would be required to keep the satellite in orbit around the Moon. An 

additional 952 m/s would be needed for launch vehicle injection for lunar transfer orbit, mid-course 

corrections, lunar orbit capture, and the maneuver to enter the extended mission orbit (Ref. 24).
NASA/TM—2011-216894 15

4.3.3 Spacecraft Pointing

SCHOLR’s temporally isolated hybridization influenced the configuration of the spacecraft. Due to 

competing altitude requirements, the payloads could not function simultaneously. To enable 

hybridization, SCHOLR was designed to be inertially pointing and configured with the external 

components of the SIP and LCP on opposite sides of the spacecraft bus. With the satellite always oriented 

to the lunar poles, the spacecraft face with the SIP would always face the lunar surface as it passes over 

the North Pole. Similarly, the side with the LCP would always view the lunar surface as it passes over the 

South Pole. This additionally establishes an axis of rotation about which SCHOLR can spin to track Earth 

with its Earth antenna. Continuously pointing the Earth antenna in this manner reduces risk and propellant 

usage over typical operations that involve flipping maneuvers to keep Earth in view (Ref. 6).

5.0 SCHOLR Subsystem Design

5.1 Subsystem Integration

A conceptual design for each of SCHOLR’s subsystems was performed after the mission architecture, 

payloads, and orbit were determined. To ensure that the subsystems would function both independently 

and collectively to accomplish SCHOLR’s mission, the mass, dimensions, power consumption, and 

pointing needs for each subsystem were tracked throughout the design phase. If any subsystem’s demands 

conflicted with another subsystem’s limitations, the integration process mediated discussions to strike a 

compromise. In this way, the integration process prevented subsystem incompatibility, thus supporting 

the success of the entire satellite system.

The primary means of subsystem organization and analysis was the Master Equipment List (MEL). As 

subsystem was specified, component details were entered and updated in the MEL. A condensed view of 

the MEL can be found in Table 6. 

TABLE 6.—SCHOLR OVERVIEW SNAPSHOT OF THE MEL.

[Note that only one of eight power configurations is displayed.]

Description

SCHOLR Communications Spacecraft

Basic mass

(kg)

Growth

(%)

Growth

(kg)

Total mass

(kg)

Power 

mode 3

(W)

SCHOLR Spacecraft 1870.73 8.8 164.95 2035.67 1020

Science Payload 74.20 28.2 20.95 95.15 230.4

Attitude Determination and Control 40.52 20.0 8.10 48.62 207.0

Command and Data Handling 59.40 23.7 14.09 73.49 163.3

Communications and Tracking 92.51 10.0 9.25 101.77 305.3

Electrical Power Subsystem 86.00 25.8 22.20 108.20 0.0

Thermal Control (Non-Propellant) 197.47 15.2 30.10 227.56 13.0

Propellant (Chemical) 874.91 0.0 0.00 874.91 0.0

Structures and Mechanisms 362.38 14.1 51.09 413.46 0.0

The MEL maintained characteristics such as mass, growth percentage, power consumption, operating 

temperatures, and basic dimensions. These measures provided a general overview of the satellite design 

and indicated violation of any critical restrictions, such as total mass or total power consumed. The 

utilization of the MEL during the design process allowed for quick and collaborative design and analysis 

of the satellite.

5.2 Structural Subsystem

The structures of SCHOLR must support and protect the spacecraft’s components from dynamic 

environments during all mission phases: pre-launch, launch, deployment, operations, and disposal. The 

key elements that drive the structures of the spacecraft include the packaging concepts, deployable 
NASA/TM—2011-216894 16

structures, materials, and volume to ultimately keep the payload safe and functional. These decisions are 

made based upon the mission’s goals, requirements, constraints, and payload configuration drivers.

5.2.1 Subsystem Requirements

Each component must be in a position that allows it to perform properly and satisfy mission 

requirements. The structures subsystem must support the spacecraft’s components in these desirable 

locations. SCHOLR must additionally deploy any necessary components while providing enough 

stiffness to keep them steady. This would allow the deployed components to achieve optimal 

performance. The structures must also provide sufficient support for all loads and vibrations during 

mission phases. It must prevent the spacecraft from collapsing, damping, or affecting the components. 

Furthermore, the structure material must protect each component from radiation, pressure, impacting 

particles, and thermal cycling (Ref. 25).

Structure requirements were derived from the mission requirements and in consideration of the types 

of spacecraft components that would require support. The general shape and dimensions were estimated 

for the main bus structure, as seen in Figure 12. A packaging configuration of all the components within 

the spacecraft was then developed. This consisted of identifying subsystem requirements for size, fields of 

view (FOV), mechanisms, and deployable structures. Finally, the material for each structure was selected.
NASA/TM—2011-216894 17

5.2.2 Structure Design

The type of structure for the main bus first determined SCHOLR’s structural system. A thrust tube 

design with inner panels, seen in Figure 13, was chosen for simplicity, strength, and adaptability to the 

launch vehicle payload adaptor. The thrust tube would house the main fuel tank, pressurant tank, and 

main thruster. The material was chosen to be aluminum lithium alloy, specifically aluminum 2090-T83. 

This has a high strength-to-weight ratio, high technology readiness level (TRL), and low cost. 

Honeycomb sandwich panels would be used on the inner panels connected to the thrust tube and the outer 

side panels for mounting components. Honeycomb panels were chosen for mounting components because 

they provide a high bending strength and stiffness for a low mass (Ref. 25).

Two booms were required to deploy the radar sounder 15-m in opposite directions. ABLE 

Engineering coilable booms were chosen because of their high reliability, strength, and TRL; low weight, 

and small-stowed size. A continuous-longeron boom was chosen because the boom was not restricted to a 

certain diameter. A continuous-longeron boom also provides high dimensional stability and a high 

stiffness-to-weight ratio. A lanyard deployment mechanism was chosen over a canister deployment 

mechanism because it weighs less and the stowage volume is smaller in length and diameter (Ref. 26).

