Centurion Final Design Document

Conceptual Design for an Orbit Transfer Vehicle
FDR Report
Prepared for:
The American Institute of Aeronautics and Astronautics
Undergraduate Team 4 Space Design – AE443S
April 14th, 2015
Prepared by:
Mruthyum (Jay) Mulakala
Samip Shah
Bentic Sebastian
Benjamin Wilson
Derek Awtry
Kevin Lohan
Yu Guan

Engineering Team
Jay Mulakala
Lead Systems Engineer
Pgs. 1-8, 69-77
X .
Derek Awtry
Orbital Engineer
Pgs. 8-20
X .
Benjamin Wilson
Propulsion Systems Engineer
Pgs. 20-29
X .
Yu Guan
Structural Engineer
Pgs. 30-43
X .
Bentic Sebastian
Power and Thermal Systems Engineer
Pgs. 38-43, 53-61
X .
Samip Shah
ADCS Engineer
Pgs. 38-53
X .
Kevin Lohan
Launching and Docking Engineer
Pgs. 61-69
X .

Table of Contents
List of Figures .............................................................................................................................................i
List of Tables..............................................................................................................................................ii
List of Acronyms.......................................................................................................................................iii
1. Executive Summary ....................................................................................................................................1
1.1 Mission Timeline..................................................................................................................................7
2. Orbital Systems ...........................................................................................................................................8
2.1. Design Approach.................................................................................................................................8
2.2. Concept Development .........................................................................................................................9
Trajectory to L1 ...........................................................................................................................9
Trajectory to L2 .........................................................................................................................12
Station-keeping Analysis ...........................................................................................................16
Orbital Maintenance in LEO......................................................................................................17
Aerobraking maneuvers .............................................................................................................18
End of Life Summary.................................................................................................................19
2.3. Critical Design Issues ........................................................................................................................20
Computation of Halo Orbits and Insertion Velocities................................................................20
Computation of Invariant Manifolds..........................................................................................20
3. Propulsion Systems ...................................................................................................................................20
3.1. Design Approach...............................................................................................................................20
3.2. Concept Development .......................................................................................................................21
Main Propulsion System ............................................................................................................21
Attitude Control Propulsion System ..........................................................................................28
4. Structural Definition..................................................................................................................................30
4.1 Design Approach................................................................................................................................30
4.2 Concept Development ........................................................................................................................30
Material Selection ......................................................................................................................30
Mass Estimation.........................................................................................................................33
Vehicle Internal Volume............................................................................................................34
Vehicle Structural Design ..........................................................................................................34
Structural Testing.......................................................................................................................39
4.3 Critical Design Issues .........................................................................................................................40
5. Communication and Systems ....................................................................................................................40
5.1 Frequency Band Selection ..................................................................................................................40
5.2 Radiometric Tracking.........................................................................................................................41
5.3 Antenna Selection...............................................................................................................................44
6. Attitude Determination and Control Systems............................................................................................45
6.1 Design Approach................................................................................................................................45
Sensor Selection.........................................................................................................................46
Actuator Selection......................................................................................................................50
Onboard Processing and Control Methods.................................................................................53
7. Spacecraft Power Management Systems...................................................................................................55

7.1. Design Approach...............................................................................................................................55
Power generation and distribution .............................................................................................55
Power Storage ............................................................................................................................57
Radiation shielding ....................................................................................................................59
Emergency mode .......................................................................................................................59
8. Spacecraft Thermal Systems .....................................................................................................................60
8.1. Design Approach...............................................................................................................................60
Thermal Control of Nuclear Reactors........................................................................................60
Thermal Control of Fuel Tanks..................................................................................................62
Thermal control of Systems Module..........................................................................................63
Thermal Control of Solar Panels................................................................................................65
9. Launching and Docking ............................................................................................................................65
9.1 Design Approach................................................................................................................................65
Launch Vehicle ..........................................................................................................................66
Discussion..................................................................................................................................68
Docking System.........................................................................................................................69
Refueling Procedure...................................................................................................................72
9.3 Critical Design Issues .........................................................................................................................73
10. Risk and Cost Analysis............................................................................................................................73
10.1 Risk Analysis and Mitigation..........................................................................................................73
10.2 Cost Estimation...............................................................................................................................76
11. Conclusion...............................................................................................................................................81
12. References ...............................................................................................................................................82

i
List of Figures
Figure 1-1. Earth-Moon Lagrange Points [1] ..................................................................................................1
Figure 1-2. Illustration of Centurion OTV in Earth Orbit ...............................................................................2
Figure 1-3. Detailed Illustration of Centurion OTV........................................................................................4
Figure 1-4. Illustration of Centurion OTV ......................................................................................................5
Figure 1-5. Illustration of Centurion OTV Trajectory.....................................................................................6
Figure 1-6. Long Term Mission Timeline .......................................................................................................7
Figure 2-1. Side view of the trajectory to L1 ................................................................................................13
Figure 2-2. View of L1 from the perspective of the Moon............................................................................13
Figure 2-3. Zoomed view of the halo orbit at L1 ..........................................................................................13
Figure 2-4. L2 trajectory as seen from the Earth...........................................................................................15
Figure 2-5. L2 halo orbit as seen from the Moon..........................................................................................15
Figure 2-6. L2 trajectory in the Earth-Moon rotating axis ...........................................................................16
Figure 3-1. Basic LANTR schematic [23]....................................................................................................22
Figure 3-2. Aerojet R-1E thrusters [30].........................................................................................................29
Figure 4-1. PAMG-XR1 5056 Aluminum honeycomb .................................................................................32
Figure 4-2. CYCOM 5320-1 toughened epoxy resin prepreg system [37]....................................................32
Figure 4-3. Systems Module with outer casing removed ..............................................................................35
Figure 4-4. Hinged Radiator for Systems Module [38].................................................................................36
Figure 4-5. Cryogenic Propellant Tank .........................................................................................................37
Figure 4-6. Propulsion System with outer casing removed...........................................................................37
Figure 4-7. Deployable radiator [39].............................................................................................................38
Figure 4-8. Stress Analysis of cryogenic propellant tank..............................................................................40
Figure 5-1. Atmospheric attenuation as a function of frequency [43]...........................................................41
Figure 5-2. Number of missions using NEN vs. DSN [47] ...........................................................................42
Figure 5-3. NEN performance compared to DSN using the S-Band [47] .....................................................43
Figure 5-4. EIRP of NEN compared to DSN in the S-Band [47]..................................................................44
Figure 5-5. High Gain Antenna with parabolic reflector...............................................................................45
Figure 6-1. Layout of ADCS Components....................................................................................................47
Figure 6-2. Surrey Rigel-L ............................................................................................................................48
Figure 6-3. Adcole Course Sun Sensor Pyramid...........................................................................................49
Figure 6-4. Configuration of attitude control thrusters..................................................................................51
Figure 6-5. Triple Mode Redundancy Configuration....................................................................................55
Figure 7-1. Power schematic of He-Xe gas for electricity production [71]...................................................56
Figure 7-2. Power distribution schematic......................................................................................................57
Figure 7-3. BFO (blood-forming-organ) dose [74] .......................................................................................59
Figure 8-1. Close-up of the radiators.............................................................................................................62
Figure 8-2 Braytpn Cycle for ESCORT System [24]....................................................................................61
Figure 8-3. Detailed wireframe of the OTV..................................................................................................63
Figure 8-4 Solar panels at system module.....................................................................................................65
Figure 0-1. Conceptual Design for NASA Docking System [81] .................................................................70
Figure 0-2. Modified Dextre Robot [86] .......................................................................................................72
Figure 9-1. Technology Risk Analysis..........................................................................................................74
Figure 9-2. Operational Risk Analysis ..........................................................................................................75
Figure 9-3. Number of Falcon 9 launches required to transport fuel for 10 missions ...................................79
Figure 9-4. Total project costs using Centurion versus conventional technologies.......................................80

