NASA Revolutionary Aerospace Systems Concepts - Academic Linkage

PROJECT LUNA

Reusable Payload Transportation Vehicle: Technical Report

Theme 2: Gateway-based Cis-Lunar Tug

Submitted By:

Glenn Andrew, Eric Baker, Brandon Caudill, John Frisch, Davis Huffman, Skylar Manteuffel,

Joshua Mataosky, Hank Rains, Siwani Regmi, Shilp Ronvelwala, and Alejandro Sosa

Faculty Advisor:

Dr. Kevin Shinpaugh

Virginia Polytechnic Institute and State University

May 2019

ContentsAbbreviations iii

1 Introduction 1

1.1 Background and Motivation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

1.2 Problem Definition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

1.2.1 Mission Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

1.2.2 Design Objectives . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

1.3 Design Maturity . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

1.4 Concept of Operations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2

1.4.1 Pre-Mission . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2

1.4.2 Reference Mission . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2

1.4.3 End of Life . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3

1.4.4 Expected Missions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3

1.5 Proposed Use Cases . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3

1.5.1 Cislunar Destinations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3

1.5.2 Uncrewed Payloads . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3

2 Orbital Trajectories 4

2.1 Gateway Orbit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4

2.2 Orbital Destinations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4

2.3 Transfer Orbits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4

2.3.1 High-Thrust Transfers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4

2.3.2 Low-Thrust Transfers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5

3 System Architecture 5

3.1 Propulsion System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5

3.1.1 Propulsion Modularity . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6

3.1.1.1 Chemical Design Performance . . . . . . . . . . . . . . . . . . . . . . . 6

3.1.1.2 Electric Design Performance . . . . . . . . . . . . . . . . . . . . . . . . 7

3.1.1.3 Hybrid Modular Performance . . . . . . . . . . . . . . . . . . . . . . . . 7

3.2 Payload-Interaction System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8

3.2.1 Robotic Arm . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8

3.2.2 Embedded Latching End Effector . . . . . . . . . . . . . . . . . . . . . . . . . . . 8

3.2.3 Automated Rendezvous and Docking . . . . . . . . . . . . . . . . . . . . . . . . . 8

3.2.4 Modular Configuration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8

3.3 Structures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9

3.3.1 Bus Structures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9

3.3.2 Propellant Tanks and Lines . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9

3.4 Refueling Module . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9

3.4.1 Bus and Tank Structures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10

3.4.2 Refueling Structures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10

3.4.3 Refueling Operations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10

3.5 Power System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10

3.5.1 Power Generation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10

3.5.2 Energy Storage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11

3.6 Communications and Data Handling . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11

3.6.1 Nominal Operations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11

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3.6.2 Contingency Operations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11

3.6.3 Data Handling . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11

3.7 Thermal Control System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12

3.7.1 Thermal Environment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12

3.7.2 Radiators and Heaters . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12

4 Conclusions 13

4.1 Mission Schedule . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13

4.2 Mass and Cost Budgets . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13

4.3 Missions Risks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14

4.4 Design Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15

References 16

A Calculated Preliminary Budgets 18

A.1 Preliminary Mass Budget . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19

A.2 Preliminary Cost Budget . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19

A.3 Preliminary Power Budget . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20

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AbbreviationsADCS Attitude Determination and Control System

AIAA American Institute of Aeronautics and Astronautics

BOL Beginning of Life

CM Chemical Module

COPV Carbon Overwrapped Pressure Vessel

CSA Canadian Space Agency

DOF Degrees of Freedom

DSN Deep Space Network

ELEE Embedded Latching End Effector

EM Electric Module

EMGF Electro-Mechanical Grapple Fixture

EM-L2 Earth-Moon Lagrange Point 2

EOL End Of Life

ESPRIT European System Providing Refueling, Infrastructure, and Telecommunications

FOV Field Of View

IMM Inverted Metamorphic Multijunction

IR Infrared Radiation

ISS International Space Station

ITO Iridium Tin Oxide

LEE Latching End Effector

LRO Lunar Reconnaissance Orbiter

MLI Multi-Layer Insulation

NDS NASA Docking System

NGC Next Generation Canadarm

NRHO Near Rectilinear Halo Orbit

OBC On-board Computer

OSR Optical-Solar-Reflector

O&M Operations & Maintenance

PTCDH Power, Thermal, Communications and Data Handling

RCS Reaction Control System

RFP Request For Proposal

RM Rrefueling Module

S&MS Structures & Mechanisms System

TCM Temperature Control Module

TCS Thermal Control System

TLI Trans-Lunar Injection

TOF Time Of Flight

TRL Technology Readiness Level

TriDAR Triangulation and LIDAR Automated Rendezvous and Docking

TT&C Telemetry Tracking and Control

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1 Introduction1.1 Background and MotivationLunar exploration is the second phase of NASA’s Mars Exploration Roadmap [1]. The potential to serve

as an outpost for deep space operations and abundance of valuable natural resources make the Moon an

ideal stepping stone for the advancement of human presence in space. The lunar exploration phase is

centered around the Deep Space Gateway, a lunar-orbiting station expected to become operational by

2022 [1, 2]. The Gateway will serve as a communications hub, science laboratory, habitation module,

and aggregation point for various types of payloads. One of NASA’s goals is to have a vehicle capable of

regularly transporting payloads between the Gateway and other cislunar destinations [3, 4]. Project

Luna aims to deliver an innovative design that can meet this goal.

1.2 Problem DefinitionTheme 2 of NASA’s RASC-AL competition challenges teams to design a reusable payload transportation

system, or Tug, capable of servicing uncrewed payloads with a wide range of masses and geometries. The

Tug must operate under a number of high-level requirements established by the RASC-AL committee

and the team, as described below. At the same time, design for the Tug was guided using a number of

self-imposed objectives and a value-oriented approach.

1.2.1 Mission Requirements

Project Luna must meet six main requirements: The Tug shall (1) provide continuous service for a

minimum of 15 years, during which the Tug shall (2) perform a minimum of one mission per year. The

Tug shall (3) be capable of being launched on a single commercial launch vehicle, which shall (4) launch

by the year 2025. The Tug’s (5) combined cost of production, launch, and first year of Operations and

Maintenance (O&M) shall not exceed $500M. Finally, (6) any necessary refueling operations shall occur

without using Gateway propellant [3, 4].

1.2.2 Design Objectives

Table 1.1: Objectives used for design optimiza-tion, along with corresponding weights. Bold ob-jectives are maximized while italic objectives areminimized.

Objective Weight

Tug Wet Mass 0.26

Risk 0.22

Autonomy 0.19

Mission Range 0.13

1st Launch Propellant 0.12

Missions Per Year 0.08

Maximize Minimize

Design decisions were guided using six primary objectives,

shown in Table 1.1. Tug wet mass should be minimized as it

affects the vehicle’s flight-frequency capability, fuel efficiency,

size, and other performance factors. Since the Tug operates

near the Gateway and directly interacts with customer pay-

loads, it is important to minimize the risk of damaging these

assets. Given that the Gateway is uncrewed for long periods

of time and relying on ground station control is not guaran-

teed, it is necessary to maximize the autonomy with which

berthing and docking operations are performed. Additionally,

maximizing the orbital range of the vehicle and the frequency

at which it can perform individual missions is important for

achieving a highly versatile, highly useful design . Finally, the total propellant housed in the first Earth

launch, as a percentage of the total propellant needed to complete the expected missions (shown in

Table 1.2), should be maximized. This allows NASA to delay or avoid costs associated with a propellant

resupply mission.

1.3 Design MaturityThroughout the design process, Project Luna’s architecture and Concept of Operations has matured

significantly as a result of better requirements definition and improved performance estimates. The

original design, driven by the importance of mass minimization, consisted of a solely electric vehicle and

a propellant depot placed in its own orbit.