5.2.3 Configuration Design

Characteristics such as location, mass properties, size, mechanical interfaces, fields of view, and 

thermal interfaces for all components must be considered to determine the best layout for the spacecraft 

and its components. This entails integrating each subsystem’s components to create the design that allows 

maximum performance of the spacecraft. The main design drivers for the configuration are the propulsion 

tank sizes and the fields of view for the science instruments, antennas, and solar array (Ref. 27).

The configuration was derived from the subsystem requirements and constraints. Design parameters 

and a preliminary spacecraft design were then estimated based on these subsystem concepts. The Falcon 9 

payload fairing, shown in Figure 14, drove the size of SCHOLR. Budgets for quantity, size, and mass of 

each component were then established in the MEL. A structural architecture and design for packaging the 

components was then selected and created in the computer-aided design software, SolidWorks (Ref. 28).
NASA/TM—2011-216894 18

5.2.3.1 Configuration Constraints

The location of the outer components of the SIP and LCP were constrained by their pointing 

requirements. The SIP should face the Moon when SCHOLR is at the North Pole, and the LCP should 

face the Moon when SCHOLR is at the South Pole. SCHOLR’s inertial pointing constrained the 

configuration of the external components because different spacecraft faces view the Moon at different 

parts of the orbit. This directly affected the FOV for specific components.

Another constraint was the available area to place external components. For example, it was 

necessary for the radiators to be a certain size to adequately control the spacecraft temperature. A conflict 

arose because of the noise introduced to the communications link by radiators located near antennas. The 

lunar antenna would be deployed on a boom to avoid the FOV of the radiators, and the Earth antenna 

would be gimbaled to avoid thermal radiation. Similar constraints exist with the solar array, SIP, LCP, 

radiators, guidance sensors, and propulsion. 

SCHOLR’s internal configuration was heavily influenced by the size of the propulsion tanks, the 

heaviest and largest components. Other components, whose volumes were tracked in the MEL, must be 

secured to the spacecraft bus and fit within the spacecraft’s skin. Component locations were also driven 

by thermal interfaces from the thrusters, battery, and radiators. Optimal functionality considerations 

shaped the internal placement of components, heat pipes, and electrical and data connections.

5.2.4 Final Configuration

SCHOLR’s final deployed configuration is shown in Figure 15. The stowed configuration design is 

shown in Figure 16 and the internal configuration can be seen in Figure 17. The FOV is sufficient for the 

mission objectives of the antennas, solar array, star trackers, and scientific instruments. The science 

instrument locations are shown in Figure 18. The radar sounder was placed in the center of the side panel 

to help balance the mass distribution. The UltraFlex solar array, Earth and lunar antennas, and coilable 

booms for the radar sounder were all analyzed to best determine the stowed and deployed configurations 

of the spacecraft. There are two Sun sensors placed on each side of the craft to provide full omnidirectional coverage. This allows the sensors to have a constant FOV of the Sun.

The subsystems are grouped together within the spacecraft to ensure simplicity during assembly. 

There is adequate space left around each component for wiring and installations. Doors are provided on 

the side panels to provide technicians access to the components during assembly. Each component is

packaged in a location that would allow it to best perform its mission objectives. Every component would

function within its operating temperature.

The C&DH system was placed near the center of the spacecraft because it is the interface between all 

the subsystems. The main thruster was positioned through the spacecraft’s center of mass to reduce the 

complexity of attitude control. Table 7 shows the final design details.
NASA/TM—2011-216894 19

TABLE 7.—FINAL STRUCTURE AND 

CONFIGURATION DESIGN DETAILS

Total SCHOLR mass.......................................................1886.35 kg

Total SCHOLR mass with growth...................................2039.05 kg

Max load.......................................................................................6 g

Total structure mass with growth.......................................413.49 kg

Main bus..............................................................................238 kg

Two 15 m booms................................................................8.88 kg

Installation mass ....................................................................95 kg

Design load.......................................................................110,917 N 

Allowable load..................................................................119,896 N

Safety factor..................................................................................1.4

Limit load .........................................................................167,855 N

Aluminum lithium alloy (Al 2090-T83) honeycomb sandwich panels

Two 15 m continuous-longeron booms with lanyard deployment
NASA/TM—2011-216894 20

5.3 Environmental Control Subsystem

The SCHOLR Environmental Control Subsystem (ECS) ensures the spacecraft temperature will 

remain within the operational limits of the on-board instruments and that the spacecraft will safely operate 

within the lunar orbit environment. The main element of the ECS, the thermal control subsystem, rejects 

excess heat or generates required heat according to the thermal operational limits of the spacecraft 

components. The rest of the ECS shields the spacecraft from the lunar orbit environment, including 

unwanted radiation and micro meteors, through use of a variety of passive systems.

5.3.1 Thermal Subsystem Design

5.3.1.1 Thermal Requirements

The thermal subsystem must maintain the spacecraft components and structural interfaces within their

operating temperature ranges during all mission phases. To accomplish this, the thermal subsystem must 

provide a means of cooling the spacecraft during operation as well as provide heat to vital components 

and systems to preserve their functionality. The maximum heat load to be rejected by the thermal system 

is 1312 Watts thermal (Wt), and the desired operating temperature for the radiators is 320 K.

5.3.1.2 Thermal Assumptions

The thermal modeling process provides mass and power estimates for the various aspects of the 

thermal control system based on a number of inputs related to the vehicle geometry, flight environment 

and component size. The assumptions utilized in the analysis and sizing of the thermal subsystem were 

based on the operational environment of the spacecraft. Since SCHOLR would operate in a highly 

elliptical polar lunar orbit with a periapsis of 100 km, the following assumptions were made to size the 

thermal subsystem:

• The view factors for the radiator to the Earth, lunar surface and solar array were assumed to be 

0.1, 0.35 and 0.1 respectively

• The maximum angle of the radiator to the Sun was 30°

• The direct energy transfer (DET) from the solar panels to the instruments is 85 percent efficient 

during times of insolation and 65 percent efficient during eclipses, when batteries supply power

(Ref. 29)

• Typical parabolic antennas convert 55 percent of the power provided to them into RF energy that 

carries the communications signal to Earth and therefore does not need to be rejected through the 

radiators (Ref. 30)

• The radiation heat transfer equation, based on the Stefan Boltzmann law, was used to size the 

radiators (Ref. 31)

5.3.1.3 Radiator Design

The primary component of the thermal subsystem is the radiator, which regulates the spacecraft 

temperature by rejecting excess heat. Most of the heat that must be controlled or rejected is produced 

when electrical power from the solar array is converted to heat due to the internal resistance of spacecraft 

instrumentation. Both waste heat from the satellite and environmental radiation are rejected by infrared 

radiation from the surface of the radiators.

