Proof of Concept for Sub-Orbital Lunar Transport System

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Gator Hopper-Final Design Report
Logan Monday, Ryan Morin, Logan Ritten, Marcos Sanchez, Ishaan
Singh, Jonathan Weese, Jacob Williams, Calvin Wuorinen
EAS 4700 Aerospace Design 1
Prof. Michael Generale and Prof. Ting Dong
December 9th
, 2020

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Contents
LIST OF ACRONYMS ................................................................................................................................5
EXECUTIVE SUMMARY............................................................................................................................6
PURPOSE...................................................................................................................................................6
CONCEPT...................................................................................................................................................6
TEAM & ORGANIZATION ..............................................................................................................................7
MISSION OVERVIEW ...............................................................................................................................8
MISSION PHASES ........................................................................................................................................8
CONCEPT OF OPERATIONS.............................................................................................................................8
VEHICLE OVERVIEW.....................................................................................................................................9
PROPULSION.........................................................................................................................................10
OVERVIEW...............................................................................................................................................10
ENGINE SELECTION ....................................................................................................................................10
PROPELLANT FEED SYSTEM .........................................................................................................................12
TANK SELECTION.......................................................................................................................................13
FEED LINE SELECTION.................................................................................................................................15
FEED LINE COMPONENT SELECTION ..............................................................................................................16
FEED SYSTEM LAYOUT................................................................................................................................17
MASS PROPERTIES ....................................................................................................................................18
TRAJECTORY AND GUIDANCE ...............................................................................................................19
OVERVIEW...............................................................................................................................................19
TRAJECTORY HARDWARE............................................................................................................................19
HAZARD AVOIDANCE SYSTEM......................................................................................................................21
MAXIMUM RANGE EXAMPLE.......................................................................................................................23
COMMUNICATION AND TRACKING ......................................................................................................27
OVERVIEW...............................................................................................................................................27
TRANSCEIVER ...........................................................................................................................................27
ON-BOARD-COMPUTER .............................................................................................................................28
HARDWARE AND ELECTRONICS ............................................................................................................30

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OVERVIEW...............................................................................................................................................30
ELECTRICAL INVENTORY..............................................................................................................................30
SOLAR PANELS..........................................................................................................................................31
BATTERIES ...............................................................................................................................................32
GIMBAL ACTUATORS..................................................................................................................................32
POWER DISTRIBUTION ...............................................................................................................................33
CABLE MASS ............................................................................................................................................34
ENVIRONMENTAL .................................................................................................................................35
OVERVIEW...............................................................................................................................................35
THERMAL ................................................................................................................................................35
ACOUSTIC................................................................................................................................................44
VIBRATION ..............................................................................................................................................45
MECHANISMS AND STRUCTURES..........................................................................................................47
OVERVIEW...............................................................................................................................................47
BASE STRUCTURE ......................................................................................................................................47
LANDING GEAR.........................................................................................................................................48
SLT STRUCTURE AND MOUNTING.................................................................................................................49
GIMBAL ASSEMBLY....................................................................................................................................51
FABRICATION AND MANUFACTURING............................................................................................................52
STRESS ANALYSES......................................................................................................................................52
FINAL PROPERTIES.....................................................................................................................................54
GROUND QUALIFICATION AND TESTING...............................................................................................55
OVERVIEW...............................................................................................................................................55
STRUCTURAL LOAD TESTING........................................................................................................................55
SHOCK LOAD TESTING................................................................................................................................56
ACOUSTICS ENVIRONMENT TESTING .............................................................................................................56
VIBRATION ENVIRONMENT TESTING .............................................................................................................57
THERMAL ENVIRONMENT TESTING ...............................................................................................................57
LIFTOFF AND ASCENT VENTING TESTING ........................................................................................................58
PROJECT MANAGEMENT ......................................................................................................................59
RISK ASSESSMENT.....................................................................................................................................59
REQUIREMENTS MATRIX ............................................................................................................................61
OISR AND SCHEDULE.................................................................................................................................61

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BUDGET ................................................................................................................................................63
OVERVIEW...............................................................................................................................................63
STAFF .....................................................................................................................................................63
CONCLUSION ........................................................................................................................................64
WORKS CITED .......................................................................................................................................65
APPENDICES..........................................................................................................................................66
APPENDIX A: VEHICLE SUB-SYSTEM RISK ASSESSMENT.....................................................................................66
APPENDIX B: SLT VERIFICATION FIGURES.......................................................................................................69
APPENDIX C: SLT MASS PROPERTIES.............................................................................................................70
APPENDIX D: VARIOUS STRESS ANALYSES......................................................................................................71
APPENDIX E: REQUIREMENTS MATRIX ..........................................................................................................72
APPENDIX F: BUDGET BREAKDOWN..............................................................................................................73
APPENDIX G: SLT MASS BREAKDOWN (WITHOUT PAYLOAD) .............................................................................75
APPENDIX H: TRADE STUDIES......................................................................................................................76
APPENDIX I: ENGINEERING CHANGE NOTICES .................................................................................................81

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List of Acronyms
CMP – Co-Manifested Payload
COPV – Composite Overwrapped Pressure Vessel
COTS – Commercial Off-The-Shelf
ECN – Engineering Change Request
EDL – Entry, Descent, Landing
G&C – Guidance and Control
HMP – Habitable Mobility Platform
LBC – Lunar Base Camp
LiDAR – Light Detection and Ranging
MLI – Multi Layered Insulation
MMH – Monomethylhydrazine
NTO – Dinitrogen tetroxide
OBC – On-board Computer
PCB – Power Control Board
RCS – Reaction Control System
RHU – Radioisotope Heater Unit
SCS - Small Communications Satellite
SLS – Space Launch System
SLT – Sub-orbital Lunar Transport

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Executive Summary
Purpose
As plans for lunar outposts ramp up and NASA’s Artemis mission comes closer to fruition, support
equipment must be designed to ensure human and mission safety. Soon, NASA plans on having
lunar camps on the surface of the moon, with astronauts permanently stationed and performing
excursions across the moon. To perform these excursions, they will use an HMP, basically a
pressurized lunar rover. What would happen if astronauts on a mission hundreds of kilometers
away from the LBC were to damage the HMP or something worse?
This project describes a preliminary design for a semi-autonomous/remotely operated SLT vehicle
that would be capable of saving these astronauts and deliver emergency supplies to them, as seen
in Figure 1. The proposed design is capable of carrying up to 200 kg of emergency supplies from
a future permanently crewed station near Shackleton Crater to a remote exploration site and return.
The proposed design supplies this vehicle to NASA no later than February 2029, to be used as a
support vehicle at NASA’s LBC’s.
Concept
The proposed vehicle features a large assortment of design parameters, with a priority in these 3
categories:
1) Hazard Avoidance: The vehicle has been designed to avoid mountains/rugged terrain on
its trajectory and search for/find a safe landing zone near the downed HMP. To accomplish
this, the vehicle is outfitted with LiDAR and other sub systems to ensure that it can detect
hazards and find a safe landing zone. It is also outfitted with RCS engines and two main
engines that act on a gimbal which can provide immediate thrust vectoring to change
course.
2) Surviving Lunar Darkness: The vehicle has been designed to have two 25 Ah batteries
and as little power draw as possible, which allows it to survive while being powered on for
more than 72 hours in complete darkness. This provides time for the astronauts to retrieve
the payload and complete system checks before it returns.

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3) Payload Capacity: The vehicle has been designed to carry a payload of 200 kg for long
distances, up to 300 km across the lunar surface. To provide the thrust required to lift the
vehicle and the payload, two 4,000 N R40B engines are used. The structure of the vehicle
has also been designed to be as light as possible.
Figure 1: SolidWorks render of the Gator Hopper with an astronaut for scale.
Team & Organization
The team has been organized in a top-down arrangement, with a Project Manager overseeing the
Scheduler, Budget Analyst, and Integration Engineer, and an Integration Engineer overseeing all
the vehicle sub-teams. A breakdown of this structure is seen in Figure 2.
Figure 2: Breakdown of team structure.

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Mission Overview
In the span of the SLT’s existence, it will be loaded onto the SLS and launched no later than
November 2029, place itself into lunar orbit, land itself on the lunar surface near the Lunar Base
Camp (LBC), and deliver emergency payloads of up to 200 kg up to 300 km and return. It has
been designed to be reusable, survive more than 72 hours in complete darkness, and produce no
waste.
Mission Phases
There are 3 main phases that have been identified for mission success. These phases are listed
below with a brief description for each:
1) Pre-Launch – The initial design, build, and test phase. Consists of designing and building
the SLT, testing the SLT at various test centers, and delivering it and loading it onto the
SLS for launch towards the moon.
2) Lunar Base Camp Rendezvous – The lunar delivery phase. Consists of launching
onboard the SLS, providing impulses to enter lunar orbit, and providing entry, descent, and
landing burns to land at the Lunar Base Camp.
3) Effective Mission Start – The beginning of the concept of operations. This is where the
payload is loaded onto the SLT, the vehicle delivers it to the astronauts, and the SLT returns
to the LBC.
Concept of Operations
Once the SLT has landed on the surface of the moon and has been refueled, it can perform
emergency missions. In the event of an emergency, the following steps will be performed:
1) Pre-Launch (at LBC): Load the emergency payload, power on the vehicle, ensure
connections are established, perform a system check, and top off the propellant tanks.
2) Launch (from LBC, 0 seconds): Ignite and gimbal the two throttleable main engines to
place SLT on trajectory toward the downed HMP.

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3) En-Route (to HMP, 167.4 seconds): Main engine cutoff, use star trackers, accelerometers,
and gyroscopes to track position, and coast until descent burns.
4) En-Route (to HMP, 642.75 seconds): Reignite the two main engines for a controlled
descent. Begin scanning ground with obstacle avoidance systems to ensure safe path and
landing area.
5) Obstacle Detected (to HMP): If an obstacle is detected, alter the current trajectory
towards a safe zone using the RCS thrusters and the main engines which act on a gimbal.
6) Landing and Waiting Return (810.15 seconds): Upon landing, shutoff hazardous
systems and vehicle before astronauts approach and then unload the payload.
7) Return to LBC: Once cleared for return, the vehicle will use the same procedure above.
Vehicle Overview
Throughout this report, images of the design will be seen. Figure 3 is a breakdown of what each
part is on the Gator Hopper SLT. In the top view provided, the solar panels and MLI has been
hidden to allow a proper view of the electronics bay underneath. In side view #2, the MLI on that
corner of the vehicle has been hidden to allow a proper view of the payload and inner design of
the vehicle.
Figure 3: Side and Top views of Gator Hopper.