Improved orbital analysis of low-thrust maneuvers indicated that a solely electric vehicle could not

meet the flight-frequency requirement without significant changes to the propulsion system. Preliminary

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analysis also indicated that hybrid propulsion was able to significantly reduce Time of Flight (TOF),

albeit greatly increasing propellant use. This trade highlighted the need for further analyses focused on

examining different alternatives for achieving hybrid propulsion, and comparing it to single-propulsion-

type designs. Additionally, better definition of Gateway requirements indicated the possibility of using

a docking port for the propellant depot [4]. This allowed the depot to be simplified, as propulsion,

attitude control, and power systems were no longer needed.

The latest iteration of Project Luna features an innovative, unique hybrid modular design with

independent chemical and electric modules. The ability of each module to complete missions indepen-

dently allows the Tug to perform optimally for time and fuel constrained missions. Additionally, the

modules can operate in a hybrid configuration, allowing the Tug to transport a wide range of payloads

while reducing TOF, thus being able to service more payloads over its lifetime. The Tug is capable of

transporting payloads of up to 15,000 kg to Low Lunar Orbit (LLO).

1.4 Concept of OperationsThe design consists of an Electric Module (EM), a Chemical Module (CM), and a Refueling Module

(RM). The first two modules form the payload-transporting Tug, while the RM remains docked at the

Gateway and stores propellant for additional missions. Operations are divided into three segments,

shown in Figure 1.1: Pre-Mission, Reference Mission, and End Of Life (EOL).

Figure 1.1: Concept of Operations Segments: Pre-Mission (1-2), Reference Mission (3-11), End of Life (12-13). Thepre-mission entails launch and RM docking. The reference mission consists of transporting an initial payload of 12,000 kgfrom the Gateway to LLO, after which a second payload of 3,000 kg is transported from LLO back to the Gateway. Endof Life entails the RM undocking and vehicle decommission operations.

1.4.1 Pre-MissionDuring the pre-mission segment, the Tug will be launched using a Falcon Heavy Expendable rocket,

which can carry up to 15,190 kg to lunar orbit [5]. After the Falcon Heavy completes the Trans-Lunar

Injection (TLI) maneuver, the Tug will separate from the launch vehicle and rendezvous with the

Gateway, expected to be stationed in a Near Rectilinear Halo Orbit (NRHO). The Tug will then assist

with the RM’s docking procedure and remain attached to the RM until the first mission is scheduled.

1.4.2 Reference MissionThe reference mission segment, used to represent a likely mission, will begin with the Tug docking with

a payload. The Tug will then begin a transfer maneuver using the CM as a kick stage to reach an

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intermediate orbit, reducing TOF. At the intermediate orbit the Tug will transition to use the EM,

which will complete the transfer to LLO where the Tug will deliver the payload. The Tug will then

dock with a new payload and proceed to use the EM to transfer back to the intermediate orbit. At this

point, the Tug will transition to use the CM to complete the transfer back to the Gateway. The burns

performed for these maneuvers are shown in Figure 2.1. If the following mission requires refueling, the

reference mission will end with the Tug docking with the RM on the Gateway to transfer the necessary

propellant for the next mission. At all times, the Tug will be fueled with an additional factor of safety

of 10% along with 10% ullage.

1.4.3 End of Life

At EOL, expected to occur after at least 15 years, the RM and Tug will both be ready for decommission.

In compliance with category II of planetary protection protocol, the Tug will assist with the RM

undocking procedure and perform a burn to take the three modules into a predetermined lunar

graveyard, minimizing the likelihood of contamination for future missions [6].

1.4.4 Expected Missions

Table 1.2: Expected payload masses and fre-quencies over the 15-year mission lifetime [3].Payloads are expected to be 5,000 kg on av-erage. The total mass expected to be trans-ported over the entire 15-year period is ap-proximately 86,000 kg.

Payload Mass (kg) Frequency

15,000 1

10,000 1

7,000 3

5,000 5

3,000 4

1,000 3

Current lunar missions such as ESA’s Lunar Lander, China’s

Chang’e 4, and THEMIS have masses ranging from 600 kg to

2,000 kg, with landers generally having higher masses [7, 8, 9, 10].

Although little information is available regarding future Gateway-

based missions, a maximum expected payload mass of 15,000 kg

was established. This mass is based on the American Institute of

Aeronautics and Astronautics’s (AIAA) 2018-2019 Undergraduate

Team Space Systems Design Competition. The project’s Request

For Proposal (RFP) suggest a payload mass of 15,000 kg for lunar

habitation modules [11]. Moreover, payload masses and mission

frequencies are expected to increase with the presence of the Gate-

way. Given this, a list of expected missions, shown in Table 1.2,

was established to determine lifetime propellant requirements. The average payload mass is estimated

at 5,000 kg, and the Tug is expected to transport approximately 86,000 kg over its lifetime.

1.5 Proposed Use CasesAlong with designing the Tug, a main goal for this project is to identify use cases that promote commercial

and governmental use of the Gateway. The Tug’s ability to perform high ∆V maneuvers and dock with

a wide variety of payloads allows for numerous use cases. Additionally, the Tug’s hybrid modular design

allows it to perform time-constrained missions while minimizing fuel consumption, meeting potentially

demanding customer needs.

1.5.1 Cislunar Destinations

With the current design, the Tug can perform round-trip missions from the Gateway to LLO, an

approximately 1,600 m/s ∆V maneuver, with payloads up to 15,000 kg in less than one year. LLOs are

of particular interest given future lunar surface and habitation missions. Apart from orbital stability

considerations which make some LLO inclinations unusable, the quasi-inertial nature of said orbits

allows the Tug to reach any desired LLO inclination provided enough time [12]. This allows the Tug to

place landers on surface access paths.

The Earth-Moon Lagrange Point 2 (EM-L2) is another location of interest. Not only does EM-L2 have

scientific significance but could also serve as an important point for establishing lunar communication

arrays [13]. The ∆V required to transfer from NRHO to EM-L2 is approximately 60 m/s, which is well

within the Tug’s range [13].

1.5.2 Uncrewed Payloads

Trends in lunar exploration activities point at future payloads such as traditional lunar landers and

satellites, as well as emerging categories such as constellation deployers and non-operational payloads [7,

3

8, 9, 10, 14, 15]. Given the rise of CubeSats and MicroSats, future lunar missions could entail deploying

a large number of said satellites around the Moon.

Provided the payload has a grapple fixture and the maneuver can be safely performed, the Tug is also

able to dock with malfunctioning payloads and transport them to the Gateway for servicing. In addition,

if a payload with an Electro-Mechanical Grapple Fixture (EMGF) has a malfunctioning communications

system, the Tug could transmit the payload’s data. Similarly, the Tug can transport non-operational

payloads to graveyard orbits, reducing orbital debris in an environment that is projected to become

increasingly crowded [1].

2 Orbital Trajectories2.1 Gateway OrbitDesign for the Tug assumed the Gateway will be stationed in a NRHO. These orbits, which are quasi-

polar members of an orbit family in the Earth-Moon system, are likely candidates for the Gateway’s

orbit for a variety of reasons. These orbits offer distinct advantages as opposed to other lunar orbits:

they exhibit relatively stable behavior, feature a near-uninterrupted view of Earth, relatively constant

sunlight exposure, and require low station-keeping costs [16, 17]. In addition, NRHOs have very low

perilune altitudes, reducing the energy needed to transfer to LLO, in turn allowing for easy staging

operations to the Moon’s surface.

2.2 Orbital DestinationsAs discussed in Section 1.5, LLOs and EM-L2 are attainable destinations which present important

scientific and commercial opportunities. Orbital estimates place the ∆V required for round-trip NRHO-

LLO and NRHO-EML2 transfers at approximately 1,600 m/s and 120 m/s, respectively [13]. As

described in Section 3.1, the Tug can perform a round-trip NRHO-LLO mission while transporting a

15,000 kg payload, meaning the vehicle can achieve higher energy maneuvers for payloads of lower mass.

Moreover, LLOs are quasi-inertial in the Moon inertial frame, allowing the Tug to cover all lunar surface

points in 14 days.

2.3 Transfer OrbitsGiven the Tug’s hybrid modular design, both high and low-thrust maneuvers are needed. High-thrust

maneuvers are performed using the CM and are characterized by short TOF and high propellant

consumption. On the other hand, low-thrust maneuvers are performed using the EM and typically

require long TOF while consuming very little propellant. Both maneuver types require fundamentally

different analysis, discussed below.