The thermal power to be rejected from the craft was determined by the power consumed by the 

spacecraft during each stage of the mission. Considering each of these factors, the maximum amount of 

power to be rejected at any time is 1312 Wt, as seen in Figure 19. 

The radiating temperature was determined by the desired operating temperature of the internal 

instruments. The sink temperature of 260 K was estimated using information about the environment 

during SCHOLR’s mission. A reflective coating with a high infrared emissivity (0.84) was chosen to 

cover the radiators to protect them from environmental radiation. 
NASA/TM—2011-216894 21

The radiator for SCHOLR was sized by solving the heat transfer equation for the radiator area required

to reject the excess heat generated within the spacecraft. These calculations returned a required radiator 

area of 4.7 m2

, which is too large to fit on the given 3.0 m2 side of the spacecraft. Therefore, trade studies 

were performed to determine the best way to accommodate the large amount of thermal power that must 

be rejected from the craft. 

5.3.1.4 Radiator Trades

Multiple trades were considered for the radiator design for SCHOLR. Design ideas investigated 

included: the use of louvers in addition to a radiator, a deployable radiator, two radiators radiating at the 

same temperature, and one radiator using phase change material (PCM). It was determined that louvers 

should not be used because they obscure the radiator’s view to deep space, increasing the required 

radiator area by 30 percent. Although using a deployable radiator would provide twice the radiating area 

for the same effective stowed size, it was deemed to be significantly too complicated and costly for the 

spacecraft and therefore was not implemented. A trade study was performed on the use of PCM to 

determine if this method would sufficiently reduce the radiator area, allowing the radiator to fit on one 

side of the spacecraft. 

The goal of the PCM study was to determine the mass of PCM that would be required to allow the 

use of a single radiator. The maximum power that could be rejected from a radiator constrained to one 

side of the spacecraft was calculated to be 775 Wt. Figure 19 shows the excess power to be dissipated 

using PCM if only one radiator were to be implemented. Since the original power to be dissipated was 

1312 Wt, the PCM would have to absorb 537 Wt. Assuming a 3-hr dissipation time due to SCHOLR’s 

orbit and a PCM heat of fusion of 215 kJ/kg, dimensional analysis determined 27 kg of PCM was

required. An additional 8 kg of structure supports the PCM, increasing the total mass to 35 kg. Based on 

the mass constraints imposed by SCHOLR’s Falcon 9 launch vehicle, the PCM could not be used to 

decrease the radiator area because of the required increase in mass.
NASA/TM—2011-216894 22

5.3.2 Final Environmental Control Design

SCHOLR would incorporate two 2.3 m2 radiators that radiate at the same temperature. Improved 

radiator efficiency can be realized by maximizing spacecraft pointing at deep space and minimize 

pointing at celestial bodies such as the Moon, Earth, and Sun. Therefore, the radiators are mounted on 

opposing sides of the spacecraft as shown in Figure 20.

Other components of the ECS include micrometeor shielding, which protects the junction between the 

heat pipes and radiators and multi-layer insulation to insulate the craft, enabling accurate thermal control 

and external coatings to reflect excess heat and radiation. Lastly, electric heaters and a control system 

were also selected for the subsystem management. SCHOLR’s ECS consists of active and passive 

controls as part of its design, as shown in Table 8.

TABLE 8.—ACTIVE AND PASSIVE CONTROLS

Active thermal control Passive thermal control

Variable heat pipes Radiators

Thermal control circuit Heat sinks

Data acquisition Multi-layer insulation

Thermocouples Micro meteor shielding

Electric heaters Reflective surface paints

5.4 Guidance, Navigation, and Control Subsystem 

The purpose of SCHOLR’s Guidance, Navigation, and Control (GN&C) subsystem is to locate, point, 

and rotate the spacecraft to complete its mission. The GN&C subsystem provides attitude control of the 

vehicle from launch through end of mission in all three axes.

5.4.1 GN&C Requirements and Assumptions

The attitude control system is responsible for maintaining SCHOLR’s attitude, solar array orientation, 

and high gain antennas’ pointing throughout the mission. The attitude control system shall correct for 

disturbance torques and would supply housekeeping, telemetry, attitude, and orbit determination data to 

the ground station.The GN&C subsystem also controls and monitors the propulsion system. SCHOLR’s 

products of inertia were assumed to be zero, meaning it was assumed the spacecraft’s mass distribution 

was symmetric about the x, y, and z axes. SCHOLR has negligible slew requirements because of the 

inertial pointing designed for the orbit. SCHOLR’s altitude of 100 km at periapsis and 5798 km at 

apoapsis factors into the calculations to size the subsystem’s instruments. 

5.4.2 Disturbance Torque Considerations

One of the major roles of the GN&C subsystem is to ensure the craft maintains the correct attitude 

during each phase of the mission. Disturbance torque affects the attitude of the satellite by varying 
NASA/TM—2011-216894 23

amounts at different points throughout the mission. Two types of disturbance torques were analyzed: 

gravity gradient and solar radiation torques. Magnetic field torques and aerodynamic torques were not 

considered for SCHOLR’s lunar orbit because there is not a significant magnetic field or an atmosphere 

associated with the Moon to cause a disturbance during the mission.