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Propulsion
Overview
To traverse the maximum round trip distance of 660 km given in ASD-SLT-020, an adequate
propulsion system was required. Because there is no lunar atmosphere, any airbreathing or rotary
propulsion systems are inappropriate which leaves chemical propulsion methods as the main
contender. Several engines with various propellant combinations were considered including
LOX/LNG, monopropellant hydrazine, LOX/IPA, and NTO/MMH. In addition, RCS engines
were deemed necessary to stabilize the aircraft during vertical launch and vertical landing while
also allowing for fine trajectory adjustments. To reduce SLT mass, it was a derived requirement
for the reaction control system (RCS) engines to utilize identical propellants to the main engines.
Engine Selection
The derived minimum thrust for the main engines was determined to be equivalent to the maximum
vehicle weight of approximately 3340 N. A wider thrust range is desired, however, as takeoff and
landing thrust should be higher and lower, respectively, than the vehicle weight to achieve
appropriate acceleration. The main engine trade study performed concluded that no cryogenic
propellant engines within the desired thrust range were available. In addition, to minimize the cost
and development time of propellant storing infrastructure, non-cryogenic propellants were desired.
For these reasons, the Aerojet Rocketdyne R-40B using NTO/MMH was selected for the main
engines. Determined to be a suitable main engine, it has a relatively high specific impulse, is
throttleable, and is flight proven. To achieve the desired thrust profile, two R-40B’s will be
utilized. Specifications for this engine are shown in Figure 4.
After main engine selection, RCS engines were selected to use an identical propellant combination,
as this reduces the number of tanks required and correspondingly minimizes vehicle mass. For the
RCS thrusters, a desired thrust range of 60-500 N was selected by comparison with similar
vehicles. Additionally, a low minimum impulse bit was desired for fine control of translation and
applied roll, pitch, and yaw moments. With these considerations, Aerojet Rocketdyne’s R-1E
engine was selected. The datasheet for the R-1E is shown below in Figure 5. Using 12 R-1E

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engines in four clusters of three orthogonally mounted thrusters allows for translation in any
direction as well as pitch, roll, and yaw capability.
Figure 4: R-40B Specifications.
Figure 5: R-1E Specifications.

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Propellant Feed System
A pressure-fed cycle was selected for the feed system due to its simplicity and reliability. The
piping and instrumentation diagram for the engine is shown below in Figure 6 along with a legend
and the naming scheme. Shown in green is the composite overwrapped pressure vessel (COPV)
storing high pressure helium (He). During normal operation, the Helium Pressure Regulator (HPR)
delivery pressure is set to approximately 14.5 bar. Which is applied equally to both the MMH and
NTO within the tanks. The pressure is transmitted across a Fuel Check Valve (FCV1) and an
Oxidizer Check Valve (OCV1) to ensure no backflow of propellant into the pressurant lines, as
premature mixing of NTO and MMH would cause undesired and likely catastrophic ignition of
the hypergolic propellants inside the fluid lines.
Similarly, in the case of HPR malfunction or unexpected propellant boiling leading to tank over
pressurization, identical pressure relief valves (FPRV and OPRV) are included on each propellant
tank set to open at 20.5 bar, which results in a burst pressure factor of safety of 1.7.
Figure 6: Propulsion System P&ID.

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These pressure relief valves are redundant safety features, as the HPR already has redundancy
features. Tank filling occurs through the Fuel Fill Valve (FFV) and Oxidizer Fill Valve (OFV),
which additionally serves to prime the RCS propellant lines for minimal delay in reaction control
pulses upon launch. The Fuel Throttling Valve (FTV) and Oxidizer Throttling Valve (OTV) are
pressure drop inducing proportional control solenoid valves used to throttle the mass flow rate and,
ultimately, thrust from the main engines. There is no similar throttling valve incorporated for the
RCS system, as thrust control can be achieved by pulsing the engine’s valves. After the fuel and
oxidizer flow through the FTV and OTV, respectively, a tee fitting splits each line in two, one for
each main engine. Again, there are check valves redundant to those in the engine in each propellant
line connecting to the engines to ensure there is no backflow or premature mixing of the hypergolic
propellants. Similarly, there are redundant check valves in the lines that connects directly to each
RCS engines. Although insulation already protects the propellant tanks from experiencing
excessive rates heat transfer, the Fuel Pressure Transducer (FPT), Fuel Thermocouple (FTC),
Oxidizer Pressure Transducer (OPT), and Oxidizer Thermocouple (OTC) are identical components
that serve as redundant dual pressure and temperature transducers that will allow for verification
of the propellant’s liquid state.
Tank Selection
Trajectory analysis for a 660 km round trip showed a required Δv of nearly 3000 m/s and, using
the ideal rocket equation along with the engine’s specific impulse and initial vehicle weight, a
required propellant mass of 1244.6 kg is found. The NTO/MMH propellant combination has an
O/F ratio of approximately 1.6, so required NTO mass becomes 790.6 kg and required MMH mass
becomes 494 kg. Additionally, adding an extra 24 kg of NTO and 15 kg of MMH ensures the RCS
engines can have a combined total firing time of 1000 s. To house these propellants before
combustion in the engines, two identical Ariane Group 769 L spherical surface tension tanks will
be welded to the vehicle frame. Specifications for these tanks are displayed in Table 1.
Because the chamber pressure of the vehicle’s engines is low, a pressure vessel of relatively low
mass can be used to store and supply the inert gas that forces the propellants from the tanks to the
combustion chamber of each engine. This component was selected based on the desired delivery

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pressure to the engines and the tank volume, as the pressure within the tanks must remain
sufficiently high to account for major and minor head losses along the feed lines even as the
propellant tanks are near empty. From these considerations, the MT Aerospace PGV Family 75 L
HPV COPV was chosen. This pressure vessel will store 49.8 kg of helium at 310 bar to ensure
sufficient tank pressure during all stages of the mission. Helium was chosen as the pressurant over
nitrogen because it has a lower molecular mass. The datasheet for this COPV is displayed in Figure
7 below.
Table 1: Propellant Tank Specifications.

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Feed Line Selection
Because the selected propellants, MMH and NTO, are toxic to humans and corrosive, great care
was taken in selecting components to ensure material compatibility. Because the pressurant line
interfaces with the high pressure COPV, ½” stainless steel tubing was chosen for its high pressure
rating. Although stainless steel will corrode over time with exposure to either propellant,
compatible check valves ensure that these sections of the lines do not encounter the propellants.
The majority of feed lines in contact with the propellants were selected as aluminum 6061 1”
tubing, as it has excellent compatibility characteristics with both propellants. The only exception
is the use of ¾” polyethylene flex hoses for direct connections to the main engine, as gimbaling
requires flexible connections. Polyethylene is compatible with MMH and NTO, but these hoses
will be replaced every five uses to ensure excessive corrosion does not occur. This maintenance
also helps to ensure that excessive hose damage is not caused from regolith ejected by the engine
plume. Lastly, there is a small section of ¼” aluminum 6061 tubing for direct connection to the
RCS engine manifolds. Each line size was selected as a function of the expected mass flow rate
through that section of tubing.
Figure 7: Helium COPV Specifications.

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Feed Line Component Selection
Each of the valves in Figure 6 had to be selected according to pressure rating, rated flow rates,
and media compatibility. Because some components could only be sourced with undesirable
connection types, Swagelok will be the primary supplier of fittings and adapters to join
components within the feed lines. The component selected for the HPR was the Marotta
MFV400 for its high pressure rating, wide range of process fluid temperatures, and built in
redundancy features. Each of the check valves in the system, FCV1-FCV25 and OCV1-OCV25,
were selected as the Marotta CVM300 or CVM600 series for their material compatibility and
appropriate range of connection sizes. Next, the Valcor V3500-123-1-W will be used as both the
FPRV and OPRV for its appropriate range of cracking pressures and its compatibility with
MMH, NTO, and He. The dual and redundant temperature and pressure probes labeled in Figure
6 as the FPT, FTC, OPT, and OTC were selected as the GP:50 Model 343, which was selected
for its capability to measure pressure and temperature simultaneously. It can measure
temperatures below and above the freezing and boiling point, respectively, of both propellants
and pressures well above expected operating conditions. In both the fuel and oxidizer line, an
Ariane Group FDV series fill and drain valve will be mounted underneath the propellant tank.
This component is compatible with both propellants but must be replaced every 40 cycles as
specified by the manufacturer. Lastly, the FTV and OTV are pressure drop inducing throttling
valves used to control main engine thrust. No appropriate OTS components could be found, so
time and resources have been allotted in the schedule and budget to account for the design,
manufacturing, and testing of a proportional control solenoid throttling valve that meets the
requirements shown in Table 2.
Parameter Value
Pressure Range 0-42 bar
End Connection Size 1”
Material Construction Titanium
End Connection type Compression Fitting
Response Time < 250 ms
Power Requirements 70 W
Table 2: Throttling Valve Design Specifications.

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Feed System Layout
The rigid tubing of each feed line will be mounted to the frame of the vehicle where possible to
prevent vibrations or excessive stress on joints caused by the tubing and inline valves’ weight. A
crude representation of this configuration is shown in Figure 8 where it is seen that the feed line
geometry mimics that of the vehicle frame. One must note that the valves and most of the fittings
are not accounted for in this representation, as 3D models of these components were not
available. Additionally, the propellant lines for the main engines are not shown but will stretch
from underneath each propellant tank towards the center of the vehicle where the main engines
are mounted.
The tubing from the COPV will stretch from the underside of the vehicle to the top of each
propellant tank where the fuel and oxidizer pressurant connections are labeled in Figure 8. The
lines that transport the fuel and the oxidizer to the RCS engines are mirrored with respect to each
other. Each of these lines begin at the outlet of the tanks denoted as the fuel and ox tank
connection in Figure 8. They then stretch to the corners of the vehicle where the RCS engine
clusters are mounted, which is detailed in Figure 9.
Figure 8: RCS Cluster Connections.

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Mass Properties
The total wet mass of the propulsion system for a routine flight is found to be 1480.2 kg. For a
moon landing from lunar orbit, however, the total wet mass of the propulsion system becomes
1240.6 kg, as less Δv is required even after a 22% factor of safety is added. This difference in Δv
accounts for approximately 240 kg less propellant for moon landing versus suborbital journey. The
mass budget breakdown for the propulsion system in a landing from lunar orbit scenario is seen in
Table 3. This will be the propulsion system mass during takeoff from Earth.
Component Mass (kg)
Propellants 1045
Inert Gas (Helium) 49.8
Propellant and COPV Tanks 77.8
RCS Engines 24
Main Engines 21
Feed System 23
Total 1240.6
Table 3: Propulsion System Mass Breakdown for Lunar Landing.
Red: Oxidizer (NTO)
Teal: Fuel (MMH)
Figure 9: RCS Cluster Connections.