ii
List of Tables
Table 1.1. AIAA Mission Requirements.........................................................................................................2
Table 2.1. Hohmann Transfer Burns and Insertions, adapted from [10] .......................................................10
Table 2.2. Delta V's for different orbits, table adapted from [10] .................................................................10
Table 2.3. Total mission ΔV for varying z-amplitudes..................................................................................11
Table 2.4. Transfer times to L1 and halo orbit durations for varying z-amplitudes ......................................11
Table 2.5. Total mission ΔV for varying z-amplitudes..................................................................................14
Table 2.6. Transfer times to L2 and halo orbit durations for varying z-amplitudes ......................................15
Table 2.7. Station-keeping values for L1 halo orbits.....................................................................................17
Table 2.8. Station-keeping values for L2 halo orbits.....................................................................................17
Table 2.9. Delta V's of possible altitudes at which to aerobrake with a 10 meter heat shield. ......................19
Table 3.1. Potential Main Propulsion System Technologies .........................................................................21
Table 3.2. Fuel Consumption to and from L2 [Isp = 911s].............................................................................27
Table 3.3. Potential ACS propulsion technologies........................................................................................28
Table 3.4. Potential ACS propellant tanks ....................................................................................................29
Table 4.1. Comparison of common material for space vehicles [16] ............................................................31
Table 5.1. NEN Frequency Band Characteristics [2] ....................................................................................41
Table 5.2. Near Earth Network Tracking Characteristics [20] ......................................................................42
Table 5.3. Types of antenna for space communication [11]..........................................................................44
Table 6.1. Characteristics of common star trackers [22] [23] [24]................................................................48
Table 6.2. Characteristics of common IMUs [26] [27] [28]..........................................................................48
Table 6.3. Characteristics of common sun sensors [5] ..................................................................................49
Table 6.4. Capture Tolerances for Docking and Berthing [29] .....................................................................49
Table 6.5. Demonstrated Accuracy of AOS Proximity Sensors [30] ............................................................50
Table 6.6. Characteristics of common thrusters [31] [32] [33] [34]..............................................................51
Table 6.7. Characteristics of commonly used control moment gyroscopes [37] [38] [39]............................52
Table 6.8. Estimated Source Lines of Code [41]...........................................................................................53
Table 6.9. Characteristics of radiation hardened flight processors [42] [43] [44] .........................................54
Table 7.1. Batteries and their characteristics [36] .........................................................................................58
Table 8.1. Diagram of upper casing, lower casing, and fuel tank..................................................................63
Table 8.2. Heat Dissipation among components ...........................................................................................64
Table 9.1. Comparison of Potential Launch Vehicles for Centurion.............................................................67
Table 9.2. Launch Vehicle Selection Factors and Weighting........................................................................67
Table 9.3. Trade Study of the Viable Launch Vehicles.................................................................................68
Table 9.4. IDSS Docking Compatability [57] [6]..........................................................................................71
Table 10.1. Risk Analysis Criteria ................................................................................................................73
Table 10.2. Technology Risk Analysis..........................................................................................................74
Table 10.3. Operational Risk Analysis..........................................................................................................75
Table 10.4. Development and Mission Costs................................................................................................78

iii
List of Acronyms
ADCS – Attitude Determination and Control Systems
AIAA – American Institute for Astronautics and Aeronautics
AMSL – Above Mean Sea Level
APAS – Androgynous Peripheral Attachment System
ATCS – Active Thermal Control System
BLS – Boeing Launch Services
BNTR – Bimodal Nuclear Thermal rocket
CALT – China Academy of Launch Vehicle Technology
CECE – Common Extensible Cryogenic Engine
CIS – Commonwealth of Independent States
CMG – Control Moment Gyroscope
CR3BP – Circular restricted three body problem
EML – Earth-Moon Lagrangian Point
ESA – European Space Agency
FDR – Final Design Report
Isp – Specific Impulse
L1 – Lagrnage point 1
L2 – Lagrange point 2
LANTR – Liquid Oxygen Augmented Nuclear thermal rocket
LEO – Low Earth Orbit
LH2 – Liquid Hydrogen
LOX – Liquid Oxygen
NASA – National Aeronautics and Space Administration
NDS – NASA Docking System
NERVA – Nuclear Engine for Rocket Vehicle Applications
NTR – Nuclear Thermal Rocket
OTV – Orbit Transfer Vehicle
PDR – Preliminary Design Report
PTCS – Passive Thermal Control System
QFD – Quality Function Deployment
RCS – Reaction Control Systems
RRM – Robotic Refueling Mission
SNRE – Small Nuclear Rocket Engine
STPO – Space Transportation Project Office
TFU – Theoretical First Unit
TRL – Technology Readiness Level
XE – Experimental Engine