5.4.2.1 Gravity Gradient Torque

Gravity gradient torque is caused by the difference in magnitude of the gravitational forces on the ends 

of the craft. Gravity gradient torque reduces by a factor of 1/r

3 as the orbit radius increases. The portion of 

the spacecraft that is closest to the Moon experiences a greater gravitational force than the portion farthest 

from the Moon, thus creating a torque on the spacecraft. This torque will rotate the spacecraft to align the 

minimum moment of inertia with the local vertical (Ref. 32). The more cubical a satellite is, the smaller 

the gravity gradient torque. 

The amount of gravity gradient torque SCHOLR would experience was calculated as a function of the 

difference in the moments of inertia of the spacecraft. The estimated worst case scenario gravity gradient 

torque SCHOLR would experience is 0.00024 Nm. The momentum due to the gravity gradient torque is a 

cyclic torque and therefore does not build up, causing the need for momentum dumps.

5.4.2.2 Solar Radiation Torque

Solar radiation torque is a function of the craft’s cross sectional area exposed to the Sun, the craft’s 

absorption and emissivity properties, the distance between the center of mass and center of pressure, and 

the solar constant. A measure of the flux from incoming solar electromagnetic radiation, the solar 

constant is dependent on the distance from the Sun, which is 1367 W/m2 near the Moon (Ref. 29). Based 

on these properties, SCHOLR would experience solar radiation torque of about 0.000039 Nm. The 

momentum due to the solar radiation torque would build up over time because it is a secular torque.

5.4.3 Final GN&C Design

To counteract these torques and maintain the desired attitude, reaction wheels were implemented to 

stabilize the craft. The total amount of torque was used to determine the required momentum storage and 

wheel torque necessary to size the reaction wheels for SCHOLR. A torque margin was incorporated when 

choosing the reaction wheels to account for the torque generated when the antennas, solar array, and 

stereo camera would be gimbaled. Based on the momentum storage of the wheels sized for SCHOLR 

(20 Nms for each wheel) and the cyclic torque acting on the spacecraft (0.000039 Nm of solar radiation 

torque), the wheels would be able to store the momentum for 512820 sec, or about 6 days before being

dumped. 

The GN&C components and key aspects of their design are listed in Table 9. All components for 

SCHOLR’s GN&C subsystem have a long heritage and are space flight qualified. Using COTS products 

reduces the risk associated with the mission. Figure 21 depicts the components of the GN&C subsystem 

and how they connect to the propulsion system through the C&DH subsystem. The GN&C software runs 

on the main C&DH computers.

5.4.4 Propulsion Subsystem

The SCHOLR propulsion subsystem enables the spacecraft to perform its mission by sending the 

craft where it needs to go, maintaining its attitude, and preventing SCHOLR from mission-ending 

collisions with the lunar surface. This subsystem is single fault tolerant where appropriate.
NASA/TM—2011-216894 24

TABLE 9.—GN&C COMPONENTS

Instrument Picture Function Key 

characteristics

Qty Mass 

(kg)

Inertial 

Measurement 

Unit

(Northrop 

Grumman SIRU)

• Used in orientation and 

location determination

• Added accelerometers also allow 

for position and velocity sensing

• Uses multiple gyroscopes on 

different axes

• Internally redundant

1 7.10

Reaction Wheels 

(Bradford W45)

• Attitude control and fine 

pointing

• Use torque motors and 

high-inertia rotors for 

momentum storage

(Ref. 33)

• 20 Nms of momentum storage

• Sized based off disturbance 

torques experienced during the 

mission

• Pyramid configuration allows 

control over all three axes and 

redundancy

4 6.95

each

Star Trackers

(Selex-Galileo 

A-STR)

• Identify star patterns and 

determine attitude based 

on star crossings

• Require relative stability to 

determine inertial reference

• Error introduced due to sunlight 

interference

2 3.00 

each

Sun Sensors

(Adcole SASS 

15671)

• Used in attitude 

determination and solar 

array orientation

• 12 units, two on each face

• Measures angle between 

mounting base and incident 

sunlight (Ref. 34)

• Require clear fields of view

12 0.32 

each

5.4.4.1 Propellant Choices

When seeking out innovative options for SCHOLR’s propulsion system, green propellants were 

proposed as a possible solution as they are an emerging technology in-space propulsion. The first option 

considered was a fuel developed by Ecological Advanced Propulsion Systems, a subsidiary of the 

Swedish Space Corporation. High Performance Green Propellant is a monopropellant designated as LMP103S, composed of ammonium dinitramide (NH4N(NO2)2), water, methanol, and ammonia. This fuel has 

both a higher specific impulse and higher density than hydrazine, causing it to be a desirable design 

choice for satellite designers who are focused on performance as well as environmental impact (Ref. 35).

While this fuel and these engines are space flight qualified, and the fuel is easier and safer to handle than 
NASA/TM—2011-216894 25

hydrazine, the engines available are not large enough for the SCHOLR mission. Thus, this propulsion 

system was not chosen for SCHOLR. 

Much activity is occurring with regards to hydrogen peroxide rocket propulsion research. This green 

propellant uses simple chemistry, and General Kinetics will soon have large engines that are flight 

qualified. Tests are underway with 250 and 1500 lbf rockets commercially available. Unfortunately, the 

necessary engines are not yet available. This system was not chosen for SCHOLR due to unavailability of

COTS parts.

Hydrazine is a heritage propellant in the world of satellite design. COTS parts are readily available 

through Aerojet and other companies; it is well researched, and is used on the Space Shuttle for attitude 

control. Due to the COTS design requirement, a hydrazine system was chosen as engines and parts are 

available in the necessary sizes for SCHOLR.