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Trajectory and Guidance
Overview
In accordance with ASD-SLT-019 and ASD-SLT-024, the SLT is required to have the ability to
land within 10 m of a designated landing point at a distance of up to 300 km from Shackleton
Crater Station. This designated landing point could potentially have hazards that will need to be
avoided in order to land safely. To achieve the goal of an accurate and precise landing, the SLT
will be outfitted with an attitude sensor system and accelerometers that will measure spacecraft
attitude and acceleration during the flight. Once the SLT is in the landing stage of the flight,
approximately 5 km above the lunar surface, the hazard avoidance system will activate. This
system will provide the SLT with a very detailed estimation of attitude and velocity in addition to
the attitude sensor system and accelerometers. Additionally, the hazard avoidance system will
discern all hazards in the area around the designated landing point to ensure a safe landing.
Trajectory Hardware
The hardware that will guide the SLT to its designated landing point is composed of three systems:
the star trackers, the gyroscopes, and the accelerometers. A star tracker is essentially a camera that
takes a picture of the stars in its field of view, then compares that picture to a known database of
star locations to determine the attitude and position of the spacecraft. The star trackers used by the
SLT are mounted at the top of the spacecraft to provide a clear, unobstructed view. The gyroscopes
measure the rotation speed at which the spacecraft is turning. In addition to providing the data
necessary for a properly timed execution of a rotational maneuver, the gyroscopes are also useful
for attitude measurement during quick-turning movements in which the star tracker is unable to
obtain an ideal image for database comparison. The accelerometers measure acceleration along an
axis, which is useful in telling the on-board computer when to start or end an engine burn.
For many spacecraft purposes, the gyroscope and accelerometer are combined into one system,
known as an Inertial Measurement Unit or IMU. However, this spacecraft will be using the Jena-
Optronik ASTROgyro, which instead combines the star tracker and gyroscope into one attitude

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sensor system. The ASTROgyro is composed of two star trackers and an inertial reference unit,
shown in Figure 10.
Figure 10: ASTROgyro unit with two star trackers (front) and inertial reference unit (rear).
The ASTROgyro automatically combines attitude data from the gyroscopes and star trackers to
provide a level of accuracy that would be difficult to obtain by each system alone. Additionally,
each unit has built-in redundancy, with two star trackers per unit and two gyroscope channels per
axis. However, the ASTROgyro is currently untested in a real-world mission scenario. Although
the testing is promising, the SLT will be outfitted with two complete ASTROgyro units due to this
unknown reliability. With these two units, the SLT will be guided by a total of four star trackers
and four gyroscope channels per axis.
The accelerometer chosen for the SLT is the Honeywell Q-Flex QA-3000-030, shown in Figure
11.
Figure 11: Honeywell Q-Flex QA-3000-030 accelerometer.

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This accelerometer was chosen due to its high measurement range, high sensitivity, and especially
its high turn-on repeatability. Turn-on repeatability refers to the ability of the accelerometer to
output the same measurement if repeatedly placed in a similar environment. In other words, this
accelerometer model will be consistently accurate over the course of the SLT’s launch and landing
life cycle. The SLT will have two accelerometers per axis, meaning that even if an accelerometer
is nonfunctional on every axis, the SLT will still be able to take accurate acceleration
measurements.
Hazard Avoidance System
The Hazard Avoidance System (HAS) is composed of three systems: The High-Altitude Laser
Altimeter (HALA), the Navigation Doppler LiDAR (NDL), and the Hazard Avoidance LiDAR
(HAL). The HALA provides accurate line-of-sight range measurements at approximately 20 km
above the surface, beginning the terrain-relative navigation segment of the landing. Once the SLT
reaches an altitude of 4 km, the NDL will use flash LiDAR technology to begin providing velocity
vectors, altitude, and 2-D ground relative attitude. The NDL optical head is placed on the outside
of the SLT to give as unobstructed a view as possible, while the chassis is placed in the electronics
bay. These two components are shown in Figure 12.
Figure 12: Navigation Doppler LiDAR optical head (left) and electronics chassis (right).

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Finally, once the SLT reaches altitudes of less than 1 km, the HAL uses scanning LiDAR
technology to map out the area around the designated landing point, then uses that map to choose
the most ideal safe landing site. Each of these systems have undergone extensive NASA
development and testing over the course of the last decade through the Autonomous Landing
Hazard Avoidance Technology (ALHAT) project, the CoOperative Blending of Autonomous
Landing Technologies (COBALT) project, and the Safe and Precise Landing Integrated
Capabilities Evolution (SPLICE) project. Because of this, the systems on the SLT will have a high
reliability. Additionally, during the landing stage of the flight, the three subsystems in the Hazard
Avoidance System will not only overlap with each other but will also be supplemented with data
from the ASTROgyro attitude sensor system and the accelerometers. This gives a total of six
systems feeding attitude and position data to the spacecraft, allowing for a safe and accurate
landing. A diagram of the landing sequence detailing the use of each system is given by Figure 13.
Figure 13: SLT landing sequence diagram showing when each guidance system is active.
The HAL is mounted on a swivel on the side of the SLT. As the SLT begins its landing sequence,
the ability to swivel allows the HAL to keep the area around the designated landing point in view
as it performs the procedures necessary to properly determine the safest and most fuel-efficient
landing area. First, the HAL examines the roughness of the terrain around the designated landing
point. The slope of the terrain in this area is then calculated. With the roughness and slope
determinations, the HAL is then able to calculate the fuel cost to land at any given area around the

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designated landing point. Finally, the ideal landing point is chosen based on this cost estimation.
Figure 14 shows examples of the maps created by the HAL as it performs these tasks.
Figure 14: Hazard Avoidance LiDAR safe-site determination example maps.
Maximum Range Example
The SLT has a maximum operational range of 300 km. At this range, the spacecraft is subject to
the highest likelihood of guidance inaccuracy during tracking. To determine the maximum
inaccuracy of the trajectory, a sample calculation must be performed to determine key aspects of
the launch and flight of the SLT. The first step is to determine θ, which is half of the angle travelled
over the lunar surface. This was determined using (1), where d is the maximum range of 300 km
and R is the radius of the Moon, 1737.1 km.
𝜃 =
45° ∗ 𝑑
2𝜋𝑅
= 4.958° (1)
Using θ, the Δv required at launch to travel 300 km on an ideal trajectory could be determined
using (2), where the lunar gravitational param μ = 4.9048695 x 1012 𝑚3
𝑠2 . This value is equal to the Δv
required for an ideal landing on this trajectory.

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∆𝑣𝑙𝑎𝑢𝑛𝑐ℎ = ∆𝑣𝑙𝑎𝑛𝑑𝑖𝑛𝑔 = √
2𝜇 sin(𝜃)
𝑅(1 + sin(𝜃))
= 669.6
𝑚
𝑠
(2)
To meet ASD-SLT-020, the SLT must be able to perform at least 1.1 maximum range round-trips
without refueling. The total Δv required for a round trip, including the reserve Δv beyond that
needed for an ideal launch and landing, was calculated using (3) and (4).
∆𝑣𝑟𝑒𝑠𝑒𝑟𝑣𝑒 = 0.1(∆𝑣𝑙𝑎𝑢𝑛𝑐ℎ + ∆𝑣𝑙𝑎𝑛𝑑𝑖𝑛𝑔) = 133.9
𝑚
𝑠
(3)
∆𝑣𝑡𝑜𝑡𝑎𝑙 = 2(∆𝑣𝑙𝑎𝑢𝑛𝑐ℎ + ∆𝑣𝑙𝑎𝑛𝑑𝑖𝑛𝑔 + ∆𝑣𝑟𝑒𝑠𝑒𝑟𝑣𝑒)
= 2(2∆𝑣𝑙𝑎𝑢𝑛𝑐ℎ + ∆𝑣𝑟𝑒𝑠𝑒𝑟𝑣𝑒)
= 2946.24
𝑚
𝑠
(4)
The time of flight of the SLT for a maximum range launch assuming impulse burns for launch and
landing was calculated with (5), where the lunar gravitational constant G = 1.62
𝑚
𝑠2
.
𝑡𝑖𝑚𝑝𝑢𝑙𝑠𝑒 = ((
1 + sin(𝜃)
2
)
3
2
𝑎𝑟𝑐𝑐𝑜𝑠 (
cos(𝜃)
1 + sin(𝜃)
) +
1
2
cos(𝜃) √sin(𝜃)) (2) (√
𝑅
𝐺
) = 642.76 𝑠 (5)
The engine burn time for ideal launch or landing was determined using (6), where the mass of the
vehicle at launch mvehicle was estimated as 2000 kg and the thrust provided from the two R-40B
engines Tvehicle = 8000 N.
𝑡𝑏𝑢𝑟𝑛 =
∆𝑣𝑙𝑎𝑢𝑛𝑐ℎ ∗ 𝑚𝑣𝑒ℎ𝑖𝑐𝑙𝑒
𝑇𝑣𝑒ℎ𝑖𝑐𝑙𝑒
= 167.4 𝑠 (6)
The total time for a trajectory that travels 300 km over the lunar surface, with ideal launch and
landing burns, is then calculated using (7)-(9).
𝑑𝑎𝑟𝑐 = ∆𝑣𝑙𝑎𝑢𝑛𝑐ℎ ∗ 𝑡𝑏𝑢𝑟𝑛𝑜𝑢𝑡 = 430.39 𝑘𝑚 (7)

25
𝑑𝑏𝑢𝑟𝑛 =
1
2
𝑎𝑡2
=
1
2
(
8000 𝑁
2000 𝑘𝑔
) (167.4 𝑠)2
= 56.046 𝑘𝑚 (8)
𝑡𝑡𝑜𝑡𝑎𝑙 =
𝑑𝑎𝑟𝑐 ∗ 2𝑑𝑏𝑢𝑟𝑛
∆𝑣𝑙𝑎𝑢𝑛𝑐ℎ
+ 2𝑡𝑏𝑢𝑟𝑛 = 810.15 𝑠 (9)
These calculations will be performed by the onboard computer as necessary when given a
designated landing point, and results will vary based on the distance of the landing point from
Shackleton Crater Station. However, using the results of these maximum range launch
calculations, the ideal trajectory can be seen in Figure 15.
Figure 15: Ideal maximum range launch trajectory. A 300km circle is shown in yellow around Shackleton Crater Station to
illustrate the maximum SLT launch range.
Under nominal operation of the ASTROgyro, the spacecraft will have an attitude estimation
inaccuracy of less than 0.0003º. If the gyroscopes are inoperable for any reason and only the star
trackers can be relied upon, the attitude inaccuracy will be less than 0.0014º. If the star trackers
are not functioning properly and only the gyroscopes can be used, the attitude inaccuracy drift will

26
be less than 0.05º per hour. To determine the ability of the spacecraft to land within 10 m of the
designated landing point, as required by ASD-SLT-024, the change in distance travelled due to
potential inaccuracies in data measurement was calculated using (10), where γ is the flight angle
relative to the local horizontal (the lunar surface). The flight angle is the independent variable that
will shift due to the hardware inaccuracies.
𝑑 = Δ𝑣𝑙𝑎𝑢𝑛𝑐ℎ ∗ cos(𝛾) ∗ (𝑡𝑖𝑚𝑝𝑢𝑙𝑠𝑒 − 𝑡𝑏𝑢𝑟𝑛) (10)
The calculations assumed that the SLT would receive no control input after the initial launch burn.
In other words, once the spacecraft was launched and was affected by any inaccuracy during that
launch, the accuracies calculated are a measurement of how far away from the designated landing
point the SLT would land with no other input. It was determined that a nominally operating
ASTROgyro will guide the SLT to a landing point within 1.038 m of the designated landing point.
Reliance on star trackers alone will guide the SLT within 4.788 m, and gyroscopes alone will guide
it within 7.952 m.
These results prove that the SLT is more than capable of landing within the required 10 m area
around a designated landing point, even if an entire guidance system is nonfunctional.
Additionally, the HAS meets a NASA requirement to land within three m of the site that the HAL
chooses as the ideal safe landing point.