Figure 2.1: Burns required to transfer from NRHO (blue) to LLO (green). First, a short high-thrust burn (red arrow)occurs at the NRHO perilune, reducing the apolune. After this, continuous low-thrust burns (red) between ν = 270◦ andν = 90◦, followed by coasting periods (black) are repeated for several passes until LLO is reached. Not to scale.

2.3.1 High-Thrust TransfersHigh-thrust maneuvers consist of a series of short burns that change the desired orbital elements. When

applicable, the Hohmann transfer method is desirable as it is energy-optimized, allowing payload mass

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to be maximized. Other methods such as the one-tangent burn can achieve shorter TOF but generally

require higher energy. For the reference mission, a single short burn is performed at perilune, lowering

the NRHO apolune until the intermediate orbit is reached.

2.3.2 Low-Thrust Transfers

The efficiency of low-thrust maneuvers provides substantial propellant savings as compared to high-thrust.

However, the total thrust produced by electric engines is 10 to 100 times lower than that of chemical

engines, which results in significantly longer TOF. Depending on the destination orbit, electric engine

thrust has to be varied so as to only change the desired orbital elements. For example, to reduce a

highly elliptical orbit down to a circular orbit, the Tug burns from ν = 270◦ through ν = 90◦, where ν

is the true anomaly of the spacecraft. This process repeats upon each pass until the desired orbit is

reached, as shown in Figure 2.1.

3 System ArchitectureThe hybrid modular architecture of the Tug features a unique design that allows for modules to operate

either independently or together. The major systems for the Tug are propulsion, payload-interaction,

power, communications and data handling, thermal control, and reaction control. Excluding MegaFlex

solar arrays, these are shown in Figure 3.1.

(a) (b)Figure 3.1: Primary systems for the (a) CM and (b) EM. The EM features a hollow section for the CM to be inserted inthe hybrid configuration so as to avoid engine plume. The hybrid configuration is shown in Figure 3.4. Secondary internalcomponents such as computers, reaction wheels, trusses, and propellant lines are not shown. Not to scale.

3.1 Propulsion SystemThe inclusion of chemical and electric propulsion was driven by the need to meet the flight-frequency

and payload mass requirements while maximizing fuel efficiency. The CM uses Aerojet’s bi-propellant

HiPAT thrusters, which exhibit a higher fuel efficiency than other similar alternatives, such as R4D

and R42 thrusters. Although one HiPAT thruster is able to produce the thrust necessary to meet the

flight-frequency requirement (445 N), the burn-time needed to achieve the reference mission in one pass

exceeds the thrusters’ demonstrated one-hour limit [18]. Since additional passes result in longer TOF,

two HiPAT thrusters are used in the Tug’s design.

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The EM uses an octagonal array of NASA’s NEXT-C ion thrusters. These engines exhibit high

specific impulse (4,190 s), low power requirement (6.9 kW), and roughly three times the thrust (236

mN) as NASA’s previous generation NSTAR engine [19]. Furthermore, NEXT-C thrusters have been

tested for 48,000 hours of continuous burn, longer than the estimated 14,000 hours needed to complete

the expected missions [19, 20, 21]. In addition to satisfying the Tug’s main 15-year lifetime requirement,

the durability of these engines would allow the Tug to continue operation past it’s nominal lifetime.

3.1.1 Propulsion Modularity

Although it is possible to design a Tug that employs a single type of propulsion system, hybrid propulsion

outperforms both solely chemical and solely electric design alternatives in two key areas: propellant

consumption and TOF. However, single-type propulsion offers important advantages in specific cases.

For this reason, the Tug architecture uses a hybrid modular design with two independent modules that

can achieve optimal performance in said specific cases. Figure 3.2 shows a performance comparison of

hybrid, solely electric, and solely chemical designs.

(a) (b)

(c) (d)Figure 3.2: Comparison of propellant consumption, TOF (left), and maximum attainable ∆V (right) between hybrid,chemical (top), and electric (bottom) designs. A chemical design achieves shorter TOF than a hybrid design but usessignificantly more propellant while achieving a lower ∆V . On the other hand, an electric design uses less propellant butachieves a lower ∆V .

3.1.1.1 Chemical Design Performance

The thrust provided by chemical engines makes them ideal candidates for time-constrained missions.

However, their relatively low specific impulse results in high propellant consumption. To achieve the

reference mission described in Figure 1.1, a solely chemical Tug would need approximately 6,200 kg of

propellant. As shown in Figure 3.2a, such a vehicle would achieve optimal TOF for all payload masses

of interest but would require significantly more propellant than a hybrid design. Moreover, a chemical

design would achieve lower ∆V for the average expected payload mass of 5,000 kg, as shown in Figure

3.2b. Finally, given the launch capability of the Falcon Heavy, a chemical design would require four

propellant resupply missions to achieve the missions outlined in Table 1.2.

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Nevertheless, high-thrust chemical propulsion offers important advantages for time-constrained, low-

mass missions, where propellant consumption and ∆V underperformance is minimized. For this reason,

the CM is able to perform missions independently.

3.1.1.2 Electric Design Performance

Electric engines feature specific impulses that are roughly 10 times larger than that of chemical engines,

making them ideal candidates for propellant-constrained missions, as can be seen in Figure 3.2c. However,

electric engines produce thrust that is much lower than that of chemical engines. Therefore, a solely

electric Tug would need 16 NEXT-C engines to be able to meet the flight-frequency requirement. This

number of engines would require approximately 110 kW to power, posing considerable thermal control

challenges. Solar panel degradation would require panels to be designed to produce 210 kW at Beginning

of Life (BOL), increasing vehicle dry mass significantly. This added dry mass would decrease the

attainable ∆V of the vehicle for payload masses below 5,000 kg, which are expected to be more common

than heavier payloads. Moreover, engine and solar panel packing pose significant design constraints that

limit the performance of an electric design.

However, low-thrust electric propulsion offers considerable advantages for high-mass missions with

lenient timelines. In this range, the ∆V disadvantage of an electric design is minimized. For this reason,

the EM is able to perform missions independently.

(a) (b)Figure 3.3: Comparison of CM (a) and EM (b) propellant use and TOF against the hybrid configuration. Using theEM for high-mass missions results in propellant savings of up to 830 kg. Using the CM for low-mass missions results inTOF savings of up to 130 days.

3.1.1.3 Hybrid Modular Performance

The propellant and TOF savings of a hybrid configuration make such a configuration ideal for the

expected payload masses, launch constraints, lifetime and flight-frequency requirements of the mission.

Additionally, the ability of both modules to operate independently allows the Tug to perform optimally

for the special cases described above. Due to the non-linear nature of ∆V and TOF when transferring

from highly elliptic orbits such as NRHO, use of the chemical and electric modules was determined to

be only beneficial in specific phases of the orbital maneuver. For this reason, in the reference mission,

the CM is only used to transfer from NRHO to an intermediate orbit, after which only the EM is used

to perform the rest of the transfer.

The fraction of total ∆V in the reference mission performed by each module was chosen so that

the TOF for a hybrid configuration is exactly 355 days. This split determines the capabilities of the

CM and EM used in the hybrid design, shown in Figure 3.3. Given the high energy requirement of a

round-trip NRHO-LLO mission, the CM can only independently perform this mission for masses lower

than 1,200 kg. For this range, using the CM results in no more than 100 kg of additional propellant

consumed, while TOF is reduced by up to 130 days, as compared to the hybrid configuration. Since

the EM produces minimum thrust, it can only meet the TOF requirement for masses below 6,000 kg.

However, using the EM results in up to 830 kg in propellant savings for masses above 6,000 kg, provided

a mission can take more than one year to complete.

7

3.2 Payload-Interaction SystemTo interact with payloads, the Tug requires versatile mechanisms to perform berthing and docking

operations. The existence of the Gateway is likely to drive an increase in the mass and variety of

payloads in cislunar space [1]. Therefore, these mechanisms must feature high mass-handling capabilities

and impose little design requirements for payload manufacturers. Additionally, payload berthing and

docking must be capable of being performed semi-autonomously, since the Gateway will be uncrewed

for extended periods of time.