5.4.4.2 Final Design

First, an ATK 80507-3 was chosen as the main tank for the SCHOLR spacecraft. It was the 

appropriate size to hold the propellant necessary for the 1.5 km/s delta-V predicted by the mission. As 

helium tends to leak due to its small atomic size, nitrogen was chosen as the pressurant. National Institute 

of Standards and Time data on nitrogen characteristics was used to calculate the mass of pressurant 

needed and to assist in sizing the pressurant tank. To verify this calculation, a tank was selected that could 

hold the entire mass of nitrogen at very high pressures. The ATK 80314-1 was selected as the pressurant 

tank. Due to the mission’s thrust/impulsive burn requirements, a large engine was selected as the main 

engine. Following the Messenger mission’s precedent, SCHOLR will carry only one main engine, the 

100 lbf Aerojet MR-104, which enables both mass and cost savings. Smaller engines were needed for 

attitude control, and due to its utility in other designs (Ref. 6), the 0.5 lbf Aerojet MR-111E, shown in 

Figure 22, was selected to fly in four pods of three. Figure 23 depicts how these thrusters are arranged so 

that thrust can be sent in any of the three axes of the satellite. An internal piping diameter of 1.3 cm and 

length of 2 m was assumed for the main engine assembly. An internal piping diameter of 0.6 cm and 

length of 4 m was assumed for the reaction control system assembly. Calculations were done using actual 

COTS parts and data for valves, fill drain valves, filters, MLI, and pressure regulators. For heaters and 

sensors, simple assumptions were made to account for their mass and power usage. Table 10 shows the 

propulsion and propellant MEL.
NASA/TM—2011-216894 26

TABLE 10.—PROPULSION AND PROPELLANT MASTER EQUIPMENT LIST

Description

SCHOLR Communications Spacecraft

June 2010

Qty Total 

mass

(kg)

Power

(W)

Shape

box, cylinder

Dim

length

(cm)

Dim 

width

(cm)

Propulsion (chemical) 92.51 197.6

Propulsion hardware (chemical) 1.90 13.1

Main engine 1.90 13.1

Main engine 1 1.90 13.1 0 0.46 0.15

Propellant management (chemical) 86.65 20.9

Fuel tanks 1 44.70 10.0 Cylindrical tank with 

spherical domes

49 ID 45.6

Pressurization system—tanks, valves, etc 1 25.33 10.3 Spherical 0 ID 16.4

Feed system—regulators, valves, etc 1 16.62 0.6 0 0 0

RCS hardware 3.96 163.7

RCS thruster subassembly 12 3.96 163.7 0 0.17 0.36

Propellant (chemical) 874.91 0.0

Main engine propellant 865.75 0.0

Fuel 865.75 0.0

Fuel usable 1 785.84 0.0

Fuel margin (main and RCS) 1 50.64 0.0

Fuel residuals (unused) 1 29.28 0.0

Pressurant 9.16 0.0

Main pressurant 1 9.16 0.0
NASA/TM—2011-216894 27

5.5 Power Subsystem 

5.5.1 Requirements

The Electric Power Subsystem (EPS) is responsible for the generation, storage, and distribution of 

power on the SCHOLR spacecraft. SCHOLR’s 3 yr mission requires a reliable power system that would 

enable continuous operation of the spacecraft subsystems and payloads. During periods of sunlight, the EPS 

would generate enough power to operate the spacecraft and recharge the spacecraft’s batteries. The EPS 

would store sufficient power to ensure spacecraft survival and allow for reduced payload function during 

periods without sunlight. At all phases of the mission, power would be distributed along a regulated 28 V 

power bus and would be converted to usable voltages at each subsystem. 

To size the EPS, a power budget, also known as a Power Equipment List (PEL), and a mission phase 

document were created to estimate the power consumption by the subsystems during each mission phase, 

Figures 24(a) and (b). The PEL was the basis for all decisions about the EPS architecture and component 

sizing calculations and includes a 30 percent growth margin to the final totals in accordance with industry 

standards (Ref. 36). As seen in Figure 24, the spacecraft consumes 1020 W while taking scientific 

measurements over the lunar North Pole and 875 W while relaying information over the lunar southern 

hemisphere. During periods of lunar eclipse, reduced payload operations were assumed to reduce energy 

storage requirements on the EPS.

5.5.2 Generation

Photovoltaic arrays were the clear choice of power generation method due to the abundance of solar 

power available to SCHOLR throughout the 3 yr mission (Ref. 37). Among the choices for photovoltaic 

arrays were rigid panels, SquareRigger, and the UltraFlex solar array. It was determined that the UltraFlex

solar array with triple-junction gallium arsenide cells, shown in Figure 25, was the best choice because it 

has extremely high power-to-weight ratio (Ref. 38).

The UltraFlex arrays have over 15 yr of development at ATK Space Systems, have flown on the Mars 

Phoenix Lander, and were selected as the arrays on the New Millennium Program Space Technology 8 

mission and the Orion Crew Exploration Vehicle (Refs. 39 and 40). The development of the UltraFlex 

array for both manned and unmanned missions would enable the SCHOLR spacecraft to reduce risk by 
NASA/TM—2011-216894 28

relying on heritage hardware. As seen in Figure 26, the solar array would be mounted on the face opposite 

the science instruments and would incorporate a single-axis gimbal to track the Sun. After selecting the 

UltraFlex array, the next step was to determine the appropriate solar array size.

Using traditional power system sizing methodologies (Ref. 41), orbital information, and power 

information from the PEL, a mathematical model of the EPS was created. To calculate the size of the 

solar array necessary to generate sufficient power, the model assumed the following:

• State-of-the-art triple-junction GaAs photovoltaic cells achieve 28 percent efficiency at beginning 

of life (Ref. 38)

• An 8 hr elliptical polar orbit with a worst case eclipse duration of 3 hr

• Six percent loss in power supplied due to aging effects of the UltraFlex solar array and 

1.5 percent annual degradation due to environmental factors (Ref. 38)

• The power architecture is 85 percent efficient from the solar panels to the instruments during 

times of insolation and 65 percent efficient during eclipses, when batteries supply power (Ref. 29)

• Less than 15° of cosine losses due to the inertial pointing of the spacecraft and the Sun-tracking 

accuracy of the array

Based on these assumptions, it was calculated that a 2 kW photovoltaic array was able to meet 

SCHOLR’s power generation requirements. The solar array area required for the 3 yr mission is 7.0 m2

. 