27
Communication and Tracking
Overview
Given ASD-SLT-025, the communication system was required to be able to transmit and receive
data and commands with SCS and remote sites either by line of sight or via relay satellite/tower as
appropriate for the situation utilizing X-band radio. It was determined that this requirement also
implied that in addition to a communication device such as a transceiver, an appropriate On-Board-
Computer (OBC), and vehicle tracking system would need to be selected as well. These derived
requirements were deemed essential to ensure that the SLT can maintain constant communication
with operational controllers at all phases of the SLT mission, as well as monitor vehicle health,
location, and issue updated commands to the SLT if needed. As discussed in the Trajectory and
Guidance section of this paper, the ASTROgyro star tracker and gyroscope unit were selected for
tracking, so the focus of this section will be on the transceiver and On-Board-Computer selected
for the SLT.
Transceiver
After an industry survey of several Off-The-Shelf (OTS) communication devices, it was
determined that a compact transceiver would be the best choice for the SLT. With this information,
various transceiver setups were compared based on their performance, size, and level of flight
testing. It was determined that the X-Link Transceiver by IQ Spacecom was the best option for
this application (shown in Figure 16). This transceiver is flight tested, has a very compact form
factor, uses X-band radio, and has a built-in patch antenna for receiving commands. Table 4
contains an outline of this devices key specifications.
Figure 16: X-Link Transceiver by IQ Spacecom.

28
Parameter Value
Downlink (SLT to Ground) Up to 100 Mbps
Uplink (Ground to SLT) 64 kbps +
Power Consumption 15W (Max)
Mass (grams) Less than 200g
Size (mm) 96 x 65 x 28 mm3
Price (USD) $53,152
Additionally, it should be noted that two transceivers will be mounted to the SLT design. This was
done to ensure redundancy, in the case that one transceiver is not responding to or receiving
transmissions. The on-board-computer system would be able to detect an error with the transceiver
and switch over to the functioning transceiver for the duration of the mission. The risk assessment
for this component can be found under the risk assessment for guidance and control in Appendix
A.
On-Board-Computer
The on-board-computer (OBC) was selected in a similar method to the transceiver. A comparison
was made between several different off the shelf (OTS) on-board computer systems, and the best
option was selected. The OBC selected was the EXA ICEPS Spacecraft System Core, shown in
Figure 17. One major plus about this OBC is the fact that it is the main computer system for the
peregrine lander, scheduled to launch on July 2021. Important specifications for this On-Board-
Computer can be seen in Table 5. The risk assessment for the OBC can be found under the risk
assessment for electrical in Appendix A.
Table 4: Transceiver Specifications.
Figure 17: EXA ICEPS Spacecraft System Core.

29
Component Specification
RAM 512 MB of DDR3L
Storage 32 GB SSD
CPU
Dual-core ARM Cortex
A9 (runs up to 733
MHz)
Price
(USD)
$34,309.46
In addition to the selection of the OBC, a plan was developed for its redundancy and storage. As seen in
Figure 18, there will be three on-board-computers placed in one housing. This housing is constructed of
7075 – T6 Aluminum construction and will be fastened onto the electronics bay with the other electronic
systems. This OBC array offers redundancy through both the software designed parity, as well as backup
hardware. As shown in Figure 19, if a bit flip or other error is detected by an OBC, it will be taken out
commission for the duration of the flight. Upon landing the OBC that incurred a problem can easily be
replaced by an astronaut.
Table 5: On-Board-Computer Specifications.
Figure 19: Redundancy of the On-Board-Computers, Diagram
Figure 18: On-Board-Computer Array Housing.

30
Hardware and Electronics
Overview
Beyond established requirements that the SLT must use MIL-DTL-38999 connectors for power
interfaces, connect with the SLS Dispenser for electrical power and bonding/grounding, and
provide 28V, 15A of electrical power to payloads for the duration of the payloads mission, several
additional requirements were self-imposed to ensure that the SLT could perform its mission
optimally. The battery system used for power storage was required to be able to provide full power
at peak power draw to the SLT for the entirety of its round-trip flight. The solar panels used for
long term power generation on the SLT were required to be able to endure the harsh environments
of space and provide over 100% of the SLT’s nominal power requirement, such that the excess
power could be used to recharge the battery system. Finally, the gimbaling system for the main
engine was required to provide up to 6° of rotation to the main engine.
Electrical Inventory
An inventory of the components of the SLT requiring electrical power, featuring the peak power
draw of each unit, the number of units, and the total power draw of the system is shown in Table
6 below.
Item Power/Unit (W) Quantity Power (W) Notes
Payload 420 1 420 1 way
Xlink Transceiver 15 2 30
R-40B valves 70 2 140
R-1E valves 36 12 432
Main Engine Proportional Control Valve 70 1 70 estimated
COPV Gas Regulator 118.4 1 118.4
OBC 100 3 300
Accelerometer 0.48 6 2.88
ASTROgyro 54 2 108
Gimbal Actuators 240 4 960

31
HALA 15 1 15 estimated
NDL 80 1 80
HAL 80 1 80 estimated
Pressure/Temperature Transducer 1.0625 4 4.25 estimated
Polyamide coil film (for ASTROgyro) 80 1 80 estimated
Table 6: Inventory of items requiring electrical power.
The peak power requirement of the SLT to and from a mission as well as the nominal power
requirement, calculated as the SLT on the ground so no power is required by the propellant valves
or the gimballing actuators, is given in Table 7 below.
PEAK POWER TO (W): 2840.53
PEAK POWER FROM (W): 2420.53
NOMINAL POWER (W): 1133.53
Table 7: Peak and nominal power requirements of the SLT.
Solar Panels
The first step for solar panel selection was to select the type of solar cells to be used between thin
glass silicon and multi-junction gallium arsenide, the two types typically used for space
applications. An analysis based on panel efficiency, weight, cost, and thermal conductivity was
conducted and determined that, despite the significantly higher cost, multi-junction cells were ideal
for this application because of their low weight and thermal conductivity and high efficiency.
Multi-junction cells are also the modern industry standard and a rapidly improving technology,
making them an ideal choice for long term viability.
From this, several potential panel arrays from NanoAvionics, designed to provide over 1133.53W,
were compared based on their power, mass, and area to determine the most ideal arrangement for
the SLT. After analysis, the 3U-FMmT Solar Panels wired together in an arrangement of 2 in series
and 78 in parallel (156 total) was determined to provide the ideal balance of high power, low
weight, and low area. For mounting purposes, the overall panel arrangement was split into four
arrays for space considerations when mounting on the SLT.

32
Batteries
When investigating the power storage system for the SLT, Lithium Cells were selected for their
high power density and better performance relative to Nickel cells and their status as a rapidly
improving technology, making them a good choice for long term viability. It was also known that
the batteries needed to have the ability to store 2840.53W at 28V for 13.5 minutes one way and
2420.53W at 28V for 13.5 minutes back, giving a total capacity requirement of 1183.7Wh or
39.99Ah. Knowing this, a set of battery systems were compared based on their capacity, weight,
maximum power output, and volume, resulting in the selection of a system of 2 Space Vector 25Ah
Li-ion Polymer 39381 Series Batteries, shown in Figure 20 below.
Figure 20: Space Vector 25AH Li-Ion Polymer Battery 39381 Series.
The 39381 uses D38999 connectors, in compliance with the established mission requirements. As
the overall battery system provides 50Ah of capacity, this gives a factor of safety of 1.25 over the
established capacity requirements of the SLT.
Gimbal Actuators
To meet the gimballing requirements for the SLT, a series of linear actuators were compared based
on the max load, speed, and weight characteristics, as well as their cost. After the comparison, the
Progressive Automations 2” stroke heavy duty linear actuator, shown in Figure 21 below, was
determined to be the best option for the SLT.

33
Figure 21: Progressive Automations 2” stroke heavy duty linear actuator, selected for use on the SLT gimballing system.
The actuator can achieve a full gimble angle in 0.44 seconds, has a maximum dynamic load of
3780N, and a full-load actuating speed of 13mm/second. With an effective temperature range of -
25°C to 40°C, the actuators will need to be insulated to perform effectively in the environment at
Shackleton Crater.
Power Distribution
For the power distribution system of the SLT, the TERMA Space power condition & distribution
unit, shown in Figure 22, was selected. The system can handle both power conditioning and
distribution. The system can also handle an input power and voltage of 3000 W and 50 V
respectively, both higher than the SLT’s peak requirements of 2840.53 W and 29.6 V. The system
has a low mass at 16.3 kg and a high transfer efficiency, as well as proven reliability in space, as
they had previously been used on the European Space Agency’s Galileo Spacecraft.
Figure 22: TERMA Space power conditioning & distribution unit selected for use on the SLT.

34
A diagram for the overall power distribution of the SLT is shown in Figure 23 below. Power flows
into the PCB from either the batteries or the solar panel array and then is distributed out to the
electrical components of the SLT. When power is coming from the solar panels, power flows to
the batteries to recharge them for later use.
Figure 23: Power distribution diagram for the SLT.
Cable Mass
The SLT’s total cable mass was calculated from the expected cable length for each component and
the cable gauge width necessary to meet the current requirements of each component, based on
selections from the GORE Space Cables catalogue. In total, the cable mass for the SLT was
0.136kg.