3.2.1 Robotic Arm

The Canadarm2 has serviced the International Space Station (ISS) since 2001 [22] and is capable of

moving payloads of up to 116,000 kg at low speeds [23]. Made of 19 layers of high-strength carbon

fiber thermoplastic [24], the Canadarm2 contains three joints that provide a total of 7 Degrees of

Freedom (DOF). Although usually controlled by ISS crew members, the Canadarm2 can perform

programmed maneuvers semi-autonomously. Given location coordinates for a target, the Canadarm2 can

autonomously observe the object, orient itself, decide what motion is needed, and perform the necessary

movements. However, this information is always relayed back to the ISS and functions can be overridden

at any time [25, 26]. Due to the success of the Canadarm program, the Canadian Space Agency (CSA)

has begun the development of the Next Generation Canadarm (NGC), specifically designed for the

Gateway. The NGC will feature improved autonomous capabilities, which could be integrated into the

Tug’s robotic arm [27].

Given that the maximum expected payload mass for the Tug is 15,000 kg, an 8 m long, 25 cm diameter

scaled-down version of the Canadarm2 is used to perform berthing operations. The mass-handling

capability of the Tug’s robotic arm is estimated to be approximately 45,000 kg, well above the payload

mass requirement. Similar to the Canadarm2, the Tug’s robotic arm docks to payloads using its Latching

End Effector (LEE) to attach to an EMGF located on the surface of said payload.

3.2.2 Embedded Latching End Effector

To dock to payloads, the chemical and electric modules use a simple version of the LEE embedded into

the front of the vehicle. This, combined with the robotic arm, ensures that payloads remain secured

during all stages of a mission. Given that these Embedded LEEs (ELEEs) are only used for securing

payloads, the arm interface mechanism, cameras, and complex robotic controllers needed for Canadarm2

operations, are removed [22, 24, 28, 29].

3.2.3 Automated Rendezvous and Docking

Berthing and docking operations require a high-fidelity imaging system capable of rendering 3D visuals of

the Tug’s target. Developed by CSA, NASA, and Neptec Design Group, TriDAR is a relative navigation

system that uses lasers and thermal imagery to achieve this task [30]. It has been used in the American

Space Shuttle Program, the Cygnus resupply spacecraft, and ISS missions [30]. Specifically designed

for non-cooperative missions which require high levels of autonomy, TriDAR does not require payloads

to have retroreflectors or targets for orientation [30]. This is an important advantage since it does not

impose additional requirements to payload manufactures, thus increasing the variety of payloads the

Tug can service.

3.2.4 Modular Configuration

Table 3.1: Number of EMGFs required for different Tugconfigurations. The hybrid and EM configurations makeuse of the robotic arm as well as an ELEE. The CM onlyuses an ELEE for payload interaction.

Configuration Payload EMGFs

Hybrid 2

Electric 2

Chemical 1

Embedded LEEs and TriDAR vision systems are present

on both the EM and CM. On the other hand, the CM

does not employ a robotic arm. Given that the maximum

payload mass expected to be transported using the CM

is approximately 1,200 kg, the Reaction Control System

(RCS) can be used to position the module and directly

dock to payloads without the need of a berthing mech-

anism. Additionally, the mass of the robotic arm would

8

significantly increase the CM dry mass, eliminating the advantage of an independent CM. For missions

in which the hybrid or electric configurations are used, payloads are required to have two EMGFs

located on perpendicular faces. For missions that use the chemical configuration, payloads only need

one EMGF. These requirements are summarized in Table 3.1.

3.3 StructuresThe vehicle’s bus is required to house all components onboard the Tug. This includes tanks for housing

the different fuel types within each module, an isogrid structure, propellant lines, and truss structures.

The main design drivers for these structures are mass minimization and launch load considerations.

3.3.1 Bus Structures

The EM and CM buses are comprised of an Aluminum 7075 isogrid shell, bulkhead frames, and truss

structures. Plume generated by electric engines could damage surface finishes in the CM, reducing the

lifetime of solar panels and Multi-Layer Insulation (MLI). Therefore, the EM bus features a hollow inner

section with an EMGF to allow the CM to be inserted in the hybrid configuration. To secure internal

components, two types of structures are used: bulkhead frames and ring trusses. Bulkhead frames are

used for sections which contain two or more tanks, while ring trusses are used to hold individual tanks.

3.3.2 Propellant Tanks and Lines

The presence of different propellants in the Tug required designing two types of tanks. For gas propellants,

a Composite Overwrapped Pressure Vessel (COPV) was designed, while a Ti-6AL-4V pressure vessel

was employed for liquid propellants. The COPV is made up of a titanium lining with a carbon-fiber

wrap, which reduces the mass of the tank. The Ti-6AL-4V alloy was chosen as the primary material for

the liquid propellant pressure vessels due to its simplicity and inexpensive manufacturing method. A

Teflon bladder is included inside these tanks with a Nitrogen pressurization system to keep a constant

internal pressure.

To distribute propellant to the different tanks for propulsion and refueling operation, a series of

propellant lines, check valves, and interfaces are included in each module, as shown in Figure 3.4. The

EM has additional propellant lines to allow MMH, NTO, and N2 to be transferred from the RM to

the CM during refueling operations. This allows both modules to be refueled while in the hybrid

configuration, reducing the amount of time and maneuvers needed for refueling operations.

Figure 3.4: Hybrid refueling configuration showing the tanks (white), propellant lines (red), and truss structures (black)on the three modules. The spherical tanks in the RM (left) connect to a refueling interface which can connect to both theEM and CM independently. When in the hybrid configuration, additional propellant lines in the EM direct MMH, NTO,and N2 propellants to the CM inside it.

3.4 Refueling ModuleIn order to meet the 15-year lifetime requirement, the Tug needs to have enough propellant to perform

the missions outlined in Table 1.2. Since carrying this propellant in the Tug would considerably increase

9

the Tug’s wet mass and have a negative effect on other key performance objectives, it is necessary to

have on-orbit refueling capabilities. The RM serves this purpose by housing the propellant required for

both the CM and EM. Given the expected missions outlined in Table 1.2 and the mass constraint set by

the capability of the Falcon Heavy launch vehicle, it is expected that two RMs will need to be launched.

The first RM will launch along with the Tug, and will house approximately 74% of the propellant needed

to complete the missions outlined in Table 1.2. This propellant is estimated to last roughly 13 years.

However, depending on the payload mass and destinations encountered during the Tug’s lifetime, it is

possible that the first RM will provide enough fuel for 15 years.

3.4.1 Bus and Tank Structures

The RM consists of an Aluminum 7075 isogrid bus with internal bulkhead frames to house all propellant

tank types, as shown in Figure 3.4. Containing a total of 8,100 kg of liquid and gaseous propellant, the

RM uses tanks with the same materials as those present in the EM and CM. The tanks are distributed

in two planes, each of them supported by a bulkhead frame. To dock with the Gateway, the RM uses

NASA’s Docking System (NDS), which is capable of transmitting power and data.

3.4.2 Refueling Structures

To produce the pressure necessary for propellant transfer, the RM contains two boost pumps consuming

a total of 3 kW during refueling operations. This power is expected to be provided by the Gateway

through the NDS. Propellant transfer from the RM to the other two modules is achieved with the use

of propellant lines and safety valves located throughout each module. The RM will contain a single

refueling interface with connections for all propellant types. This way, the CM and EM can be refuelled

independently when separate and simultaneously when in the hybrid configuration. A set of propellant

lines inside the EM will directly transport the propellant needed for the CM through the EM structure.

3.4.3 Refueling Operations

After the Tug completes a mission and requires refueling, it will rendezvous with the Gateway and

proceed to dock with the RM. To do this, the Tug attaches to an EMGF on the RM and connects to

the refueling interface. Following successful docking, the RM will begin propellant transfer operations

by initializing its boost pumps and opening propellant fill/drain lines and check valves. It is important

in this phase that all sensor data streams are monitored for anomalies as any failure in hardware carries

significant risk. For this reason, there are check valves located before and after each component in the

propellant transfer line, ensuring any hardware failure can be mitigated.