To enable a 4 yr extended mission, the solar array area should be enlarged to 7.4 m2 to account for the 

extra environmental degradation of the array.
NASA/TM—2011-216894 29

5.5.3 Storage

The current best technology for long-term energy storage in space is lithium-ion batteries due to its 

high specific energy and its ability to withstand the many charge and discharge cycles of a long-term 

space mission (Ref. 42). Assuming the specific energy of the Li-ion battery pack is 110 Wh/kg, the peak 

spacecraft power consumption during an eclipse is 875 W, and a worst case eclipse duration of 3 hr, it 

was calculated that a 4.4 kW-hr battery pack would be necessary to operate the spacecraft through 

periods of lunar eclipse (Ref. 43). This battery pack is oversized so that the battery is never greater than 

60 percent discharged, a condition that would damage the battery pack’s ability to store charge near the 

end of the 3 yr mission (Ref. 44). If the power load can be reduced during eclipse cycles, the size of the 

battery will decrease proportionally.

5.5.4 Distribution

The SCHOLR EPS would utilize a direct energy transfer (DET) architecture to transmit power from 

the solar array to the spacecraft. This calls for a simple and efficient connection between the solar array 

and the spacecraft’s power bus. Excess generated power would be dissipated by a shunt regulator located 

at the solar array (Ref. 45). The power bus would be regulated near 28 VDC by the battery charge 

controller and shunt resistors. The battery discharge controller would limit the bus voltage to slightly 

below 28 V when the spacecraft is in the lunar shadow. While the spacecraft is in the sunlight, shunt 

resistors would limit the bus to slightly over 28 V. A 28 V bus was chosen because flight qualified 

hardware for this voltage has flown on other missions and is readily available (Ref. 46). Most spacecraft 

subsystems and components would not link directly to this regulated bus, but would be buffered by DCDC converters that both protect components from voltage variation in the spacecraft power bus and 

convert the voltage to the appropriate level for that subsystem. This power system architecture, 

represented in Figure 27, completes the EPS design that would meet the requirements of the SCHOLR 

mission.
NASA/TM—2011-216894 30

5.6 Command and Data Handling Subsystem

The Command and Data Handling (C&DH) subsystem of SCHOLR spacecraft is the “brain” 

responsible for receiving and distributing commands, data packaging and storage, and timekeeping

(Ref. 47). The data bus is the “nervous system” responsible for internal spacecraft communications. The 

command aspect of C&DH consists of decoding, validating, and executing commands received from 

Earth. Data handling consists of processing data received from the instruments and payloads, then 

packaging the data into frames, which can be readily transmitted or stored for future transmission

(Ref. 48). The data bus is the interface used to communicate telemetry, commands, and high bandwidth 

scientific data between C&DH and other subsystems. The challenge with designing SCHOLR’s C&DH 

and data bus arises because successful integration of the science and communication missions requires a 

high performing, reliable system.

5.6.1 Requirements

SCHOLR requires an especially capable C&DH subsystem because the unique mixture of science and 

communications payloads would result in a high demand for processing power, data storage, and highrate interconnectivity between subsystems. The C&DH subsystem must integrate the science instrument 

and the relay communications payloads into one data bus architecture, enabling SCHOLR to perform two 

functions on one satellite. The single board computer must perform with computational robustness. The 

data bus must support high data rates (>200 mbps) and low data rates (<1 mbps) between instruments in 

accordance with their bandwidth requirements. Other requirements are relatively standard across all 

C&DH subsystems: it must be space qualified and thermally stable, store science and telemetry data, 

consume low power, and have a real-time operating system (Ref. 43). From these requirements, the

C&DH design process and architecture were conceived.

5.6.2 Design Process

A preliminary C&DH design was created based on the designs of previous missions, especially the 

LRO and Lunar Crater Observation and Sensing Satellite (Ref. 49). From this starting point, a core set of 

technologies was evaluated and trade studies were performed to determine where state-of-the-art 

advances could improve upon the LRO data bus design. Specifications such as mass, cost, power, and 

performance were analyzed, with space qualification being an absolute requirement. The most attention 

was given to trades for data bus architecture and technology as shown in Table 11. 

TABLE 11.—TRADE STUDY FOR SYSTEM ARCHITECTURE (REF. 50) 

Architecture 

type

Characteristics Pros Cons

Centralized Central computer is connected to 

every subsystem

Failure of one system does not 

affect other subsystems

Difficult to add additional 

subsystems to a design

Ring Subsystems connected in series Easy addition of new subsystems Failure of one system may lead to 

failure of entire spacecraft

Federated bus Subsystems connected to each 

other as well as central computer

Direct data paths between all 

subsystems, easy to troubleshoot

Requires subsystems to have 

specific interfaces to communicate 

via a common bus

Distributed bus Federated bus, multiple processors Allows for simultaneous 

command execution

Difficult to test, more complex 

design

5.6.3 Final C&DH Architecture

The final C&DH architecture is based on the centralized concept and is shown in Figure 28. There are 

separate controllers for the environmental control system, housekeeping, Ka-band, S-band, and Optical 

communications alongside the single board computer and solid-state drive. Each of these components 

takes the form of a PCI card inserted into the backplane of a rigid and space qualified enclosure 

responsible for protection and thermal management of the electronics (Ref. 49). The backplane acts as an 
NASA/TM—2011-216894 31

interface for each of the components through which power is distributed and communications between 

cards is transmitted. Central to this configuration is the Single Board Computer, which is based off of a 

RAD750 PowerPC microprocessor from BAE Systems. The RAD750 was selected because it has a long 

flight history, outperforms the similarly priced options (Ref. 51), and is far less expensive than other 

processors (Ref. 52 and 53). Tests performed with LRO’s C&DH system using the RAD750 provide 

assurance that the RAD750 is capable of handling the simultaneous high bandwidth data transmission, 

normal housekeeping operations, and collecting and processing science data loads that the SCHOLR 

spacecraft would encounter (Ref. 49). Since LRO was developed, the RAD750 processor performance has 

increased to 200 MHz, SRAM capacity has increased, and SpaceWire has reached the next level of 

integration with the processor (Ref. 54). These advancements will enhance spacecraft performance. For 

redundancy, each critical component will have a cold backup that can take over upon a hardware failure.