1
1. Executive Summary
Hyperion Ventures’ aims to provide a transportation vehicle that satisfies the requirements set forth by the
American Institute for Astronautics and Aeronautics (AIAA). The task was to develop an Orbital Transfer Vehicle
(OTV) capable of transporting payloads between Low Earth Orbit (LEO) and two Lagrange points, either EML1 or
EML2. There are currently 5 Lagrange points around Earth, as shown in Figure 1-1. Two of these five positions offer
an area where the combined gravitational pull of the Earth and Moon offer a stable orbit configuration, while the other
3, L1, L2, and L3 are unstable but offer the ideal location for a potential space station due to their positioning and
accessibility [1].
Several missions have been planned over the
past few decades to utilize these points, from a Deep
Space Climate Observatory to the James Webb Space
Telescope to a design proposed by Boeing that would
serve as a refueling depot and servicing station. The
platform would serve as a base for deep space
exploration, robotic relay stations for moon rovers,
telescope servicing, and even mars base missions. In
order to meet these growing demands, Hyperion
Ventures is tasked with developing an Orbit Transfer
Vehicle (OTV) capable of transporting unmanned and
manned payloads between Low Earth Orbit (LEO) and
Earth-Moon Lagrangian points L1 (EML1) or L2 (EML2). The benefits for these points have been researched for
decades and motivation for development at these Lagrange points have grown. The American Institute of Aeronautics
and Astronautics’ (AIAA) Request for Proposal (RFP) clearly dictates the constraints and requirements for an OTV
mission to these Lagrangian points for potential use in future missions. The specific design constraints are listed in
Table 1.1. To satisfy these constrains, Hyperion Ventures has designed a vehicle to satisfy the AIAA criteria, called
Centurion, Figure 1-2.
Figure 1-1. Earth-Moon Lagrange Points [1]

2
Table 1.1. AIAA Mission Requirements
Number Condition: Reference:
1 The OTV will be stationed in 400 km AMSL circular LEO with 28° inclination. Section 2
2 The OTV payload capability shall be 50,000 lbs from LEO to EML1 and 15,000 lbs
from EML1 to LEO.
Section 3
3 The OTV must be capable to remain at EML1 or EML2 for at least 30 days. Section 2
4 Each transfer should not exceed 6 days. Section 2
5 The life of the OTV shall be 5 years and the OTV shall be capable of at least 10 missions
to EML1 or EML2.
Section 4
Centurion is a modular design vehicle equipped with some of the latest technologies, including a nuclear
thermal propulsion system. The structure weighs approximately 89,000 kg of which about 49,000 kg is the fuel
onboard the vehicle. It is capable of docking with a variety of payloads and capsules and has the ability to transport
that cargo to Lagrange points L1 or L2. The total cost to design, fabricate, and launch Centurion would be about $2.5
billion, with subsequent missions costing about $95 million each.
Centurion redefines modularity and simplicity. The vehicle is composed of different modules designed
specifically for this mission, the optimal attitude determination and control systems for the vehicle, power and thermal
systems that reduce the weight of Centurion while not compromising safety, an orbital plan that can almost cut travel
time in half compared to conventional technologies, a revolutionary new propulsion system never before used in
action, and a launching and docking mechanism that allows the vehicle to dock with ease.
Figure 1-2. Illustration of Centurion OTV in Earth Orbit

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The primary structure of Centurion is composed of aluminum and titanium. Concerns regarding failure under
stress and thermal conditions have been taken into consideration to reduce stress concentrations while maintaining a
strong, stable structure. The base of the vehicle consists of three nuclear thermal engines supplied by a large liquid
hydrogen fuel tank containing over 49,000 kg of fuel. Liquid hydrogen was chosen as the primary fuel due to its low
cost and low molecular weight necessary for use in the nuclear thermal engines.
The nuclear thermal propulsion system is one of the key, distinguishable aspects of Centurion. The
technology was initially proposed in 1955 by the Hungarian engineer, Theodore von Karman [2]. This new system
was then tested in 1960 through the Nuclear Engine for Rocket Vehicle Application or NERVA program. These tests
validated the applicability of a nuclear thermal engine on rockets. The benefits for such an engine range from fuel
savings to reduced costs, and cut the cost of our missions by a factor of 3. The reactor core is composed of highly
enriched uranium–carbide fuel in a graphite matrix. Liquid hydrogen is injected into the core where it is heated to
above 2200°C and ejected out of the nozzle. The main concerns surrounding the use of a nuclear thermal propulsion
system include the political hurdles in gaining approval to send an active nuclear reactor into space and maintain the
safety of crew members from the intense neutron and gamma-ray radiation fields produced by the reactor. Radiation
concerns can be addressed through the application of radiation shields around the reactor and can further be reduced
through a combination of a tungsten and lithium hydrogen shield. The cost to develop this engine in 1971 was
estimated to be around $2.2 billion in FY1971 dollars, but within the past few decades, that price has gone down by
more than 90% due to increased research by numerous companies and development by Pratt & Whitney [3]. The
highly advanced nuclear thermal propulsion system is the solution that Hyperion Ventures’ proposes to address the
high costs and fuel associated with a mission to EML1 or EML2.
The lower module of Centurion, containing the nuclear thermal engines and the fuel, is completely covered
in radiation shielding and thermal shielding to protect the various components and computers aboard the vehicle, and
to protect the payload and other external and internal structures. Water will be used as a secondary cooling system to
ensure the engines do not overheat and to maintain a safe temperature. The primary concern for nuclear thermal
engines include crew safety and component deterioration due to radiation exposure. The nuclear thermal engines come
with their own radiation shielding to shield surrounding components from accidental radiation exposure. Additional
thermal shielding surrounds the fuel tanks and engines to protect the vehicle and other internal instruments. For our
missions, safety is a high priority, and has been taken into consideration in every aspect of Centurion’s design.