3.5 Power System

Table 3.2: Power required by Tug systems for CM and EM nominaloperations, as well as contingency operations. During contingencyoperations, batteries will be used. The maximum power required isthe maximum power expected to be used simultaneously.

System EM (W) CM (W) Contingency (W)

Propulsion 56,289 289 26

Robotic Arm 2,000 — —

C&DH 250 250 40

TCS 300 30 210

Batteries 276 276 —

Max Required 57,115 845 276

The power system for the Tug’s mission incorpo-

rates both reliable power generation and storage.

To meet the power requirements for the entire

mission lifetime, solar cell degradation had to

be accounted for. Additionally, the power sys-

tem must fit in the launch vehicle, have reliable

deployment, and be able to rotate to maximize

solar exposure. Power requirements for the Tug’s

systems are shown in Table 3.2.

3.5.1 Power Generation

To supply power for electric and hybrid configurations, two 15 m diameter MegaFlex arrays, produced

by Northrop Grumman, are used over similarly structured UltraFlex arrays. At 30% efficiency, these

panels are attached on either side of the EM and produce 105 kW at BOL [31]. Although MegaFlex has

a lower Technology Readiness Level (TRL) than UltraFlex, it has a higher power-generation capability

of up to 450 kW [32]. The deployment time for MegaFlex arrays is estimated to be 30 min, based on

UltraFlex deployment times. The expanded arrays can withstand up to 0.5 g of linear acceleration, well

under the 0.3 g expected to be encountered during transfer maneuvers [33, 34].

10

The CM uses Inverted Metamorphic Multijunction (IMM) cells, produced by SolAero. At 32%

efficiency, IMM has 2.5% higher efficiency and 42% less mass than traditional triple junction cells [35].

The 2.9 m2 rectangular panel is attached radially and produces 1.6 kW at BOL.

3.5.2 Energy Storage

To determine battery requirements, NASA research and orbital simulations for NRHO and LLO eclipse

periods were used. Only total eclipses and eclipses under 0.3 sun fraction are examined for the EM and

CM, respectively. These are the thresholds at which each module’s power-generation system cannot

meet contingency power requirements. The EM will experience a maximum total-eclipse time of roughly

3 hours and will experience about 3,000 cycles per mission [16]. An independent CM experiences a

maximum of 3.5 h and 3,500 cycles per mission. Given that the batteries power contingency operations,

the batteries for the EM and CM must provide 830 and 970 Wh, respectively.

To meet these requirements, batteries are used as a secondary power source, reducing risk of failure

and enabling longer lifetime [36]. Three battery types were considered: silver zinc, lithium-ion, and

nickel-hydrogen. Lithium-ion batteries were chosen since they have the best specific energy, best energy

density, and comparable cycle lifetime [37, 38, 39]. Three batteries with a total capacity of 8.7 kWh are

on the EM while only one battery with a 1.9 kWh capacity is needed for the CM.

3.6 Communications and Data HandlingThe Tug’s communication system is used to transmit and receive data to and from ground stations or

other spacecraft and assist with docking and rendezvous procedures. The communications architecture

describes how data will be transferred between the modules, the Gateway, and terrestrial ground stations

[40].

3.6.1 Nominal Operations

During nominal operations, the Tug will communicate with Deep Space Network (DSN) ground stations

via the Gateway. The Gateway is intended to act as a communications relay for lunar missions [41].

NRHO orbit allows the Gateway to constantly be in DSN’s field of view [2]. The EM and CM also have

the capability to transmit data to each other. Data transmitted includes essential Telemetry, Tracking,

and Control (TT&C) information required for rendezvous procedures along with the modules’ health

and safety [42]. S-band communication uses two omni-directional antennas on either side of each module

in conjunction with a 2.5 m high-gain antenna to direct essential TT&C data to the other module and

the Gateway. S-band is also used to transmit scientific and engineering data of interest to operators and

the customer.

Given the failure risk of payload capture maneuvers, X-band video feed will be adapted to all modules

to ensure redundancy and reliability when performing these maneuvers. X-band will not be sent directly

to ground station from the modules as the bit rate requirement is too large for the data-handling

components on board, but will instead be relayed through the Gateway’s ESPRIT module.

3.6.2 Contingency Operations

In the case of a nominal communications architecture failure, loss of connectivity, or attitude control

failure, ground crew can rely on the contingency link to obtain essential TT&C data. Using omni-

directional antennas, the Tug will transmit essential health and safety, attitude, and location data [40].

The contingency communication architecture operates only in S-band and is independent of the nominal

architecture.

3.6.3 Data Handling

Data handling is required for the Tug’s motion-control, computations, and electronics. Since the Tug

will be uncrewed, On-Board Computers (OBCs) are necessary for the Tug’s autonomous data transfer

streams.

Data handling will be supported by two RAD 5545 OBCs assigned in a tree architecture with bridges

to SpaceWire routers, channeling computational power to the electronics. The RAD 5545 is a reliable

on-board computer and has supported missions such as the Lunar Reconnaissance Orbiter (LRO) [43].

11

A tree architecture is preferred as opposed to ring or bus architectures, due to the spacecraft having

a lower per endpoint data bandwidth and enough capacity to accommodate the data. Switches or

other concentrated routers are not needed as the data may be communicated between nodes without

using the entire the tree’s bandwidth [44]. This allows the system cost to be reduced. Both RAD 5545

computers will support the Sun sensors, star trackers, RCS, propulsion system, cameras, reaction wheels,

electric power system, and communications bus. Although a single RAD 5545 computer can support the

Tug’s operations, an additional one is included for redundancy, mitigating risk of equipment shutdown.

All OBCs are radiation-hardened and each supports 5.6 giga-operations per second and roughly 3.7

GFLOPS [45].

The EM will have a supplemental RAD 5545 Space VPX with two use interfaces to support the

axis control and movement selection of the robotic arm. The Space VPX model was chosen to support

robotic arm operations due to its ability to withstand greater thermal variability, mitigating the risk of

arm failure [45].

3.7 Thermal Control SystemThe Thermal Control System (TCS) is used to ensure the spatial and temporal temperature gradients

do not exceed the operational limits of each component [40]. A reliable TCS is critical to mitigating the

risk of component failure during operations. Active TCSs are more complex and expensive, thus posing

a considerable failure risk. For this reason, the Tug uses a passive TCS.

Thermal requirements considered in preliminary analysis include operational temperature ranges and

heat dissipated by major components of the design. The driving thermal requirements for each module

are shown in Table 3.3. Survival temperature ranges were not needed because operational temperature

ranges account for worst-case scenarios.

Table 3.3: Thermal requirements for the EM, CM, and RM. Requirements were grouped to form five conduction paths.EM internal components were grouped based on proximity. Each row of the table lists the requirements for a singleradiator.

System Operational Temperature Range (◦C) Worst-Case Thermal Dissipation (W)

Electric Module

Communications -10 to 50 80

Batteries, Data Handling -10 to 35 170

Propulsion -5 to 50 500

Chemical Module 10 to 20 200

Refueling Module 10 to 20 70

3.7.1 Thermal Environment

Table 3.4: The hot case was modelled at the subso-lar point in a 50 km LLO. The cold case was modelledduring an eclipse in NRHO. Extrema of thermal envi-ronments drive radiator and heater requirements [46][47] [16].

Variable Hot Case Cold Case

qsolar (W/m2) 1,420 0

Albedo 0.13 0.06

qIR (W/m2) 1,320 5.2

The Tug’s thermal environment is primarily composed of

solar radiation qsolar, lunar Infrared Radiation (IR) qIR, and

lunar albedo. The TCS was designed using both hot and

cold cases. The hot case was modelled at the subsolar point

in a 50 km LLO. In this case, the Tug has a high view factor

of the Moon, experiences high qIR, and is directly exposed

to the Sun [47]. The cold case was modelled at a point in

NRHO where the Tug is in the Moon’s shadow. In this case,

the Tug has low view factor of the Moon, experiences low

qIR, and is not exposed to solar radiation. Although certain NRHOs do not exhibit lunar eclipses, the

thermal environment was modelled using an NRHO with lunar eclipses [16] to design for the worst-case

scenario, given that the Gateway’s orbit has not yet been decided.