The centralized architecture of the data bus incorporates a central processing unit that is connected to 

every subsystem through a dedicated link. This limits the ability of an instrument failure on one link to 

affect instruments other communications links (Ref. 50). For high data rate instruments, redundant 

SpaceWire links connect to two centralized routers that will each provide communications at up to 

400 mbps (Ref. 55). Low data rate instruments will utilize redundant MIL–STD–1553 links for 

communications up to 1 mbps (Ref. 56). This architecture enables the RAD750 computer to process data 

from both the science and communications payloads, as well as control the other subsystems and 

monitoring their state. 

5.6.4 Atomic Clock

It was determined that SCHOLR would carry a single atomic clock for the purpose of knowing 

precise spacecraft location. While this atomic clock may be used to study the Moon’s gravitational field 

or provide precise metadata for measurements taken by the science payload, the primary purpose is to 

enable the extremely accurate pointing desired by the optical communication transmitters and receivers. 

As optical communication is only a technology demonstration, there would not be a second atomic clock 

for single string redundancy. 
NASA/TM—2011-216894 32

5.7 Communications Subsystem 

SCHOLR’s communications subsystem supports both relay and science data from the two payloads 

as well as its own TT&C. DTE communications is necessary to send lunar surface data received by the 

LCP back to Earth through a bent pipe relay or by using store and forward methods. When orbiting the 

South Pole of the Moon, SCHOLR would serve as a relay for any lunar surface assets and communicate 

back to Earth if SCHOLR is in line of sight with Earth’s ground stations. If there is no line of sight, 

SCHOLR would store the data and forward it when access becomes available. When supporting North 

Pole science missions, data collected by SCHOLR would be transmitted back to Earth at a high data rate. 

The satellite requires a DTE link to enable SCHOLR to serve both its science and communications relay 

objectives while maintaining communications with Earth ground stations when in line of sight.

5.7.1 Requirements

To satisfy mission and science requirements, the communications subsystem and mission node 

communication links must maintain the following requirements:

• Provide a communications relay capability for times when lunar surface assets have poor or nonexistent line-of-sight with Earth

• Provide a high data rate link at 200 Mbps in Ka-band to transmit all science and relay data

• Provide a 4 Mbps and 16 Kbps S-band link to Earth to ensure that TT&C functions are met

• BER of no greater than 10–8 and a link margin of no less than 3 dB to ensure link closure (Ref. 6)

• Support a technology demonstration payload to help develop optical communications in space

The communication links required are illustrated in Figure 29. 
NASA/TM—2011-216894 33

5.7.2 Assumptions

The design of SCHOLR’s DTE communications was sized to the specifications of an 18 m diameter 

antenna located at the White Sands Complex in New Mexico. This ground antenna was previously used 

by LRO and has been used in conceptual lunar relay studies (Ref. 18). The 18 m ground antenna was 

duplicated and placed at each DSN site for link analysis. SCHOLR’s communications system assumed 

this ground network of 18 m dishes would be constructed in the future at DSN sites. 

The worst case scenario for signals was accounted for during the link analysis. The impact of 

atmospheric absorption was considered, which would provide the link with a worst case scenario to allow 

a robust link analysis. The parabolic antenna efficiency was assumed to be 55 percent (Ref. 21). After 

these initial assumptions were made, an initial communications system was developed, including antenna 

size, frequency, and placement on the craft. 

5.7.3 Design

Design features that had to be initially considered were the use of dual feed antennas and overall 

pointing of the craft while in lunar orbit. SCHOLR’s pointing requirements stem from the science 

payload, radiators, as well as communication needs. The antennas must be pointing in the correct 

direction to reduce the use of booms and gimbals and optimize access time to Earth. 

To meet the high data rates required of the DTE link, the antenna design includes a dual feed 

parabolic high gain antenna (HGA) capable of Ka-band and S-band frequencies. Use of the dual feed 

antenna reduces the number of parabolic antennas down to two; a DTE and LCP antenna. The Ka-band 

frequency was determined to be 26 GHz due to the capabilities of the 18 m dish at White Sands (Ref. 18). 

The frequency and link nodes are illustrated in Figure 29.

A range of antenna diameters from 0.2 to 1 m was evaluated to ensure gain was high enough to meet

the link margin and BER requirements. The analysis indicated that a 0.5 m diameter antenna would be 

required to receive and transmit at 200 Mbps in Ka-band. Accounting for the assumed antenna efficiency, 

the gain produced by this frequency and diameter is 40.08 dB (Ref. 57). The parameter used for sizing the 

Ka-band DTE link was EIRP, which was calculated to be 55 dBW or using approximately 30 W of 

transmit power with the calculated gain. 

The high efficiency K-band traveling-wave tube amplifier (TWTA), first used on the LRO (Ref. 58), 

was used to reduce transmit power. The modulation used for all links was QPSK with LDPC coding, 

which provides a coding gain that improves the BER (Ref. 20). The physical aspects of each major 

component of the subsystem are presented in the block diagram in Figure 30.

The medium data rates for the DTE link were determined to require S-band, and the chosen rates for 

transmitting and receiving are illustrated in Figure 29. These data rates are chosen based on the S-band 

capabilities of the 18 m White Sands antenna (Ref. 59). Using the 0.5 m antenna and ensuring that 

transmit power is minimized to remain within the budgeted power, a transmit power of 5 W was assumed

to achieve a gain of 18.8 dB. 5 W of power is the minimum power that can be used to provide enough 

access time to allow all data to be sent back to Earth and still meet link margin and BER requirements. 

The standard system noise temperature that a satellite receives at is 290 Kfor the uplink in S-band

(Ref. 60). Using this temperature as the lower limit, G/T can be determined for the link. A G/T of

–30 dB/K maintains the desired temperature and allows for a link with proper gain in order to maintain 

link margin and BER requirements.