4
Figure 1-3. Detailed Illustration of Centurion OTV

5
The central and upper modules of the OTV consists of the communication systems, the Attitude
Determination and Control Systems, the power systems, sensors, and the docking module. Star trackers are used as
the absolute attitude determination sensors due to their accuracy [4]. Inertial Measurement Units (IMU) and sun
sensors are used as redundancy systems in the case of failure [5]. Thrusters and control moment gyroscopes have also
been implemented on Centurion due to the large amount of torque that can be generated and the fine attitude control
that will be essential when docking. Autonomous control systems will be used as control methods to analyze the sensor
data, implement control algorithms, and send instructions to the various actuators. Necessary computers have been
implemented on Centurion to handle these demands.
At the top of OTV is the current NASA Docking System (NDS). It was a docking system initially designed
by the United States in 1996 and redesigned in 2012 for future space exploration vehicles and serves as the
international spacecraft docking standard. It is also known as the international low impact docking system due to its
ability to dock safely and securely without damage to either vehicle [6]. It has been used in the past and is currently
in operation aboard the International Space Station. The system itself is androgynous, combining low impact docking
technology with the ability to both dock and berth. Once the payload and vehicle are docked, power, data, commands,
communications, water, and fuel can be transferred between the payload and vehicle, allowing for manned payload
missions. The NDS serves as the best docking system for Centurion mission due to its versatility and compatibility
with international standards, allowing Centurion to accommodate a wider range of payloads.
Figure 1-4. Illustration of Centurion OTV

6
Centurion’s primary mission is to transport cargo to and from Lagrange points EML1 or EML2, depending
on the mission. It will be docked with a refueling station in Low Earth Orbit (LEO) and will conduct its missions from
that base. For missions to L1, the OTV will take a minimum of 3.6 days to travel from LEO to EML1. For missions
to L2, the OTV will take a minimum of 5.2 days to travel from LEO to EML2. Once at EML1 or EML2, the OTV
will spend about 30 days to deploy and setup the payload. Once back in LEO, Centurion will refuel and will be ready
for its next mission.
Figure 1-5. Illustration of Centurion OTV Trajectory
In order to assemble and deploy the OTV to the refueling station in LEO, the vehicle will be launched from
Cape Canaveral, Florida to maintain the 28 degree inclination. This site has served as a great launching point for
many other missions in the past and serves as the perfect point to reach the desired inclination. This launch site is also
ideal for its location and limited collateral damage in the case of an emergency, allowing for debris to be dropped into
the ocean and to avoid human causalities. Centurion will be launched using a Delta IV launch vehicle for assembly.
It is expected to be in production by 2025 and would allow the mission to stay on track [7]. This vehicle would best
serve as our launch vehicle due to its large weight capacity and date for production. The launch vehicle would be
capable of carrying up to 125,000 kg into orbit and would be able to reduce the number of launches required to
assemble Centurion. For payloads, our design allows for a large range of various payloads to dock with the vehicle,

7
allowing it to be versatile and adaptable over time. These payloads can be launched using a variety of different launch
vehicles, but the Falcon 9 Heavy is recommended due to its versatility and compatibility.
Centurion uses some of the most advanced systems that puts it above the competition. From the nuclear
thermal engines to the modified NASA docking system, Centurion is at the forefront of space exploration. All of these
systems working cohesively together make up Centurion, a revolutionary new vehicle that will enable future deep
space missions.
1.1 Mission Timeline
Figure 1-6. Long Term Mission Timeline
Figure 1-6 showcases the long term timeline for the mission. The OTV is expected to be in operation by
2027. The major limiting factors in the development of our timeline is the development of the Nuclear Bi-Modal
thrusters and the further development of the NASA docking system. The nuclear thrusters should be in production by
2023. The NASA docking system is currently in production, but further research and development will be committed
to ensure a longer lifetime of the docking system. This should be completed by 2019, allowing us to maintain our
expected 2027 deadline for the launch of Centurion.
Orbital Assembly will begin around mid-2025 as mission tests are completed. Assembly of Centurion will
take place at Fort Lauderdale, Florida. Due to the small size of Centurion with its greatly reduced fuel mass as
compared to conventional solution, the entire completed assembly will be positioned on a Delta IV fairing. This would
allow for the most cost efficient launch and will only require a single launch to put Centurion in Orbit. Once in orbit,
Centurion will serve for a minimum of 5 years, transporting cargo to and from EML1 or EML2 for a minimum of 10
missions. The projection lifespan of Centurion far exceeds the 10 missions, but will be used as a guideline to maintain
the timeline. By 2032, Centurion will begin its orbital decommissioning. The OTV will take one final cargo to EML1.

8
After delivering its cargo, the OTV will reposition itself and head to EML4. At this point, the OTV will remain for
the duration of its life as to ensure safety from the nuclear material aboard the OTV and proper disposal of nuclear
fuel.
2. Orbital Systems
2.1. Design Approach
Hyperion Ventures is tasked with developing an Orbit Transfer Vehicle (OTV) capable of transporting
unmanned and manned payloads between Low Earth Orbit (LEO) and Earth-Moon Lagrangian points L1 (EML1) or
L2 (EML2) and back. Each mission to and from the Lagrange points can, at the most, take 6 days. Once at L1 or L2,
Centurion will need to stay there for at least 30 days, return to LEO, refuel and repeat the mission. 10 missions will
need to be completed in a matter of 5 years.
In order to get to L1, a few different orbital transfer methods were considered. First, the Hohmann Transfer
was considered, as it is the most efficient orbital transfer method (one that requires the lowest ΔV) to transfer between
two points in space. For three body dynamics, the Hohmann transfer will not give the most accurate results, but they
will be very close for initial estimates [8]. The reason for the inaccuracy is that as the spacecraft gets closer to the
Moon, its path will be perturbed by the Moon’s gravitational field. Next, the invariant manifolds were considered as
a low energy transfer option. The Earth-Moon manifolds, however, are not accessible from LEO. The manifolds are
generally over 75,000 km above Earth’s surface [8], so it would require two burns to get to L1 and another burn to
insert the spacecraft into a halo orbit. Finally, another three burn transfer was considered, where the initial LEO
departure burn would take the spacecraft close to the Moon, and then the spacecraft would make another burn to get
itself on a path to L1, then make one more burn to insert itself into an orbit around L1. For transfers to the second
Lagrange point, the same general process was followed.
The orbits that will be considered at the Lagrange point will be halo orbits and quasi halo orbits -- also known
as Lissajous orbits. Because Centurion does not need to stay at the Lagrange point for a very long amount of time,
Centurion will only be able to orbit a Lagrange point a couple times, as long as the orbit has a large enough period.
By extension, station-keeping costs when in an orbit around L1 or L2 will be small, because of the short amount of
time in orbit.