3.7.2 Radiators and Heaters

Radiators on the modules were pointed in the zenith direction, minimizing view factors of both the Sun

and the Moon [46]. Preliminary sizes of radiators were determined by simplifying thermal requirements,

assuming global parameters for each module. The EM was designed to have three independent conduction

12

paths. Internal components placed near one another were linked to the same conduction path. The CM

and RM were each designed to have a single conduction path. The final radiator sizes and corresponding

thermal coats are shown in Table 3.5. Additionally, the MegaFlex solar arrays are coated with 5-mil

A276 white paint [48].

Table 3.5: Thermal radiators placed on the EM, CM, and RM. An Indium Tin Oxide (ITO) Optical Solar Reflector(OSR) surface finish was used for the CM [48].

System Radiator Area (m2) Thermal Coat αrad εradElectric Module

Communications 0.45 5-mil A276 0.36 0.87

Batteries, Data Handling 0.70 5-mil A276 0.36 0.87

Propulsion 1.18 5-mil A276 0.36 0.87

Chemical Module 0.61 5-mil OSR/ITO Pilkington 0.23 0.78

Refueling Module 0.79 5-mil A276 0.36 0.87

All three modules will use Polyimide/Kapton flexible Heaters. These heaters have an operating

temperature range -190◦C to 200◦C, ideal for both hot and cold cases [49]. The minimal power required

for the heaters is provided by the solar panels for the EM and CM, and by the Gateway for the RM.

Temperatures of individual components are monitored by thermistors. Finally, power distributed to the

heaters is regulated by a Temperature Control Module (TCM) [46].

4 Conclusions4.1 Mission ScheduleThe first launch, containing the Tug and RM, is expected to occur in 2025. After launch, the Tug and

RM will perform Gateway-rendezvous and docking maneuvers that will take no longer than 5 days.

Following this, the Tug will be conducting regular transportation operations until 2040, when the mission

is expected to be terminated. In 2040, the Tug will perform the EOL maneuver described in Section

1.4.3. As stated, the propellant housed at first launch is estimated to last between 11 and 14 years, after

which a resupply mission would be needed. However, this timeline can change depending on the specific

missions assigned, which will be defined in more detail in the coming years.

Figure 4.1: Estimate mission timeline including development, nominal operations, and EOL phases. The Tug is expectedto provide service for 15 years, during which a second RM will be launched to provide the propellant needed for anyremaining missions.

4.2 Mass and Cost BudgetsThe projected mass and cost budget are shown in Table 4.1. The total propellant mass in the CM is 965

kg, which has a propellant mass ratio ζ = 0.57. The EM dry mass is larger than the CM due to the large

solar panels and robotic arm. However, the high fuel efficiency of electric propulsion results in a ratio

ζ = 0.15. On the other hand, the hybrid configuration exhibits a ratio ζ = 0.28, ideal for transporting

large payloads in time-constrained missions. The different propellant types inside the RM represent 83%

of its wet mass, and the module’s dry mass is primarily represented by tank and docking structures.

At the first Earth launch, the three modules will together house 87% of the propellant expected to be

needed for the missions outlined in Table 1.2.

Cost is primarily driven by engines, solar panels, and payload-interaction mechanisms. The robotic

arm as well as MegaFlex and NEXT-C arrays in the EM account for 86% of the module’s cost, and 66%

13

of the combined cost of the three modules. The overall cost is estimated at $497M, slightly below the

stipulated limit.

Table 4.1: Projected mass and cost for Tug subsystems, Earth launch, and 1st year of O&M. Overall vehicle cost isestimated to grow 38% [50] while mass growth varies per subsystem, as shown in Table A.1. Average mass growth isestimated at 15% [51]. Preliminary budgets and growth allowance are shown in Appendix A.

Chemical Module Electric Module Refueling ModuleSubsystem Mass (kg) Cost ($k) Mass (kg) Cost ($k) Mass (kg) Cost ($k)

Pro

puls

ion Engines 11 23,115 523 92,460 — —

MMH Propellant 582 121 — — 3,190 890NTO Propellant 364 91 — — 1,994 674Xenon Propellant — — 500 847 2,738 6,373Nitrogen Gas 17 <1 1 <1 129 <1

AD

CS

Reaction Wheels 47 8,280 47 8,280 — —TriDAR 32 TBD 32 TBD — —RCS Engines 9 TBD 9 TBD — —RCS Propellant 19 5 42 12 333 93Star Trackers 7 94 7 94 — —IMUs 2 TBD 2 TBD — —Sun Sensors <1 66 <1 66 — —

S&

MS

Robotic Arm — — 424 48,006 — —ELEE 146 5,393 146 5,393 — —Isogrid Bus 136 32 449 139 150 113MMH Tank 57 742 — — 422 1,132NTO Tank 22 554 — — 161 944Xenon Tank — — 13 6,524 132 5,260RCS Tank <1 123 <1 279 <1 684Nitrogen Tank 14 1,936 <1 — 109 4,109Housing Structures 34 TBD 143 TBD 99 TBDRefueling System 12 TBD 12 TBD 171 TBDNDS — — — — 349 TBD

PT

CD

H Solar Panels 4 307 827 58,002 — —C&DH 46 2,530 46 2,530 46 2,530Batteries 88 927 88 927 — —TCS 20 1,380 130 6,900 46 494Module Total 1,670 45,696 3,442 230,458 10,068 23,294Launch Cost ($k) 150,000 Total Mass (kg) 15,1821st Year O&M ($k) 48,000 Total Cost ($k) 497,500

4.3 Missions RisksOne of the most important sources of risk for Project Luna are the incorporation of several cutting-edge,

low-TRL components into the design. As of 2015, MegaFlex solar panels are at TRL6+ and more up-to-

date information is unavailable [32]. Similarly, NEXT-C engines are not scheduled to be commercially

available until late 2019, and have never been simultaneously tested in arrays of more than three engines

[52, 19]. Regular, semi-autonomous payload interaction operations are have not been extensively tested

and are prone to failure, absent rigorous qualification testing.

Figure 4.2: Matrix of main missions risks. Risks are in order of descending risk score.

Another source of risk is uncertainty in the Tug’s customer base. While outlining expected missions

served as a design guideline, little information is available on the actual missions expected to be assigned

14

to the Tug. There is risk the RM will not have enough fuel to meet the 15-year lifetime requirements,

requiring the launch of more than one additional RM. Separately, the storage of five different propellants

on-board the RM complicates refueling procedures. Though check-valves have been integrated into the

fuel lines to reduce the chances of leaks, an anomaly during refueling operations could be catastrophic.

Finally, key aspects of the Gateway remain undetermined, including its orbit and telecommunication

capabilities. The Tug relies on the Gateway’s ESPRIT module to act as a communications relay. If the

Tug is not in the Field of View (FOV) of the Gateway for enough time or is not fully compatible with

the ESPRIT module, the Tug’s ability to downlink data to terrestrial ground stations will be hindered.

These risks are ranked in Figure 4.2

4.4 Design SummaryProject Luna delivers a unique, highly innovative design that is ideal for the payload-transportation

missions envisioned as part of NASA’s Moon Exploration phase. The Tug’s hybrid modular architecture

delivers optimal performance for a long-duration mission that includes a large variety of payloads,

high-energy orbital transfers, and an ambitious flight-frequency requirement. The highly versatile design,

which allows for independent operation of the EM and CM, results in the ability to maximize propellant

and TOF savings when desired. In addition, the Tug can operate in a hybrid configuration, making use

of the specific advantages offered by chemical and electric propulsion. In this configuration, the Tug can

complete a round-trip NRHO-LLO mission with a payload of up to 15,000 kg while still meeting the

TOF requirement. Maximum modular performance for the three configurations is shown in Table 4.2.