35
Environmental
Overview
Traveling and working on the moon brings about dramatic environmental concerns that must be
mitigated in order to ensure successful operation of the SLT. The lunar transport must be qualified
for all thermal, vibrational, shock and acoustic loads in adherence with ESD 30000 Mission
Planner’s Guide. The SLT will use a combination of passive and active elements to help mitigate
all environmental concerns. Passive methods will be emphasized as much as possible to limit
complexity of design and power consumption. An in depth discussion of mission environments
paired with all control methods used on the SLT for thermal, vibrational and acoustic conditions
will be presented below.
Thermal
The moon poses very harsh thermal conditions that the SLT must be capable of dealing with. On
top of being able to safely traverse the journey to the moon, the SLT must also be capable of
surviving up to 72 hours in the harsh lunar night conditions as well as expel excess heat from all
interior electronics during lunar day conditions. Below is a table that shows all the predicted
thermal environments for the duration of the mission.
Pre-Launch
0 - 27°C. All facilities are climate controlled per ESD30000
standards and temperature range is nominal.
Launch
0 - 27°C. SLT is encapsulated and in a thermally controlled
payload fairing. Vibroacoustic effects are primary concern
during launch.
Cruise to Moon -40 - 60°C. Depending on whether in sunlight or shadow
Lunar Orbit 100° C.
Shackleton Crater
Surface
-183 - 120°C. Dependent upon whether SLT is under shadow.
Table 8: List of expected thermal environments.

36
It can be seen that the SLT can expected temperatures throughout the mission ranging from -183
- 120°C. This large fluctuation in temperature alone would be enough to cause catastrophic damage
to several systems of the SLT included propulsion and interior electronics.
SolidWorks transient studies were done on the dinitrogen tetroxide fuel tank to assess exactly how
quickly these two extreme temperatures could affect the oxidizer. NTO was chosen for this study
because it had a much more sensitive operating temperature range. Due to the nature of how heat
transfers in space, Radiation will be the primary source via which heat gets expelled or absorbed.
For this study, the emissivity of the titanium tanks was taken into account, an initial nominal
temperature of 15°C was chosen, and the maximum & minimum lunar temperatures were taken
into account. Results of this study are shown below in Figure 24.
Figure 24: Screenshot of the SolidWorks thermal study.

37
Figure 25: SolidWorks thermal analysis of NTO tank given coldest condition of -183 °C. Starting temp. 15 °C. Emissivity of
tanks: 0.5 NTO freezing point: -11.2 °C. Time to freeze w/o insulation: 40 min. Time rate change: 0.015 °C/s.
Figure 26: SolidWorks thermal analysis of NTO tanks given hottest lunar condition of 120 °C. NTO boiling point: 21.1 °C Time
to boil: 2 min.
Several tools were considered to mitigate this risk but after trade studies, it was chosen that MLI
(multi-layer insulation) was the most effective tool at protecting the SLT from these temperature
changes. MLI is made of several layers of protective film that limits the rate at which radiative
heat transfer can occur. Research was done to assess the best materials for insulation as well how
much was needed and its effectiveness. Ultimately, it was decided that layers of kapton and mylar
would be used to construct the multi-layer insulation. kapton is a thin polyimide gold plated film
that was developed by Dupont in the late 1960s and is able to remain stable in temperature ranges

38
from -269 to +400°C. It is regularly used as an insulator in vacuums due to its light weight, thermal
conductivity, dielectric qualities and low outgassing rates. Mylar has very similar properties with
temperature stability in the range of -70-180°C and is flexible while maintaining a high tensile
strength. kapton was chosen as the outermost layer of the MLI since both properties had relatively
similar qualities in tensile strength but kapton had a slightly greater temperature range to which it
could exhibit strong mechanical properties. The appearance of the final SolidWorks assembly was
changed to simulate MLI. As can be seen in the following figure, all elements in gold/bronze are
encased in MLI. It is worth noting that the tanks and electronics are fully encased. Furthermore,
the landing pads and hinges of the landing gear are also enclosed to help protect against any
potential damage from regolith upon SLT launch and land. It was found that the MLI could be
provided by RUAG, Europe’s largest manufacturer of insulation.
Figure 27: Image of SLT showcasing MLI.
In order to determine how many layers of MLI would be needed, a lumped system analysis was
used. The equations for this analysis are below.

39
(11)
(12)
From these formulas, it could be seen that the rate of radiative heat transfer was dependent upon
the thermal resistance equation which in itself depends on the value Ɛ, or emissivity. Through
research, the emissivity was found to change depending on the number of layers of MLI. A
theoretical curve showing the rate of change of the emissivity depending on layers of MLI is seen
next.
Figure 28: Theoretical curve of effective emissivity given number of layers.
From this data and the given equations, a trial & error process occurred to determine the optimum
number of MLI layers. 40 layers of insulation were found to be the optimum amount due to the
effective emittance of 0.001. When substituting this value into the thermal resistance equation and
then solving for effective radiative heat transfer, it was found that this number of layers blocked
over 99% of heat transfer. This insulation decreases the time rate change of temperature at the

40
moon’s lowest temp from 0.015 °C per second to 0.000075. The time it takes for fuel to freeze
now increases from 40 to 133 hours or more than 5 and a half days.
Given the number of layers of MLI, it was important to then determine the mass that this would
add to the total SLT. Mass approximations for mylar and kapton were taken from NASA Multi-
layer insulation material guidelines. From this guideline, it was seen that that kapton has a mass
of 19
𝑔
𝑚2 and mylar has a mass of 17
𝑔
𝑚2. Furthermore, given the size of the SLT and the parts of
the vehicle requiring insulation, it was estimated that 38 m2
of insulation would be required. Given
20 layers each per material and taking into account the amount of material per layer and number
of layers, the MLI resulted in a mass of 25.2 kg. This MLI would be attached to the SLT via hook
& pile fasteners and metal grommets at corners to allow for ease of assembly and
manufacturability. Fully suited astronauts would be able to peel off the MLI from hook & pile
fasteners to get to the interior payload fairly easily.
The next major component of thermal control was radioisotope heater units. Powered by the
radioactive decay of plutonium-238, these would be able to provide the small amount of additional
heat needed to protect all internal components from coldest lunar night conditions. Weighing only
40 g and providing 1 W of heat each, these heater units are versatile members that can be placed
anywhere additional heat is needed. Sealed in protective pouches, the risk of radiation is
completely nullified. It was determined that 25 heater units would be sufficient to satisfy internal
heat concerns of the SLT. This would be enough to provide 10 units to each propellant tank and 5
units for the COPV tanks. A visual depiction of these RHUs is presented below.
Figure 29: Radioisotope heater unit.

41
Several SLT instruments such as the ASTROgyros were required to be placed outside of the
protective layer of MLI and exposed in order to gather their necessary data. For these components,
it was necessary to find a new way to provide heat without interfering in the operation of the
components. For this, Polyimide patch heaters were selected. These heaters are made with electric
coils sandwiched between 2 layers of kapton. The coils within the kapton will be powered by the
onboard batteries and will be able to provide heat via conduction to all exterior electronic
components. This heating tape will be provided by Omega Engineering and would add under 100
g worth of mass to protect exterior electronics. Unlike the MLI and the RHUs, this is an active
heating component.
Several control methods have been discussed for how to provide heat and limit heat transfer. It is
also important though, to be able to rid interior heat from the SLT. The MLI and RHUs can be so
effective at their job that the interior electronics would be incapable of cooling down on their own.
Because of this, it becomes necessary to include a cold plate attached to the electronics. This cold
plate works by absorbing electrical waste heat and then dissipating it through flow paths.
Heat pipes would be the designated flow path for these cold plates. The heat pipes are capable of
transferring the heat built up in the cold plates. The pipes work by having an interior liquid that
will come in contact with the heat source. This liquid, typically ammonia due to its thermal
properties, would then turn into a vapor and travel along the heat pipe to the other end where the
excess heat could potentially be expelled by a radiator. The ammonia then turns back into a liquid
and starts the process all over again. This can occur indefinitely and passively to continually
prevent interior heat build up. The heat pipes will be made with 12 mm aluminum tubes filled with
liquid ammonia at a mass of 0.36
𝑘𝑔
𝑚2. This would add an additional mass of 5.3 kg to the SLT.
Figure 30: Cold plate with heat pipes attached.

42
The heat pipes must then be connected to a tool that can take the excess heat received from the
cold plates and expel it. For that, a radiator will be used. A teflon coated radiator would be capable
of expelling up to 300W of heat per square m that is exposed to space. The current SLT
configuration has a radiator attached to a side panel of the SLT behind one of the solar panels. It
is approximately 1.6 m2
in size and comes in at a mass of 7.57 kg.
Figure 31: Teflon coated side panel radiator.
One final thermal component was needed to ensure satisfaction of all thermal control criteria. A
louver was mounted on top of the exterior radiator. The purpose of a louver is to limit the amount
of the radiator that is capable of seeing and expelling to space. This ensures that not too much heat
is expelled at any given period and protects the radiator if it is pointed in the direction of the sun
for a long period of time. The louver can open and close completely passively through the use of
bimetallic springs. Bimetallic springs couple two different metals with distinct rates of thermal
expansion and converts that into mechanical energy without any necessary power. When the
radiator begins to heat up due to the heat pipes, the springs will begin to expand and open the
Louver flaps, allowing the radiator cool back down. Once cool, the springs contract again and
close the flaps of the louver. This ensures that heat is always being transferred out exactly at the
rate that it should be. The following figures will provide visual representations of the
Louver/radiator system on the SLT.
Figure 32: Bimetallic springs.

43
Figure 33: Louver assembly.
Figure 34: SolidWorks model of Louver in its open and closed states.

44
Figure 35: Louver/radiator assembly modeled onto SLT full assembly.
The full array of thermal control elements has been presented, now vibrational and acoustic
considerations are of importance. Vibrational and acoustic forces would by far be the strongest
during SLS liftoff and the initial transonic portion of the flight. It is imperative, during these times,
to have control methods that would be capable of limiting the intense forces from mission
operations. The acoustic regime could be mitigated primarily via acoustic foam and vibrations
could be stopped with isolators.
Acoustic
As was previously mentioned, the SLT will experience the most acoustic excitation during liftoff.
To mitigate this, NASA has its own sound overpressure suppression system. This system quickly
expels millions of gallons of water to limit the sound waves happening during launch from
interfering with the actual payload. For redundancy though, acoustic foam should be applied along
the interior surface area of the payload fairing. Typically, it is the responsibility of the launch
provider to provide sound abatement. It was found though that melamine foam and foams with
similar performance characteristics would be compatible with the SLT. These foams limit mass
due to their low density, while still fulfilling the acoustic damping requirements of the SLT. Below
is a chart showing the sound absorption coefficient of Melamine foam depending on the frequency
of sound in Hz. It is estimated that the foam should be 50.8 mm thick. If the launch provider were
to select Melamine, this would result in an additional mass of 51.56 kg.