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2.2. Concept Development
It takes less time to reach L1 then it does to reach L2 because of the relative distances of each. A direct transfer
to L1 requires about 4 days, with a ΔV around 3.7 km/s. An efficient transfer to L2 requires about 6 days, with a ΔV
around 4.0 km/s. From these figures alone, the transfer time to get to L2 is just under the maximum required time limit
of a transfer to or from a Lagrange point. An investigation into the various trajectories and halo orbits to get to L1 and
L2 will be discussed in the following section.
Trajectory to L1
To get to L1, three different trajectories were considered. Fist, a simple Hohmann transfer was used to get to
the Lagrange point. In order to calculate the Hohmann transfer ΔV and the TOF required to get there, the vis-viva
equation was used. A modified form of this equation is shown in Equation (2-1),
ΔVH = √𝜇 (
2
𝑟1
−
1
𝑎 𝐻
) − √𝜇 (
1
𝑟1
) (2-1)
where 𝑟1 is the radius of the perigee of the Hohmann transfer, 𝑎 𝐻 =
𝑟1+𝑟2
2
is the semi major axis of the Hohmann
transfer (with 𝑟2 being the distance from the Earth to L1), and 𝜇 is the gravitation parameter of the Earth [9]. When
the Moon is closest to the Earth, the distance between the two is around 362,600 km, and therefore points L1 and L2
lie at 308,079.4 km and 423,198.8 km respectively. These Lagrange points were found by computing the force balance
between the Earth and the Moon, and seeing where those two forces are in equilibrium at each point (see MATLAB
code Lag1.m and Lag2.m for computation). With these values, a Hohmann transfer will have a ΔV of 3.0616 km/s.
The second burn at apoapsis of the transfer will be the insertion burn into the halo or quasi-halo orbit. The TOF is
another variable of interest, because the transfer time has to be less than 6 days. Equation (2-2) shows the TOF for a
Hohmann transfer.
𝑡 𝐻 = 𝜋 (
𝑎 𝐻
3
𝜇
)
1
2
(2-2)
where all the parameters are the same as in Equation (2-1). The time of flight for a Hohmann transfer to L1 with the
Moon at periapsis is 3.597 days, which is well within the 6 day limit. In Table 2.1, proposed halo orbit insertion burns
(represented by ΔV2) for different types of orbits are shown. These types of orbits were taken from [10], where orbit
1 is a halo orbit, orbit 2 is a medium quasi-halo orbit, and orbit 3 is a large quasi-halo orbit. Originally, the orbits

10
proposed started from an Earth altitude of 200 km, but substituting in a 400 km burn, ΔV1 will decrease and thus total
ΔV will decrease as well.
Table 2.1. Hohmann Transfer Burns and Insertions, adapted from [10]
Orbit Option 𝚫𝐕𝟏 (km/s) 𝚫𝐕𝟐 (km/s) Total 𝚫𝐕
1 3.059 0.5807 3.6397
2 3.059 0.5646 3.6236
3 3.059 0.5491 3.6081
The next transfer method considered is the three burn transfer around the Moon. The first burn is the initial
burn from LEO to a 200 km lunar altitude, the second burn is from the 200 km lunar altitude to the Lagrange point,
and the third burn is the insertion burn into the halo orbit. Using data from [10], it can be seen in Table 2.2 that there
is an opportunity to use less energy to get to L1 while still being under the required time limit, and this means less
fuel consumption thus driving the cost down.
Table 2.2. Delta V's for different orbits, table adapted from [10]
Orbit Option 𝚫𝐕𝟏 (km/s) 𝚫𝐕𝟐 (km/s) 𝚫𝐕𝟑 (km/s) Total 𝚫𝐕 (
𝐤𝐦
𝐬
) TOF
(days)
4 3.117 0.26887 0.30131 3.6872 4.87
5 3.117 0.27180 0.2721 3.66701 5.09
6 3.117 0.27158 0.24842 3.6370 5.42
7 3.117 0.26690 0.21310 3.5970 5.80
Note that these results were based on an initial 200 km altitude LEO, so when calculated with an initial 400 km altitude
orbit, the totals for fuel consumption should go down across the board, and the TOF will increase.
2.2.1.1. STK Analysis to L1
In terms of modeling a two burn transfer to L1, STK/Astrogator was used. A tutorial was initially used to
simulate the satellite’s path to L1 [11], and then that tutorial was modified for this project’s specific parameters. The
Launch targeter was changed to an initial conditions targeter, where a perigee altitude of 400 km and a 28 degree
inclination were used to find the orbit from which Centurion would need to start. The Astrogator targeter then found
the trajectory to get to L1, used a differential corrector to model a 3 revolution halo orbit, and exited the Lagrange
point orbit to get back into a circular orbit around the Earth at a 400 km altitude and a 28 degree inclination.
Multiple trajectories were modeled in STK to simulate different halo orbits in L1. Table 2.3 shows the data
obtained from these simulations, in which only the z-amplitude was varied. ΔV1 is the halo orbit insertion burn, ΔV2

11
is the halo orbit exit burn, and ΔV3 was the total mission ΔV (to and from L1). The initial burn to LEO and from LEO
was 3.069 km/s and 3.058 km/s respectively.
Table 2.3. Total mission ΔV for varying z-amplitudes
Orbit Option Amplitude (km) 𝚫𝐕𝟏 (m/s) 𝚫𝐕𝟐 (m/s) 𝐓𝐨𝐭𝐚𝐥 𝚫𝐕𝟑 (m/s)
8 5,000 620.017 644.108 7411.944
9 7,500 622.213 644.063 7421.159
10 10,000 625.070 647.151 7422.128
11 15,000 632.660 654.051 7446.725
12 20,000 642.570 663.845 7482.638
Note that the simulations included two station-keeping burns, which will be discussed in a later section. The
highlighted option shows that the orbit with the 5,000 km z-amplitude had the lowest total ΔV, and thus the orbit that
will be utilized in our mission. Table 2.4 shows the transfer times to and from L1, as well as the amount of time
Centurion spent in the halo orbit. Each halo orbit modeled was 3 revolutions long.
Table 2.4. Transfer times to L1 and halo orbit durations for varying z-amplitudes
Orbit Option Amplitude (km) 𝐓𝐢𝐦𝐞 𝐭𝐨 𝐋𝟏 (𝐝𝐚𝐲𝐬) 𝐓𝐢𝐦𝐞 𝐋𝟐 (days) 𝐇𝐚𝐥𝐨 𝐎𝐫𝐛𝐢𝐭 (𝟑 𝐫𝐞𝐯)
8 5,000 4.482 4.227 35.999
9 7,500 4.352 4.352 36.146
10 10,000 4.348 4.349 36.108
11 15,000 4.340 4.342 36.079
12 20,000 4.329 4.327 36.140
Each trajectory simulated was in the halo orbit for over 30 days, and each transfer was under 6 days meeting all the
requirements. Figure 2-1 and Figure 2-2 show the trajectory of orbit option 8 as seen from different reference frames.
Figure 2-1 shows the trajectory to L1 from a rotating reference frame. Because the Moon’s orbit is not
circular, the Moon’s position with relation to the Earth will move back and forth, and the position of L1 will do the
same. Figure 2-2 shows the trajectory as viewed form the Earth-Moon plane. A zoomed in view of just the halo orbit
is shown in Figure 2-3. This specific halo orbit had a z-amplitude (above the Earth-Moon plane) of 5,000 km.