Launching a RM in 2025 along with the Tug allows Project Luna to complete a large number of

missions, delaying or possibly avoiding the need for propellant resupply missions. With a total expected

cost of $497M, possibly avoiding additional launch costs is an important advantage offered by the Tug’s

design. Finally, the use of highly autonomous berthing and docking systems ensures that payload

capture operations can be performed regularly while requiring minimal to no human intervention. This

allows the Tug to complete missions at a higher frequency.

Table 4.2: Maximum performance estimates for a round-trip NRHO-LLO mission. The maximum payload mass is limitedby the TOF requirement for the EM and by the maximum tank capacity for the CM, shown in bold.

Configuration Max Payload Mass (kg) Propellant Used (kg) TOF (days)

Hybrid 15,000 1,300 355

Electric 6,000 220 365

Chemical 1,200 875 30

15

Bibliography[1] NASA Global Exploration Roadmap. NASA, 2018. url: https://www.nasa.gov/sites/default/

files/atoms/files/ger_2018_small_mobile.pdf.

[2] William Gerstenmaier. Cislunar and Gateway Overview. url: https://www.nasa.gov/sites/

default/files/atoms/files/cislunar-update-gerstenmaier-crusan-v5a.pdf (visited on

12/02/2018).

[3] Rascal Theme 2 Requirements. Rascal, 2019. url: http://rascal.nianet.org/competition-

basics/.

[4] Troutman P. 2019 RASC-AL Q&A Session Transcript. NASA. 2018.

[5] Mike Carney. Launch Vehicle Performance Website. NASA, 2019. url: https://elvperf.ksc.

nasa.gov/Pages/Default.aspx.

[6] Lisa Pratt. About the Categories. NASA, 2018. url: https://planetaryprotection.nasa.gov/

about-categories/.

[7] Gunter’s Space Page. Chang’e 3, 4 (CE 3, 4) / Yutu. url: https://space.skyrocket.de/doc_

sdat/change-3.htm (visited on 12/01/2018).

[8] David Szondy. Astrium presents study for European lunar landing in 2019. url: https://

newatlas.com/astrium-lunar-lander-study/24685/ (visited on 12/01/2018).

[9] Chung Sara J. Hatch Jin H. Ma Theodore H. Sweetser Stacy S. Weinstein-Weiss Stephen B.

Broschart Min-Kun J. and Vassilis Angelopoulos. PRELIMINARY TRAJECTORY DESIGN FOR

THE ARTEMIS LUNAR MISSION. Tech. rep. July 2009. url: http://www.igpp.ucla.edu/

public/THEMIS/SCI/Pubs/artemis/broschart_etal.pdf.

[10] Lunar Reconnaissance Orbiter (LRO): Leading NASA’s Way Back to the Moon. Tech. rep. June

2009. url: https://www.nasa.gov/pdf/360020main_LRO_LCROSS_presskit2.pdf.

[11] Request for Proposal Reusable Lunar Surface Access Vehicle. AIAA. 2018. url: https://www.

aiaa.org/docs/default- source/uploadedfiles/education- and- careers/university-

students/design-competitions/undergraduate-team-space-systems-2018-2019.pdf?

sfvrsn=9b9f6ea6_0.

[12] D. Davis et al. “Orbit Maintenance and Navigation of Human Spacecraft at Cis-Lunar Near

Rectilinear Halo Orbits”. In: (2018).

[13] Steven McCarty Kurt Hack Ryan Whitley Diane Davis Cesar Ocampo Melissa McGuire Laura

Burke. Low Thrust Cis-lunar Transfers Using a 40 kW-Class Solar Electric Propulsion Spacecraft.

NASA. url: https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20170009215.pdf

(visited on 2019).

[14] TESS Discovering Exoplanets Orbiting Nearby Stars. Tech. rep. 2018. url: http : / / www .

northropgrumman . com / Capabilities / ScienceEnvironmentSatellites / Documents / TESS _

Factsheet.pdf.

[15] JAXA. Asteroid Explorer ”Hayabusa2”. url: http : / / global . jaxa . jp / projects / sat /

hayabusa2/index.html (visited on 12/01/2018).

[16] Ryan J. Whitley Kevin A. Bokelmann Diane C. Davis Christopher F. Berry Jacob Williams

David E. Lee. TARGETING CISLUNAR NEAR RECTILINEAR HALO ORBITS FOR HUMAN

SPACE EXPLORATION. NASA. url: https://ntrs.nasa.gov/archive/nasa/casi.ntrs.

nasa.gov/20170001352.pdf (visited on 04/20/2019).

[17] Roland Martinez Ryan Whitley. Options for Staging Orbits in Cis-Lunar Space. NASA Johnson

Space Center. 2015. url: https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/

20150019648.pdf.

16

[18] Bipropellant Data Sheets. Aerojet. 2005. url: http://www.rocket.com/sites/default/files/

documents/Capabilities/PDFs/Bipropellant%20Data%20Sheets.pdf.

[19] Timothy Rechark Scott Graham. Gridded Ion Thrusters (NEXT-C). NASA. 2008. url: https:

//www1.grc.nasa.gov/space/sep/gridded-ion-thrusters-next-c/.

[20] Walter Santiago Hani Kamhawi James Polk John Snyder Richard Hofer Frank Picha Jerry Jackson

May Allen Daniel Herman Todd Tofil. Overview of the Development and Mission Application of

the Advanced Electric Propulsion System (AEPS). NASA. 2018. url: https://ntrs.nasa.gov/

archive/nasa/casi.ntrs.nasa.gov/20180001297.pdf.

[21] BHT-8000 BUSEK Hall Effect Thruster. BUSEK. 2016. url: http://www.busek.com/index_

htm_files/70000703%20BHT-8000%20Data%20Sheet%20Rev-.pdf.

[22] Canadian Space Agency. Canadarm 2: Servicing the International Space Station since 2001.

Canadian Space Agency. url: http://www.asc-csa.gc.ca/eng/iss/canadarm2/about.asp.

[23] Brian Peacock Nancy J. Currie. “International Space Station Robotic Systems Operations – A

Human Factors Perspective”. In: ().

[24] Canadarm and Canadarm2 - A Comparable Table. url: http://www.asc-csa.gc.ca/eng/iss/

canadarm2/canadarm-canadarm2-comparative-table.asp.

[25] Shelley Canright and Brian Dunbar. The Canadian Crane - Educator Features. 2009. url: https:

//www.nasa.gov/audience/foreducators/k-4/features/F_Canadian_Crane.html.

[26] Canadian Space Agency. Canadarm. 2018. url: http://www.asc-csa.gc.ca/eng/canadarm/

default.asp.

[27] Canadian Space Agency. The Next-Generation Canadarm. 2013. url: http://www.asc-csa.gc.

ca/eng/canadarm/ngc.asp.

[28] Canadian Space Agency. Canadarm2’s Data Sheet. Canadian Space Agency. url: http://www.asc-

csa.gc.ca/eng/iss/canadarm2/data-sheet.asp.

[29] B. Walker and R. Vandersluis. Design, Testing and Evaluation of Latching End Effector.

[30] Rendezvous and Docking Sensors - TriDAR. 2018. url: https://neptec.com/products/

docking/.

[31] UltraFlex Solar Array Systems. Northrop Grumman, 2018. url: https://www.northropgrumman.

com/Capabilities/SolarArrays/Documents/UltraFlex_Factsheet.pdf.

[32] Michael E. McEachen David M. Murphy Michael I. Eskenazi and James W. Spink. UltraFlex and

MegaFlex – Development of Highly Scalable Solar Power. Orbital ATK, Space Systems Group,

Space Components Division, 2015. url: https://ieeexplore.ieee.org/stamp/stamp.jsp?

arnumber=7355945&tag=1.

[33] Scott W. Benson. Solar Power for Outer Planets Study. NASA Glenn Research Center. 2007. url:

https://www.lpi.usra.edu/opag/nov_2007_meeting/presentations/solar_power.pdf.

[34] Cygnus OA-4 Array Deploy 3FPS. Orbital ATK. 2015. url: https://www.youtube.com/watch?

v=bnDeJ7saLQY.