45
Frequency (Hz) Absorption coefficient
125 0.1
250 0.22
500 0.54
1000 0.76
2000 0.88
4000 0.93
Table 9: Melamine ultra lite foam sound absorption coefficient.
Figure 36: SolidWorks representation of the acoustic foam within the fairing.
Vibration
Vibration was the final consideration that had to be taken into account. To limit the potential
negative effects of vibration on the SLT, Moog shock isolators were installed underneath the
interior electronics. These isolators are capable of damping accelerations due random vibration
within fractions of a second and decrease the frequency of vibrations by an order of magnitude.
Data created by the manufacturer was found and used as a reference for performance analysis, as
seen in Figure 37.

46
Figure 37: Manufacturer’s data. On the left is graph of acceleration damping from random vibrations and on the right is shows
the difference in frequency of vibration experienced with and without vibrators.
As was previously mentioned, these isolators will be placed underneath the interior electronics of
the SLT and protect it from any random vibrations. Moog shock isolators, seen in Figure 38, weigh
80 g each. With 4 being used, this will add an additional 320 g.
Figure 38: Moog shock isolators and where they will be placed within the SLT.

47
Mechanisms and Structures
Overview
The two main derived requirements from the given requirements (ASD-SLT-09 to ASD-SLT-015)
were that the SLT must fit inside the 8.4 m SLS capsule for delivery and that the landing gear and
base structure mass must not exceed 450 kg. The structure of the SLT was driven by needing a
low mass but high strength structure that could withstand the forces brought onto the SLT during
operation. Due to this, 7075-T6 Aluminum alloy was used for the base structure of the SLT, as
well as most other parts including the landing gear struts, footpads, and other mounting brackets
and parts. The base structure is made of an aluminum truss structure, while there are some
connecting parts that are made of Ti-6Al-4V Titanium alloy where higher strength and corrosion
resistance was needed. The landing gear also used Aluminum Honeycomb structure for
compression.
Base Structure
The base structure is made of thin-walled hollow rectangular truss members, welded together at
the ends with added members for mounting of the payload. Similarly, welded members are located
above the payload area to house the electronics. The base structure can be seen in Figure 39 below.
Figure 39: Base structure of SLT.

48
Landing Gear
The landing gear struts are made of thin-walled hollow cylindrical tubes, which are welded from
the base structure to the landing gear mount, with the exception of the primary strut which is
welded to a titanium connector part. The landing gear can be seen below in Figure 40. Note that
the landing gear struts are mounted to an aluminum landing gear mount, and that the struts are
statically connected by welding, so they do not rotate or translate in any way.
Figure 40: Landing gear on one side of the SLT.
The two vertical struts under the landing gear mount house an aluminum honeycomb cartdridge to
allow for compression of the SLT upon landing, which are connected to the footpad on a rotational
hinge that would allow the SLT to land on varying elevations and on slopes. The vertical struts are
specifically thin-walled hollow square cross sections to accommodate the honeycomb structure.

49
SLT Structure and Mounting
From Figure 41, the mounting of various components of the SLT can be seen. The main propellant
tanks are mounted on opposite corners, welded to the side support beams of the base structure. The
two R-40B Engines are mounted on the two corners where the propellant tanks are not mounted,
with the gimbal assemblies mounted under the two beams that support the payload. There is a
cluster of 3 RCS engines mounted on the corners without propellant tanks as well, with the total
number of RCS engines at 12, allowing movement and rotation in all three X-Y-Z directions.
Figure 41: SLT isometric view without insulation panels.
The green rectangle in Figure 41 are the vertical truss members that the propellant tanks and
payload support beams are welded to.
The payload is mounted on the support beams as shown in Figure 42. The figure shows the SLT
with MLI panels hidden, but the astronauts should be able to easily retrieve the payload by lifting
up the MLI panels on this side of the SLT. The electronics bay is mounted above the payload on

50
similar support beams, which are welded to the vertical truss members of the base structure. The
COPV tank is welded below the payload to a support mount. The red rectangle is where the COPV
tank and the two main engines are mounted under the payload. The blue rectangle is where the
electronics are mounted on top of the payload.
In Figure 43, the solar panels and insulation panels are shown back on the SLT. Three solar panels
are mounted on the three struts where the Hazard Avoidance LiDAR does not block it, and the last
solar panel is mounted on top of the SLT on the thin mounting plate that is shown. It can also be
seen that the Gator Hopper logo is clearly visible on many surfaces on the SLT.
Figure 42: Side view of the SLT.

51
Figure 43: SLT isometric view with solar panels and insulation.
Gimbal Assembly
Most off the shelf gimbal bearings are not meant to operate below -20°C, so a custom gimbal
bearing will be developed for the Gator Hopper SLT. The custom gimbal bearing uses a ball and
socket gimbal structure, which was chosen to allow all degrees of freedom while keeping the
design simple. This type of gimbal bearing will also take off some of the loads off the electrical
actuators. The gimbal assembly can be seen in Figure 44.
Figure 44: Gimbal Assembly with engine.

52
The vertical bar above the thrust chamber allows for the propellant feed systems to mate with the
engine. The gimbal bearings will use teflon fiberglass inserts for lubrication between the ball and
socket, which have coefficients of friction of 0.05 to 0.10, which is the third lowest of any solid
material. This material was also used for lubrication on the Saturn V gimbal bearings.
Fabrication and Manufacturing
Most of the SLT and its components were chosen to be welded together. This was due to how
efficient welding is for the purpose of a long-term operation like ours, knowing that the risk of
failure is low. The biggest downside to this however is that welding aluminum can be difficult;
due to this we have made sure to include ample time and money to the schedule and budget to
ensure that the aluminum parts included can be welded together correctly.
There are also some parts of the SLT that are secured by fasteners, namely the footpads and the
primary landing gear strut. We believe that these fasteners will be big enough for astronauts to
repair easily on the moon, but we have added ample time to the schedule for testing to be done on
the fasteners with astronaut suits.
For the custom designed gimbal assembly structure, specifically the large ball and socket bearing
which takes the load off the gimbal actuators, it will be manufactured using a casting method with
a thin sheet separating the ball from the socket. This will allow the structure to be stronger than if
it were bolted together.
Finally, the titanium parts will be hard to machine but we have added ample time to the schedule
to account for the time it would take to reach out to contractors that specialize in manufacturing
with titanium alloy and to allow time for the manufacturing to complete.
Stress Analyses
The two main components analyzed for the stress analysis was the aluminum bar that supports the
payload, and the aluminum honeycomb structure that performs the compression for the SLT.

53
Figure 45: Stress analysis on payload support beam.
Figure 45 shows the stress analysis done on the payload support beam. The max force applied to
the SLT was assumed to be the weight of the entire SLT plus two RCS thrusters firing down,
possibly malfunctioning. This is the worst-case scenario, at around 3800 N. Since the entire SLT’s
weight does not actually act on this support bar, the weight is distributed along the length of the
beam, depicted by the yellow arrows in Figure 45. Note that the beam is welded on both ends.
Along with the distributed weight, there is an engine on this beam that fires upwards with a force
of 4000 N, distributed along a small circle at where the red arrow points on Figure 45. With a
design factor of 1.4, from the stress analysis done in SolidWorks it can be seen that the member
does not come close to yielding, with a safety factor of almost 10. Since the design factor is 1.4,
this design could be adjusted to be even lighter, as the excess strength is not needed. The walls of
the aluminum members are already thin, but some time will be allotted in the schedule during
testing to possibly reduce the size of the SLT members even further. Note that the red areas in
Figure 45 are a local maximum stress, not an ultimate stress, and the safety factor was taken at
those red areas.

54
Table 10 shows the compressive strengths of various different aluminum honeycomb structures
that are manufactured by Plascore, Inc. Using the same max load of about 3800 N, divided by 4 to
split up to each landing strut and dividing over the cross section of the aluminum honeycomb, the
compressive strength is 0.092 MPa, which yields a safety factor of 217.39. The aluminum
honeycomb structure is unlikely to fail, but it is designed to be able to be replaced with a new
cartridge upon mission completion.
Table 10: Aluminum honeycomb compressive strengths.
Final Properties
In Appendix B, Figure 49 show the lander comfortably fits inside the 8.4 m diameter SLS payload
capsule. Figures 50 show the center of mass of the SLT with empty tanks; note that the center of
mass is mostly central with some skewing towards the side Hazard Avoidance LiDAR and the
NDL. Figures 51 show the center of mass when the tanks are full, where the center of mass is
mostly central with some skewing towards the heavier oxidizer tank. Figures 52 show that the SLT
easily clears the 30-cm-tall obstructive object even when the SLT is compressed, with 21 in. of
clearance from the object to the bottom of the SLT, and 8 in. clearance from the ground to the
engines. Additional mass properties and stress analyses are found in Appendices C and D.

55
Ground Qualification and Testing
Overview
In order to verify the suitability of the SLT to survive and complete the assigned mission, a set of
rigorous testing procedures have been set up. The SLT shall be subjected to end-to-end testing,
for each subassembly and the final assembly, covering structural loads, shock loads, vibroacoustic
environments, thermal environments, and liftoff and ascent venting pressure profiles. All testing
requirements are found in the General Environmental Verification Standard (GEVS), the ESD
30000 Space Launch System (SLS) Mission Planner’s Guide, and NASA standards. In order to
reduce cost and travel time all testing of the SLT will be conducted at the appropriate facilities at
NASA’s Johnson Space Center.
Structural Load Testing
For structural load testing, this testing was chosen to be completed first as it allows flight
manager(s) and engineers to understand where any weaknesses due to stress or coupled loads
during liftoff and transonic flight could be found in either the subassembly or final assembly as
structural failures will lead to overall catastrophic failures. Each subsystem and the final assembly
will be tested to 1.25 times the limit loads and a stress analysis, coupled load analysis, a finite
element model, and a dynamic clearance analysis shall be developed to ensure positive margins
and appropriate Factors of Safety. The SLT shall be qualified for structural load factors as seen in
Table 11.
Table 11: Structural load factors the SLT shall be qualified for.