12
For this simulation, a Hohmann like transfer to L1 was automatically performed by the targeter, and the initial
burn had a ΔV of 3.069 km/s to get to L1. As you’ll recall, the analytic solution using a Hohmann transfer to L1 was
3.0616 km/s, which is a difference of only 0.0074 km/s from the simulation. This means that the results obtained from
calculating a simple Hohmann transfer were very accurate.
Trajectory to L2
For a maneuver to L2, similar methodology as the maneuver to L1 was used, where a few types of transfers
were considered: the Hohmann transfer, and the three burn transfer. The three burn transfer around the Moon was
almost immediately ruled out because the TOF was too large. The fastest transfer using this method took 14.7 days,
which is well over the 6 day limit for a transfer [10].
Next, a straight Hohmann transfer to L2 was computed. When the Moon is closest to the Earth, the distance
from the Earth to L2 is 423,198.8 km. Using Equation (2-1), the ΔVH was calculated to be 3.0920 km/s, a slightly
higher ΔV than for a transfer to L1. Using Equation (2-2), the time of flight to L2 would be 5.741 days, which is just
barely under the 6 day limit. Much like the trajectories considered to L1, this trajectory also takes into account that
the L2 insertion happens when the Moon is closest to the Earth. Of course, this Hohmann transfer calculation does not
take into account the Moon’s gravity, and the perturbations that it causes. That being said, the total ΔV when
considering the gravitational effects will be lower. This Hohmann transfer is the lowest energy transfer to get to L2.
Some transfers have been proposed that would put the spacecraft into a low lunar altitude then the velocity of the
spacecraft would then go around the Moon and to the Lagrange point. These transfers, first proposed by Robert
Farquhar in 1972, were also ruled out because of the TOF being around 212 hours or 8.83 days – well over the 6 day
limit [12].

13
Figure 2-1. Side view of the trajectory to L1
Figure 2-2. View of L1 from the perspective of the Moon
Figure 2-3. Zoomed view of the halo orbit at L1

14
In terms of the halo insertion burns, those could be on the order of 800 m/s [10] [13]. Possible halo orbits
and quasi-halo orbits around L2 have been proposed, with periods of around 15 days. This means that only 2 full
revolutions would be required to meet the minimum 30 day limit at the Lagrange point. The z-amplitudes for insertion
could range anywhere from 2,600 km to 11,000 km for a quasi-halo orbit, and 8,000 km for a halo orbit [13]. This
means that the required burn at LEO would not have to change much from the original calculation, bringing the total
ΔV of the L2 mission to around 7.8 km/s.
2.2.2.1. STK Analysis to L2
The trajectory was modeled to L2 using STK/Astrogator, and it used the same start conditions as the model
to L1. The orbit started out in a 400 km altitude, 28 degree inclined orbit, and then the orbit was propagated to EM-
L2. The only difference was the Trans-Lunar Injection (TLI) ΔV was changed to reflect the increased distance to L2
in relation to L1.
Multiple trajectories were modeled to L2, much like the procedure for L1. Table 2.5 shows different mission
ΔV’s for varying halo orbits. ΔV1 represents the halo orbit insertion burn, ΔV2 is the halo orbit exit burn, and total ΔV
is the total mission ΔV to L2 and back. The burn from LEO is 3.094 km/s, and the return ΔV is 3.099 km/s.
Table 2.5. Total mission ΔV for varying z-amplitudes
Orbit Option Amplitude (km) 𝚫𝐕𝟏 (m/s) 𝚫𝐕𝟐 (m/s) 𝐓𝐨𝐭𝐚𝐥 𝚫𝐕 (m/s)
13 5,000 1124.806398 1018.229755 8352.291161
14 7,500 1122.051898 1030.947916 8372.583998
15 10,000 1118.016671 1003.734879 8339.402704
16 15,000 1106.846749 988.600830 8321.101857
17 20,000 1097.023643 974.495641 8306.030167
These simulations also included two station-keeping burns, which will be discussed in more detail later.
Highlighted is the orbit with the lowest mission ΔV, and thus the orbit chosen for the mission. Table 2.6 shows the
halo orbit duration and the time to and from L2.

15
Table 2.6. Transfer times to L2 and halo orbit durations for varying z-amplitudes
Orbit Option Amplitude (km) 𝐓𝐢𝐦𝐞 𝐭𝐨 𝐋𝟏 (days) 𝐓𝐢𝐦𝐞 𝐋𝟐 (days) 𝐇𝐚𝐥𝐨 𝐎𝐫𝐛𝐢𝐭 (𝐝𝐚𝐲𝐬)
13 5,000 5.269 5.868 44.277
14 7,500 5.267 5.860 45.089
15 10,000 5.266 5.895 45.171
16 15,000 5.269 5.925 44.985
17 20,000 5.275 5.958 44.841
Note that every orbit simulated had a time back from L2 under 6 days and each halo orbit lasted for 44 to 45 days,
thus meeting the requirements. In terms of modeling this orbit,
Figure 2-4 shows entire trajectory of orbit option 17, and
Figure 2-5 shows a closer view of the 20,000 km amplitude, 3 revolution halo orbit.
Figure 2-4. L2 trajectory as seen from the Earth
Figure 2-5. L2 halo orbit as seen from the Moon

16
Figure 2-6. L2 trajectory in the Earth-Moon rotating axis
Seen in
Figure 2-6 is a view of the Earth and the Moon as seen from the Earth-Moon rotating axis, with Centurion
almost at its closest approach to the Moon and the Earth. Because the Moon is shifting, L2 will also shift from being
closer and farther away from the Earth. Therefore Centurion, if it is to orbit L2, will not stay in the same position with
relation to the Earth. This specific halo orbit’s period is 44.841 days.
Station-keeping Analysis
Since Centurion will only be in a halo or quasi-halo orbit for a minimum of 30 days, station-keeping costs of
our mission will be very low, on the order of 10 m/s [13]. The STK targeter included a station-keeping maneuver.
Table 2.7 shows all the station-keeping values obtained for all the L1 trajectory simulations.