[35] IMM Solar Cells Datasheet. 2018. url: https://solaerotech.com/wp-content/uploads/2018/

04/IMM-alpha-Preliminary-Datasheet-April-2018-v.1.pdf (visited on 11/28/2018).

[36] Harvey Frank Gerald Halpert and Subbarao Surampudi. Batteries and Fuel Cells in Space. The

Electrochemical Society. url: https://www.electrochem.org/dl/interface/fal/fal99/IF8-

99-Pages25-30.pdf (visited on 04/20/2019).

17

[37] Ralph D. Lorenz Viktor V. Kerzhanovich Andrew J. Ball James R. C. Garry. Planetary Landers

and Entry Probes. Cambridge University Press. url: https://www.e-reading.club/bookreader.

php/138786/Ball_-_Planetary_Landers_and_Entry_Probes.pdf (visited on 04/20/2019).

[38] Concha M. Reid. Batteries at NASA – Today and Beyond. NASA Glenn Research Center. url:

https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20160004091.pdf (visited on

04/20/2019).

[39] Energy Storage Technologies for Future Planetary Science Missions. NASA. url: https://

solarsystem . nasa . gov / resources / 549 / energy - storage - technologies - for - future -

planetary-science-missions/ (visited on 04/20/2019).

[40] The Space Technology Library Editorial Board. Space Mission Analysis and Design. The Space

Technology Library, Jan. 1999.

[41] Jason Crusan. Future Human Exploration Planning: Lunar Orbital Platform-Gateway and Science

Workshop Findings. url: https://www.nasa.gov/sites/default/files/atoms/files/

20180327-crusan-nac-heoc-v8.pdf (visited on 12/02/2018).

[42] John E. Keesee. Satellite Telemetry, Tracking and Control Subsystems. Massachusetts Institute of

Technology, 2013. url: https://ocw.mit.edu/courses/aeronautics-and-astronautics/16-

851-satellite-engineering-fall-2003/lecture-notes/l20_satellitettc.pdf (visited on

12/02/2018).

[43] Herbert J. Kramer. LRO (Lunar Reconnaissance Orbiter) + LCROSS. NASA, 2009. url: https:

//earth.esa.int/web/eoportal/satellite-missions/l/lro.

[44] Joseph R Marshall. SpaceWire Satellite Usage. BAE Systems, 2013. url: http://citeseerx.ist.

psu.edu/viewdoc/download?doi=10.1.1.691.8781&rep=rep1&type=pdf.

[45] BAE. RAD5545 Specifications. BAE Systems, 2019. url: https://www.baesystems.com/en-

us/our-company/inc-businesses/electronic-systems/product-sites/space-products-

and-processing/radiation-hardened-electronics.

[46] David G. Gilmore. Spacecraft Thermal Control Handbook 2nd ed. The Aerospace Corporation,

2002.

[47] Lunar Reconnaissance Orbiter (LRO) Thermal Design Drivers and Current Thermal Design

Concept. NASA Goddard, Edge Space Systems, Lockheed Martin. Aug. 2005. url: https://tfaws.

nasa.gov/TFAWS05/Website/files/ThermalPaperSession/tfaws_2005_mastropietro.pdf.

[48] Charles Baker. Lunar Reconnaissance Orbiter Project: Thermal System Specification. Goddard

Space Flight Center. Apr. 2019. url: https://snebulos.mit.edu/projects/crater/file_

cabinet/0/02004/02004_02_rCm.pdf.

[49] Polyimide/Kapton Flexible Heaters. epec Engineered Technologies. url: https://www.epectec.

com/flexible-heaters/polyimide-kapton-heaters.html (visited on 05/10/2019).

[50] Daniel Judnick Ingrid Hallgrimson Megan Youngs Marc Hayhurst Robert Bitten and Stephen

Shinn. Historical Mass, Power, Schedule & Cost Growth for NASA Instruments & Spacecraft.

NASA Goddard and The Aerospace Corporation, Aug. 2016.

[51] Vincent Larouche. NASA Mass Growth Analysis - Spacecraft & Subsystems. NASA, 2014.

[52] Scott W. Benson Michael J. Patterson. NEXT Ion Propulsion System Development Status and

Performance. NASA Glenn Research Center, 2007. url: http://citeseerx.ist.psu.edu/

viewdoc/download?doi=10.1.1.485.3395&rep=rep1&type=pdf.

18

A Calculated Preliminary BudgetsA.1 Preliminary Mass Budget

Table A.1: Table separates the each subsystem and its components’ masses. The CM, EM, and RM are each listed,presenting their total mass value. The masses listed are before the growth estimate is applied.

Subsystem Chemical Module (kg) Electric Module (kg) Refueling Module (kg) Growth (%)

Pro

puls

ion Engines 10 480 — 9

MMH Propellant 539 — 2,953 8NTO Propellant 337 — 1,846 8Xenon Propellant — 458 2,512 9

Nitrogen Gas 16 1 119 8

AD

CS

Reaction Wheels 44 44 — 8TriDAR 30 30 — 8

RCS Engines 8 8 — 8RCS Propellant 17 39 308 8Star Trackers 7 7 — -3

IMUs 2 2 — 8Sun Sensors < 1 < 1 — 8

S&

MS

Robotic Arm — 347 — 22ELEE 120 120 — 22

Isogrid Bus 112 368 123 22MMH Tank 47 — 346 22NTO Tank 18 — 132 22Xenon Tank — 11 108 22RCS Tank < 1 < 1 < 1 22

Nitrogen Tank 12 1 90 22Housing Structures 28 117 81 22Refueling System 10 10 140 22

NDS — — 286 22

PT

CD

H Solar Panels 3 678 — 22Batteries 72 72 — 22C&DH 47 47 47 -3TCS 15 100 35 30Total 1,494 2,941 9,127

A.2 Preliminary Cost BudgetTable A.2: Table separates each subsystem’s components and the their associated costs. The costs presented are beforethe growth estimate. The CM, EM, and RM are all separated, showing their respectively total cost value.

Subsystem Chemical Module ($k) Electric Module ($k) Refueling Module ($k) Growth (%)

Pro

puls

ion Engines 16,750 67,000 — 38

MMH Propellant 88 — 645 38NTO Propellant 66 — 488 38Xenon Propellant — 614 4,618 38

Nitrogen Gas < 1 < 1 < 1 38

AD

CS

Reaction Wheels 6,000 6,000 — 38TriDAR TBD TBD — 38

RCS Engines TBD TBD — 38RCS Propellant 4 8 67 38Star Trackers 68 68 — 38

IMUs TBD TBD — 38Sun Sensors 48 48 — 38

S&

MS

Robotic Arm — 34,787 — 38ELEE 3,908 3,908 — 38

Isogrid Bus 23 101 82 38MMH Tank 538 — 820 38NTO Tank 401 — 684 38Xenon Tank — 4,727 3,811 38RCS Tank 89 202 495 38

Nitrogen Tank 1,403 < 1 2,978 38Housing Structures TBD TBD TBD 38Refueling System TBD TBD TBD 38

NDS — — TBD 38

PT

CD

H Solar Panels 222 42,031 — 38Batteries 672 672 — 38C&DH 1,833 1,833 1,833 38TCS 1,000 5,000 358 38Total 33,113 166,999 16,880

19

A.3 Preliminary Power BudgetTable A.3: The table below lists the maximum power requirements for subsystems on each module before the growthestimate. Power will be produced on the CM through a SolAero IMM solar panel. Power for the EM will be provided bytwo Northrop Grumman MegaFlex Arrays. The RM will take power when neccessary from the Gateway.

Subsystem Chemical Module (W) Electric Module (W) Refueling Module (W) Growth (%)Engines 0 55,200 — 16

Reaction Wheels 150 150 — 16TriDAR 40 40 — 16

RCS Engines 35 35 — 16Star Trackers 27 27 — 16

IMUs 36 36 — 16Sun Sensors 1 1 — 16Robotic Arm — 2,000 — 16

Refueling System 0 0 3,000 16NDS — — TBD 16

Solar Panels 0 0 — 16C&DH 250 250 250 16TCS 30 300 270 16

Maximum 250 55,200 3,000 —

20