To provide a backup communication capability, a very low data rate link was added with a rate of 

16 Kbps as illustrated in Figure 29. Using four omni-directional antennas, similar to the ones in Table 12, 

communications can be established in the event that failures occur with the HGA. During launch and at 

the beginning of the trans-lunar injection burn, when the HGA is stowed away, the omni-directional

antennas would be used for primary communications to ensure the life systems of the satellite are working 

and still functional. 
NASA/TM—2011-216894 34

TABLE 12.—DTE COMMUNICATIONS MAIN COMPONENTS

Parabolic HGA • 0.5 m diameter

• Dual-feed

• Supports Ka and S-band

• Passive power

Omni-directional 

antennas

• Four antennas spaced 90° apart to provide 360° coverage

• Supports S-band

• Passive Power

TWTA • Mass: 1.3 kg

• Efficiencies as high as 45% (Ref. 58)

6.0 Summary

The SCHOLR mission was designed to close gaps in the current understanding of the Moon by 

collecting new lunar science data and providing a communications relay service to lunar surface assets on 

the southern hemisphere. The SCHOLR mission and subsystems have been fully explored and defined to 

the detail of a conceptual design. This design includes an orbit optimized to perform both lunar relay 

functions as well as collect important science data. A science payload was selected to unlock the 

mysteries of water at the lunar poles and a lunar communications payload was designed to enable lunar 

surface exploration on the southern hemisphere. An optical communications technology demonstration 

was also incorporated into the design to advance in-space communications capabilities. For further project 

development, targets of opportunity will be sought from NASA and other agencies. NASA’s Space 

Communication and Navigation program is closely reviewing the project results and considering the 

potential continuation of design to fulfill the lunar mission communication requirements. 
NASA/TM—2011-216894 35

Appendix A.—Nomenclature

AIRS Active Infrared Spectrometer

BER bit error rate

C&DH Command and Data Handling

COMPASS Collaborative Modeling for Parametric Assessment of Space Systems

COTS commercial off the shelf

DC direct current 

DET Direct Energy Transfer 

DSN Deep Space Network

DTE direct to Earth 

Eb/N0 signal to noise (dB/W)

ECS Environmental Control System

EIRP Equivalent Isotropic Radiated Power

EPS Electric Power Subsystem

FOV field of view

GN&C Guidance, Navigation, and Control

GRS Gamma Ray Spectrometer

G/T gain over system noise temperature (dB/K)

HGA high gain antenna

JPL NASA Jet Propulsion Laboratory 

K Kelvin

Kbps Kilobits per second

KREEP Potassium, Rare Earth Elements, Phosphorous

LADEE Lunar Atmosphere and Dust Environment Explorer

LCP Lunar Communications Payload

LDPC low-density parity-check 

LOLA Lunar Orbiter Laser Altimeter

LRO Lunar Reconnaissance Orbiter

LROC Lunar Reconnaissance Orbiter Camera

Mbps Megabits per second

MEL Master Equipment List

MOC Mission Operations Center

NASA National Aeronautics and Space Administration

PCM Phase Change Material

PEL Power Equipment List

QPSK quadrature phase shift keying

RF radio frequency

RHCP Right Hand Circular Polarization 

SCaN Space Communications and Navigation

SCEM Scientific Context for Exploration of the Moon

SCHOLR Science Hybrid Orbiter and Lunar Relay
NASA/TM—2011-216894 36

SIP Science Instrument Payload

SOAP Satellite Orbit Analysis Program

STK Satellite Tool Kit

TRL Technology Readiness Level

TT&C Telemetry, Tracking and Command 

TWTA Traveling Wave Tube Amplifier

Wt Watts thermal
NASA/TM—2011-216894 37

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SCience Hybrid Orbiter and Lunar Relay (SCHOLR) Architecture and Design 

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6. AUTHOR(S)

Trase, Kathryn, K.; Barch, Rachel, A.; Chaney, Ryan, E.; Coulter, Rachel, A.; Gao, Hui; 

Huynh, David, P.; Iaconis, Nicholas, A.; MacMillan, Todd, S.; Pitner, Gregory, M.; Schwab, 

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This publication is available from the NASA Center for AeroSpace Information, 443-757-5802

13. SUPPLEMENTARY NOTES

14. ABSTRACT

Considered both a stepping-stone to deep space and a key to unlocking the mysteries of planetary formation, the Moon offers a unique opportunity for scientific study. Robotic 

precursor missions are being developed to improve technology and enable new approaches to exploration. Robots, lunar landers, and satellites play significant roles in 

advancing science and technologies, offering close range and in-situ observations. Science and exploration data gathered from these nodes and a lunar science satellite is 

intended to support future human expeditions and facilitate future utilization of lunar resources. To attain a global view of lunar science, the nodes will be distributed over the 

lunar surface, including locations on the far side of the Moon. Given that nodes on the lunar far side do not have direct line-of-sight for Earth communications, the planned 

presence of such nodes creates the need for a lunar communications relay satellite. Since the communications relay capability would only be required for a small portion of the 

satellite’s orbit, it may be possible to include communication relay components on a science spacecraft. Furthermore, an integrated satellite has the potential to reduce lunar 

surface mission costs. A SCience Hybrid Orbiter and Lunar Relay (SCHOLR) is proposed to accomplish scientific goals while also supporting the communications needs of 

landers on the far side of the Moon. User needs and design drivers for the system were derived from the anticipated needs of future robotic and lander missions. Based on these 

drivers and user requirements, accommodations for communications payload aboard a science spacecraft were developed. A team of interns identified and compared possible 

SCHOLR architectures. The final SCHOLR architecture was analyzed in terms of orbiter lifetime, lunar surface coverage, size, mass, power, and communications data rates. 

This paper presents the driving requirements, operational concept, and architecture views for SCHOLR within a lunar surface nodal network. Orbital and bidirectional link 

analysis, between lunar nodes, orbiter, and Earth, as well as a conceptual design for the spacecraft are also presented. 

15. SUBJECT TERMS

Lunar satellites; Lunar communication; Moon; Relay; Lunar exploration 

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