17
Table 2.7. Station-keeping values for L1 halo orbits
Orbit Option 𝚫𝐕𝟏 (m/s) 𝚫𝐕𝟐 (m/s)
8 9.586 9.952
9 13.588 13.069
10 7.352 14.37
11 7.834 24.048
12 10.311 37.805
Table 2.8 shows the station-keeping burns needed to keep Centurion in an orbit around L2. These values were also
obtained using the STK simulations.
Table 2.8. Station-keeping values for L2 halo orbits
Orbit Option 𝚫𝐕𝟏 (m/s) 𝚫𝐕𝟐 (m/s)
13 8.385644 6.922246
14 8.729282 16.927658
15 9.412483 14.138888
16 12.764731 18.727220
17 17.744123 22.860030
Note that all of these values are very low, considering the entire mission is on the order of multiple km/s, and these
values are all on the order of 0.01-0.02 km/s. These values are by no means negligible, but they are low enough so
that a variation in any of these numbers would not be detrimental to the mission design.
Orbital Maintenance in LEO
When our spacecraft gets back into LEO after a mission from L1 or L2, the spacecraft will stay there for
approximately 4 to 4.5 months. The orbit would degrade over that period approximately 12 km in altitude, assuming
that the vehicle starts at a 400 km altitude orbit with an empty fuel tank. The meaning the ΔV required to get back to
a 400 km altitude orbit would be 6.7974 m/s. The amount of fuel lost would be on the order of 50 kg, using (2-3). If
instead the vehicle refueled immediately when back in LEO, the orbit would only degrade 3 km due to the larger
overall mas, and the ΔV required to get back to the 400 km orbit would be 1.697 m/s. The amount of fuel lost due to
just the burn would be on the order of 10 kg At first glance it seems that refueling immediately would save the most
fuel, but the amount of fuel boiled off needs to be accounted for. Centurion would lose on the order of 2000 more

18
kilograms of fuel due to boil off if it were to refuel immediately. Therefore, the vehicle will refuel at the end of the 4
months between missions to save the most fuel.
Aerobraking maneuvers
Included in this section are the details on aerobraking - the calculations and feasibility of the aerobrake from
L1. The time required to get back from L2 will be too close to the maximum allowable limit to even consider
aerobraking. Some things that need to be considered for this proposal are the risks involved with aerobraking, the time
it takes to aerobrake, and the overall effect on the entire system.
In order to calculate the savings in fuel of the aerobrake, a simple equation was used,
ΔV = kρo√2𝜋𝜇 (
1 + 𝑒
√ 𝑒
) √𝐻 (2-3)
where k = ballistic coefficient of heat shield, 𝜌 𝑜 = density of air at periapse, 𝜇 = the gravitation constant of the earth,
𝑒 = eccentricity of the orbit, and 𝐻 =
𝑅𝑇
𝑔
is the scale height, where T is the temperature at periapsis. The ballistic
coefficient is a function of the coefficient of drag, surface area, and the mass of the vehicle, namely,
𝑘 =
𝐶 𝑑 𝑆
2𝑚
(2-4)
where 𝐶 𝑑 is the coefficient of drag, S is the surface area, and m is the mass [14].
Using values from the HPOP STK density models and thermodynamic tables at different altitudes and initial
guesses for the spacecraft and orbital trajectories - a heat shield diameter of 10 meters, coefficient of drag of around
2, total mass of around 70,000 kg, and an eccentricity of .96 - values for the ΔV of an aerobrake come out to very low
numbers. Table 2.9 shows the ΔV values and the altitude at which they were calculated.

19
Table 2.9. Delta V's of possible altitudes at which to aerobrake with a 10 meter heat shield.
Altitude (km) Density (kg/km^3) 𝚫V (m/s)
50 102700.0 1,026.4
60 30960.0 295.6152
80 18449.456 157.9763
100 560.276 4.7543
120 22.234 0.2563
140 3.839 0.552
Note that the Jacchia 1960 density model from STK was used for altitudes above 80 km, and thermodynamic tables
from [15] were used for altitudes below 80 km. Also, these values were computed using a MATLAB code, which can
be found in Appendix 12.1. The only way that aerobraking will be able to have a tangible effect on our system is that
if the ΔV savings were on the order of 1 km/s, which only happens at an altitude of 50 km. This is way too low for
our spacecraft if we want the spacecraft to stay intact and be reusable, because the density at that altitude would be
very large and thus cause the spacecraft to heat up and either start breaking apart or fail completely.
End of Life Summary
Since Centurion has an active nuclear thermal propulsion system, a viable and safe end of life plan has been
implemented so that the vehicle will never again be able to get near the atmosphere of the Earth. Therefore, at the end
of 10 missions, Centurion will be able to take a payload on a one way trip to L1, drop it off, and then maneuver to the
Earth-Moon Lagrange Point 4 (L4). L4 was chosen because it is stable, unlike L1, L2, and L3. Anything within a
certain vicinity of L4 will stay there, and therefore it is a good spot to place Centurion. Since this is only a one way
mission, the payload to L1 would probably be some sort of satellite or science mission that would study the moon, or
beyond.
The ΔV required to get to L2 was found by using a minimum energy transfer trajectory using Lambert’s
theorem [9]. The transfer trajectory ΔV was found to be around 682 m/s from L1, bringing the total mission ΔV to
4.914 km/s for a mission to L4 with a pit stop at L1. Note that the transfer ellipse was calculated using the Moon’s
sphere of influence and therefore doesn’t take into account three body dynamics. For this reason the ΔV is not
completely accurate, but is a good first cut estimation of the ΔV required to get there. Even so, the ΔV required to get
to L4 would be far less than the ΔV required to return to LEO from L1.