56
The static structural load tests will be performed at the Structures Test Laboratory while the
dynamic structural load tests, will be performed at the General Vibration Laboratory.
Shock Load Testing
For shock load testing, this was deemed as necessary to be completed next due to the separation
events that will occur during SLS flight very close to the payload. Ensuring these tests are passed
will allow the flight manager(s) to have confidence that the SLT will survive the high shock
levels during SLS flight and separation events. Each subsystem and the final assembly will be
tested to 1.4 times the maximum expected shock for each of the three axes. The SLT will be
tested in the appropriate electrical and mechanical operational modes. Before and after testing,
the SLT shall be examined and functionally tested. The SLT shall be qualified for shock load
factors that could affect payloads during separation events. Representative payload separation
system induced shock environments are shown in Table 12. These tests will be performed at the
Spacecraft Vibration Laboratory.
Table 12: Representative payload separation system induced shock environments.
Acoustics Environment Testing
For Acoustics Environment testing, this was deemed as necessary to be completed next due to
the internal acoustic environment the SLT will be subject to during liftoff and transonic flight
causing potential damage to the SLT. The minimum overall acoustic test should be at least 138
dB and during testing, the SLT shall be in an operational configuration. The SLT shall be
qualified the payload's internal acoustic environment is seen in Figure 46. Acoustic load testing
will be performed at the Spacecraft Acoustic Laboratory.

57
Figure 46: The payload’s internal acoustic environment that the SLT shall be qualified for.
Vibration Environment Testing
For vibration environment testing, this follows the acoustic testing as the acoustic testing covers
the random vibration environment for any frequency greater than 100 Hz. For frequencies less
than 100 Hz, the random vibration environment can be seen in Table 13. Additionally, the SLT
shall be subjected to a sine sweep vibration design qualification test. Vibration environment tests
will be performed at the Spacecraft Vibration Laboratory.
Table 13: The random vibration environment for any frequency less than 100 Hz.
Thermal Environment Testing
For thermal environment testing, these tests were decided to be performed near the end as it is the
harshest tests the SLT will be subjected to. Additionally, thermal tests will also be conducted at
the individual component level as well as the subassembly and final assembly level. Four (4)

58
thermal-vacuum temperature cycles shall be performed at the subassembly and final assembly
level while eight (8) cycles will be performed at the component level. When possible, each test
subject will be operating, and its performance shall be monitored. The SLT shall be exposed for a
minimum of twenty-four (24) hours at each extreme of each temperature cycle. The SLT will be
tested in a temperature range of -185 °C to 200 °C (to ensure a factor of safety of at least 1.5). Low
temperature thermal environment tests will be conducted in Chamber B and high temperature tests
will be conducted in Chamber P.
Liftoff and Ascent Venting Testing
Finally, this testing was decided to be performed last as it focuses mostly on how the SLT will
perform during liftoff and ascent. The SLT will be subjected to the pressure envelope on the slide
and these tests will be conducted at the White Sands Test Facility. During ascent, SLS Block 1B
crew configuration payloads shall be qualified to withstand the pressure envelope shown in Figure
47.
Figure 47: The pressure envelope for SLS Block 1B crew configuration payloads during ascent.
A pressure analysis will be conducted to ensure a positive margin at loads equal to twice the
maximum pressure differential during launch. These tests will be conducted at the White Sands
Test Facility.

59
Project Management
Risk Assessment
A risk assessment was performed for the entire SLT, the result of which can be seen in Table 14.
The definition of the different levels of criticality can be seen in Table 15. For the sake of
discussion, particular attention was paid to criticality 1 and 1R items. These particular risk items
were outlined in Table 15. With the known risks for the SLT, a significant amount of redundancy
included into its design. Not only with having backup components wherever possible, but also
ensuring that many of the parts selected had built in redundancy.
Table 14: Risk Assessment Matrix.
Due to the redundancy that was planned into every sub-system, and the reliability of the parts and
materials chosen, it was determined that no risk goes above a medium level risk for the Gator
Hopper SLT. Even catastrophic risks have a very low probability. This means that although there
are a total of 8 criticality 1 items and 5 criticality 1R items, the likelihood of them occurring is
very low. For example, although all three On-Board-Computers failing would be catastrophic, and

60
are labeled as a criticality 1R item, this is very unlikely to happen. That is because the computers
have software designed parity, as well as three total computers for redundant hardware backups.
Similarly, extreme caution, care, and where possible, physical redundancy was added to any other
items that fall under these high levels of criticality. These items are either very unlikely to fail,
have built in redundancy, have physical backup systems, or some combination of these to ensure
the Gator Hopper SLT is overall very reliable.
Criticality Level Definition
Criticality 1 Loss of life or vehicle if the component fails.
Criticality 2 Loss of mission if the component fails.
Criticality 3 All others.
Criticality 1R Redundant components, the failure of both could cause loss of life or vehicle.
Criticality 2R Redundant components, the failure of both could cause loss of mission.
Table 15: Definitions for Criticality Levels.
Subgroup Criticality 1 Criticality 1R
Electrical 1) Power distribution board fails
1) Both batteries fail
2) All three On-Board-Computers fail
Propulsion
1) Backflow up propellants from
combustion chamber
2) Pressure regulator failure
1) Both main engines
fail/malfunction
2)Complete failure to propellant lines
from regolith
Guidance and
Control
1) Hazard avoidance LiDAR fails 1) Both ASTROgyro systems fail
Structures
1) Footpad detachment during flight
2) Material failure during flight
3) Failure at welded points during
flight
None
Environmental
1) Failure of ignition
overpressure/sound suppression
(IOP/SS) system upon launch
None
Table 16: Risk assessment, SLT Criticality 1 and 1R items.

61
Requirements Matrix
As can be seen in Table # in Appendix E, all of the requirements for the SLT’s development have
been completed except for a few excepts. ASD-SLT-32 has been left in progress because it has
been determined that the customer will be responsible for acquiring any specialized equipment.
Additionally, ADS-SLT-33 has also been left incomplete since it cannot technically be fulfilled
until the final delivery of the SLT.
OISR and Schedule
Upon completion of the course, the project will enter the execution stage where the initial steps of
contacting manufacturers begin. From this point until 4/6/28 the Gator Hopper SLT will be under
construction and development. The open items for the project can be seen in Table 17, and the
schedule for their completion can be seen in Figure 48.
The project will conduct manufacturing and research and development over the first 18 months
which is expected to be completed on 4/27/2022. Upon completion, the next 6 months will consist
of preparation for subsystem assembly testing. The subsystem assembly testing is expected to take
21 months and last from 10/13/2022 to 5/22/2024. Following the subassembly testing, the
integrated system testing will begin and last approximately 48 months with the estimated
completion date being 1/26/2028. After testing has been completed, the SLT will then go through
delivery preparation and delivery will be completed on 4/6/2028. The vehicle will then be in
management reserve for a little over a year and then will enter launch preparation a month before
the estimated launch date of 11/1/2029.

62
Figure 48: Planned schedule through launch.
Table 17: OISR for the Gator Hopper SLT.

63
Budget
Overview
The general breakdown for the Gator Hopper SLT project budget can be seen in Table 18, with a
more detailed breakdown in Appendix F. With all expenses considered it was determined that a
total of $1,356,491,356 will be spent over the duration of the SLT’s development and launch.
Compared to the overall budget of $3,900,000,000, the Gator Hopper SLT project will finish
$2,653,525,047 under budget. When only the total cost of the SLT was considered (development
and testing excluded), it was found to cost $18,410,906 for a single unit. When this was compared
with the budget for one unit being $900,000,000 it was determined that the Gator Hopper SLT
would be completed for $885,589,094 under budget. This means that the complete SLT project
would come in 65% under budget, and the unit cost of one Gator Hopper SLT would come in 98%
under budget.
Table 18: General expenses breakdown for Gator Hopper SLT design and development.
Staff
50 individuals have been identified as the appropriate number for a successful mission as seen in
Table 19. These individuals were allocated based on similar projects in the industry. The
industry standard was taken for their salaries.

64
Table 19: Budget breakdown for staff.
Conclusion
The vehicle that has been designed and described above provides NASA with a proper support
vehicle in order to ensure human and mission safety for NASA’s lunar bases. It can deliver up to
200 kg of emergency equipment up to 300 km away from LBC near Shackleton Crater, it can
survive for more than 72 hours in complete darkness while waiting for its return trip, and it can
avoid hazards during flight and landings. The proposed SLT is under the mass restrictions given
by NASA as seen in Appendix G, and the entire project only uses 35% of its entire budget. If
NASA adopted this design and implemented it, it would move NASA one step closer towards
100% mission assurance and safety.

65
Works Cited
[1] https://rps.nasa.gov/power-an vd-thermal-systems/thermal-systems/light-weight-radioisotope-
heater-unit/
[2] https://www.nasa.gov/centers/johnson/pdf/639595main_EA_Test_Facilities_Guide.pdf
[3] https://weather.com/weather/monthly/l/5ef49d14d6fa75c9dff0a704b55e34c0508ee35e5bab186fb
5f800fa680e6693
[4] https://www.lpi.usra.edu/meetings/lpsc2013/eposter/2617.pdf
[5] ESD 30000 SLS Mission Planner’s guide
[6] http://www.nanoflexpower.com/gallium#:~:text=GaAs%20is%20the%20highest%20performance
,in%20a%20given%20surface%20area
[7] https://ntts-prod.s3.amazonaws.com/t2p/prod/t2media/tops/pdf/LEW-TOPS-50.pdf
[8] https://now.northropgrumman.com/aerospace-technology-coming-soon-to-a-solar-installation-
near-you/
[9] https://www.researchgate.net/publication/296692705_Silicon_space_solar_cells_progression_and
_radiation-resistance_analysis
[10]Galium Arsenide Solar cells on unmanned aerial vehicles
[11]https://www.pv-tech.org/editors-blog/what-the-us-navys-solar-drones-tell-us-about-thin-film-
solars-potential
[12]Increasing Markets and Decreasing Package Weight for High Specific Power Photovoltaics
[13]https://www.solarpowerworldonline.com/2016/03/kind-solar-panels-nasa-use/
[14]Astronaut CAD https://grabcad.com/library/nasa-z2-emu-mmu-by-tommy-1
[15]https://www.terma.com/media/177707/power_conditioning_and_distribution_unit.pdf
[16]NanoAvionics GaAs Solar Panels Data Sheet
[17]https://www.gore.com/system/files/2019-10/GORE%20Space%20Cables%20-
%20Catalog%20%28Traditional%20Space%29_10-28-
2019%20%28A4%20Electronic%29_0.pdf
[18]https://ntrs.nasa.gov/citations/20140009928
[19]https://asc-sensors.de/datenblatt/honeywell/beschleunigungssensor/q-flex/qa-2000.pdf
[20]https://asc-sensors.de/datenblatt/honeywell/beschleunigungssensor/q-flex/qa-3000.pdf
[21]http://www.matweb.com/index.aspx
[22]https://medium.com/teamindus/structural-evolution-of-the-teamindus-spacecraft-that-will-land-
on-the-moon-b5aa6bc73ccc
[23]https://www.nasa.gov/centers/wstf/pdf/210566main_veh_ascent_descent.pdf
[24]GEVS, NASA-STD-5001B, NASA-STD-7001B, NASA-STD-7002B
[25]https://www.spacevector.com/CMS/images/39381_Li-Po_Battery_Brochure.pdf
[26]http://www.gsyuasa-lp.com/SpecSheets/MA190.pdf
[27]https://phys.org/news/2018-01-advanced-multi-junction-solar-cells-high.html
[28]https://www.cubesatshop.com/product/iceps-spacecraft-system-core/
[29]https://www.astrobotic.com/peregrine
[30]https://www.rocket.com/sites/default/files/documents/In-
Space%20Data%20Sheets%204.8.20.pdf
[31]https://www.facebook.com/63375477753/photos/xaero-b-is-powered-by-a-more-powerful-
engine-than-any-previous-masten-vehicle-th/10151635344362754/
[32]https://www.space-propulsion.com/spacecraft-propulsion/bipropellant-tanks/index.html#282
[33]https://www.mt-aerospace.de/downloadcenter.html?file=files/mta/tankkatalog/MT-
Tankkatalog_01b_4-3_03.pdf

66
Appendices
Appendix A: Vehicle Sub-system Risk Assessment
Levels of risk based on severity and probability of occurrence
Risk Assessment for Mechanism and Structures

67
Risk Assessment for Environment and Testing
Risk Assessment for Electrical

68
Risk Assessment for Guidance and Control
Risk Assessment for propulsion

69
Appendix B: SLT verification figures
Figure 49: Lander size verification inside SLS capsule.
Figure 50: SLT center of mass with empty tanks.
Figure 51: SLT center of mass with full tanks.

70
Figure 52: SLT obstacle clearance.
Appendix C: SLT Mass properties
• Mass = 2384.51 kg (Including payload, during flight)
• Center of Mass: (m)
o X = 0.36
o Y = -0.59
o Z = -1.01
• Principal axes of inertia and principal moments of inertia: (kg-m2
)
• Taken at the center of mass.
o Ix = (0.71, -0.02, 0.70) Px = 1160.91
o Iy = (0.70, -0.03, -0.71) Py = 2099.21
o Iz = (0.04, 1.00, -0.01) Pz = 2243.40
• Moments of inertia: (kg-m2
)
• Taken at the center of mass and aligned with the output coordinate system.
o Lxx = 1624.64 Lxy = -21.42 Lxz = 468.88
o Lyx = -21.42 Lyy = 2242.65 Lyz = -14.83
o Lzx = 468.88 Lzy = -14.83 Lzz = 1636.23

71
Appendix D: Various Stress Analyses
Figure 53: Stress analyses of side of base structure and titanium connector.
Figure 54: Stress analysis of where the propellant tank mounts.

72
Appendix E: Requirements Matrix
Number Name A O D T R Verification Artifact Status
ASD-
SLT-01 Launch Vehicle Capability X X X SolidWorks Verified
ASD-
SLT-02 Structural Load Environment X
MATLAB /
Calculations Verified
ASD-
SLT-03 Shock Loads X
MATLAB /
ASD-
SLT-04 Acoustics Environment X
MATLAB /
ASD-
SLT-05 Vibration Environment X
MATLAB /
ASD-
SLT-06 Thermal Environment X
MATLAB /
ASD-
SLT-07 Liftoff and Ascent Venting X
MATLAB /
ASD-
SLT-08 Electrical Bonding / Grounding X
MATLAB /
ASD-
SLT-09 SLT Vehicle Mass X X SolidWorks Verified
ASD-
SLT-10 Spacecraft Coordinate System X X SolidWorks Verified
ASD-
SLT-11 SLT Center of Gravity X SolidWorks Verified
ASD-
SLT-12 Debris X Design Review Verified
ASD-
SLT-13
SLT Moments and Products of
Inertia X SolidWorks Verified
ASD-
SLT-14 SLT Payload Mass Capability X SolidWorks Verified
ASD-
SLT-15 SLT Payload Volume Capability X
Design Review /
Solidworks Verified
ASD-
SLT-16 SLT Payload Interface X X
Design Review /
ASD-
SLT-17 SLT Payload Power Capability X
Design Review /
ASD-
SLT-18 Payload Communication Capability X X Design Review Verified
ASD-
SLT-19 STL Operational Range Capability X STK Verified
ASD-
SLT-20 Propulsion System Capability X X Calculations Verified
ASD-
SLT-21 Nominal Operating Environment X X Calculations Verified

73
ASD-
SLT-22
Off-Nominal Operating
Environment X X Calculations Verified
ASD-
SLT-23 Operational Cycle Capability X X Calculations Verified
ASD-
SLT-24 Navigation System Capability X X Design Review Verified
ASD-
SLT-25 Communication System Capability X X X Design Review Verified
ASD-
SLT-26 Landing System Slope Capability X X X X SolidWorks Verified
ASD-
SLT-27
Landing System Obstacle
Clearance Capability X X X SolidWorks Verified
ASD-
SLT-28
Landing System Obstacle
Adaptation Capability X X X SolidWorks Verified
ASD-
SLT-29 End to End Testing X In Presentation Verified
ASD-
SLT-30 SLT Logo X SolidWorks Verified
ASD-
SLT-31 Mission Operations X X Design Review Verified
ASD-
SLT-32 Specialized Support Equipment X
Customer Buys From
Masten
In
Progres
s
ASD-
SLT-33 Delivery X
Verification Upon
Delivery Waiting
ASD-
SLT-34 Budget X X
Design Review /
Appendix F: Budget Breakdown
Category/Item Name Quantity Cost Per Unit Total Cost Supplier
Mechanisms and Structures
5x5in 0.25" thick Alum Extrusion 6ft long 8 $233.18 $2,798.16 McMaster
4x4in 0.25" thick Alum Extrusion 6ft long 6 $220.32 $1,982.88 McMaster
5x5in Solid Alum Extrusion 6ft long 1 $628.90 $943.35 McMaster
4.5" OD .25" thick Cylindrical Alum Extrusion 6ft
long 7 $247.59 $2,599.70 McMaster
24"x24" 1" thick aluminum Honeycomb 2 $75.96 $227.88 McMaster
18" x 18" x 0.75" Alum Sheet 1 $432.35 $518.82 McMaster
24" Long x 2" Diameter Alum Rod 1 $98.40 $118.08 McMaster
6" Long x 3" Diameter Alum Rod 1 $69.84 $83.81 McMaster
0.5" x 1" x 24" Alum Bar 1 $47.25 $56.70 McMaster
5" Diameter x 6" Long Alum Rod 1 $186.74 $224.09 McMaster
0.005" Thick Teflon PTFE 1 $15.20 $18.24 McMaster

74
Propulsion
R40B 2 $5,000,000 $10,000,000 Aerojet Rocketdyne
R-1E 12 $300,000 $3,600,000 Aerojet Rocketdyne
769 Litre Biprop Tank 2 $100,000 $200,000 Ariane Group
Helium/Nitrogen High Pressure Tank 1 $30,000 $30,000 MT Aerospace
Monomethylhydrazine 769 L $147/LB $219,310 LookChem
Nitrogen Tetroxide 769 L $120/LB $292,956 LookChem
Pressure Regulator 1 $5,000 $5,000 Marotta
Check Valves 30 $400 $12,000 Marotta
Pressure Relief Valve 2 $750 $1,500 Valcor
Fill/Drain/Vent Valve 4 $1,000 $4,000 Ariane Group
Dual Pressure & Temperature Transducer 4 $3,000 $12,000 GP:50
Proportional Control Throttling Valve 2 $10,000 $20,000
1/2" Stainless Steel Tube 3.91 m $130 $130 McMaster
1" Aluminum Tube 20.8 m $136 $136 McMaster
1/4" Aluminum Tube 9.9 m $72 $72 McMaster
Hardware & Electronics
2" Stroke Heavy Duty Linear Actuator 4 $305 $1,220 Progressive Automations
25 Ah Li-ion Polymer Battery 39381 Series 2 $50,000 $100,000 Space Vector
3U-FMmT Solar Panels 156 $7,850 $1,224,600 Nano Avionics
Power Conditioning and Distribution Unit 1 $1,000,000 $1,000,000 TERMA Space
Comms, Tracking, Trajectory, and Guidance
ASTROgyro Attitude Control System 2 $150,000 $300,000 Jena-Optronik
QA-3000-030 Accelerometer 6 $1,000 $6,000 Honeywell
High-Altitude Laser Altimeter 1 $300,000 $300,000
Navigation Doppler Lidar 1 $300,000 $300,000
X-Band Transceiver 2 $53,152 $106,304 IQ Wireless
On-Board-Computer (OBC) 3 $34,309 $102,928 EXA
Hazard Detection System 1 $300,000 $300,000
SLT TOTAL $18,410,906
Manufacturing and Labor
Salary of Mission Operation Engineer 90 $88,000 $7,920,000 NASA
Salary of Quality Assurance Engineer 90 $94,000 $8,460,000 NASA
Salary of Electrical Engineer 90 $81,000 $7,290,000 NASA
Salary of Manufacturing Engineer 90 $54,080 $4,867,200 NASA
Salary of Design Engineering 90 $100,000 $9,000,000 NASA
Ground Operations and Testing

75
Vibration Test 2 $100,000,000
$200,000,00
0 NASA Johnson Space Center
Acoustic Test 2 $100,000,000
$200,000,00
Thermal Vaccuum Testing 3 $100,000,000
$300,000,00
Structural Load Testing 2 $100,000,000
$200,000,00
Pressure Profile Testing 2 $100,000,000
$200,000,00
Shock Testing 2 $100,000,000
$200,000,00
Other Mission Expenses
Ground Processing 1 $300,000 $300,000 United Launch Alliance
Environmental
Mylar/kapton MLI 38 $20.00 $30,400 RUAG
Radioisotope heater units 25 $4,333 $216,650
Kapton heating tape 288 $56 $16,128 Omega Engineering
Moog Shockwave Isolator 4 $500 $2,000 moog
Melamine acoustic foam
~250
m^2
$247.6 per square
m $62,000 Polytechinc
Radiator 7.57kg $22000 per kg $167,000 Paragonsdc
Louver 6.5m^2 $500 per m^2 $3,250 Paragonsdc
hook & pile fasteners & grommets $1,000
heat pipes 15m^2 $500 per m^2 $8,000
Advanced cooling
technologies
TOTAL $1,356,491,356
Appendix G: SLT Mass breakdown (without payload)
Sub-Team
Mass (kg)
Launch from
Earth Mission
Propulsion (valves, feed lines,
engines...) 1240.6 1480.2
Electronics (solar panels, wires...) 42.871
Structures (base structure, gimbal
system...) 453.127
Guidance & Control (LiDAR, star
trackers...) 72.19
Comms & Tracking (OBC, transciever) 1.592
Environmental (insulation, radiator...) 53.02
Total Mass 1863.4 2103

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