20
2.3. Critical Design Issues
Computation of Halo Orbits and Insertion Velocities
There are many different ways to compute halo orbit periods and trajectories, but all of these methods are
complex. These periodic orbits are solutions to the circular restricted three body problem (CR3BP). Once an
approximation is achieved (by way of finding different amplitudes and constants associated with the orbit), a full
solution to the CR3BP can be found through various techniques [9]. These solutions and their initial guesses will be
tabulated and analyzed to determine the best possible transfer and orbit combination for the mission.
Computation of Invariant Manifolds
The invariant manifolds that exist between the Earth and the Moon, or any three body system in general, are
associated with a type of periodic solution to the CR3BP, i.e. a halo orbit. So once a possible orbit is known, the
invariant manifolds can be computed based on this [9]. As mentioned earlier, for the Earth-Moon case, the invariant
manifolds exist some 75,000 km above the Earth’s surface, and are therefore not able to be immediately utilized. A
transfer between LEO and these invariant manifolds will be considered as a possible transfer option. A transfer
between LEO to Low Lunar Orbit and then to the invariant manifolds to L2 will also be considered.
3. Propulsion Systems
3.1. Design Approach
To engineer a propulsion system that must carry a 22,500 kg payload from Low Earth orbit to an Earth-Moon
Lagrange point in an efficient and timely manner is no small feat. Centurion Orbital Transfer Vehicle (OTV) must be
able to travel to and from LEO to either Earth-Moon Lagrange point 1 (EML1) or Earth-Moon Lagrange point 2
(EML2) a minimum of five times. A one way trip must be executed in six days or less. To accommodate the constraint
of a six day transit time to and from EML1 or EML2 Centurion must be able to produce a total of 8.31 kilometers per
second of ∆V per round trip, section 2. Additionally, Centurion’s propulsion systems must be able to complete all
mission requirements while using minimal propellant mass.
Centurion will have two independent propulsion systems. The two systems will be the main propulsion
system and the attitude control propulsion system. The main propulsion system will be responsible for all major orbit
transfer maneuvers. Such maneuvers include: leaving LEO, arriving at the Lagrange point, leaving the Lagrange point,
and returning to LEO. For the main propulsion system a Bimodal Nuclear Thermal Rocket (BNTR) called Escort has

21
been selected. Separate from the main propulsion system will be the attitude control propulsion system. The attitude
control propulsion system will be responsible for coarse attitude adjustments and will work in conjunction with control
moment gyroscopes to meet all of the requirements of the attitude control system. For the attitude control propulsion
system Aerojet R-1E bipropellant thrusters have been chosen.
3.2. Concept Development
While developing the concept of Centurion several types of propulsion were considered for both the main
and attitude control systems. For the main system, ion chemical, and nuclear thermal propulsion technology were
considered. When developing the attitude control system, ion, cold gas, and chemical were considered.
Main Propulsion System
Technologies considered for use in the main propulsion system were compared using a few key parameters,
some of which can be found in the Risk Analysis Section. First was the fuel mass required to perform a round trip.
Because of the large payload any solution will require a large fuel mass. By placing fuel consumption as a primary
design driver, second only to safety, it could be ensured that Centurion would have the most cost effective design.
Second was the thrust rating of the engine. Again, with a large, potentially manned, payload this was an important
factor. Too large of a thrust could be harmful to the crew and too small of a thrust could make the mission take too
long. Thus, the last major consideration was the time it would take to achieve the largest ΔV, however, this is not
included in Table 3.1 because it was only useful for ruling out ion thrusters.
Table 3.1. Potential Main Propulsion System Technologies [16] [17] [18] [19] [20] [21] [3] [22]
Type Thruster Isp
[s]
Propellant Max Thrust
[N]
Fuel Mass to L2
[kg]
Ion Aerojet NEXT 4100 Xenon 0.235 8,700
Busek BHT-20k 2320 Xenon 0.807 16,600
NASA NSTAR 3195 Xenon 0.094 11,600
Bipropellant CALT YF-73 420 LOX/LH2 44,150 229,000
Astrium Aestus 324 N2O4/MMH 29,600 436,000
Aerojet CECE 465 LOX/H2 111,000 183,000
Monopropellant Aerojet MR-80B 225 Hydrazine 3780 1,410,000
AMPAC MONARC 445 235 Hydrazine 445 1,190,000
Nuclear Thermal CIS NTR 955 LH2 66,700 47,500
NERVA XE 850 LH2 1,112,000 62,500
Escort 911 LH2 333,600 49,600
Table 3.1 illustrates the strengths and weaknesses of each propulsion technology considered. Ion thrusters
are the most fuel efficient due to their high specific impulses. However, they produce so little thrust that thousands of

22
them would be required to be competitive with the weakest chemical thruster considered. Bipropellant chemical
thrusters provide appropriate levels of thrust but their specific impulse is comparatively low. Because of this the fuel
mass required for a round trip is prohibitively large. Monopropellant chemical thrusters follow a similar pattern, with
even lower specific impulses than bipropellants and lower thrust ratings they were quickly out of the running. Nuclear
thermal propulsion technology has none of these issues. With specific impulses nearing 1000 seconds and thrust
ratings that rival bipropellant thrusters these engines are the perfect match.
3.2.1.1. Nuclear Thermal Propulsion Systems
Nuclear thermal propulsion systems are fairly simple. In a standard thrust producing system there are three
main parts; the fuel tank and feed system, the reactor, and the nozzle. The fuel tank and fuel system house the liquid
hydrogen propellant and delivers it to the reactor. The reactor provides heat to expand the liquid hydrogen. Once
heated, the hydrogen is forced through the nozzle to produce thrust just like in a conventional chemical thruster. In
fact, the only aspect of a nuclear thermal propulsion system that differs from a traditional chemical system is that
nuclear systems are driven by nuclear fission rather than chemical combustion. The advantage of nuclear propulsion
comes from the fact that nuclear fission occurs at higher temperatures than combustion reactions. This increase in
temperature directly translates to the increased fuel efficiency of nuclear systems.
There are four types of thermonuclear propulsion systems that can be used. As a baseline there is the simple
thrust producing version. Second is the liquid oxygen augmented nuclear thermal rocket (LANTR), Figure 3-1. This
Figure 3-1. Basic LANTR schematic [23]

23
system produces more thrust than conventional NTR systems by using liquid oxygen as an afterburner [23]. Third is
the bimodal system. This is the type of system that has been selected for use on Centurion. Bimodal systems make use
of the high temperatures in the reactor to generate electric power by means of a closed Brayton cycle generator. Finally
there is the trimodal system. This system uses the liquid oxygen afterburner and closed Brayton cycle generator for
increased thrust production while generating power for the vehicle as well [24].
3.2.1.2. Past and Present Nuclear Thermal Systems
The Rover and NERVA (Nuclear Engine for Rocket Vehicle Application) programs stand as the most
significant endeavor in creating a flight ready nuclear thermal propulsion system to date. In 1955 the Rover program
began investigating the feasibility of using nuclear reactors for space propulsion. Out of Los Alamos National Labs
the Rover program succeeded in creating a series of liquid hydrogen cooled nuclear reactors called Kiwi. These
reactors served as the basis for the NERVA program. In 1961 NERVA began using the Kiwi reactors to create flight
ready nuclear thermal propulsion systems [25]. Aerojet and Westinghouse were contracted to develop the flight ready
systems. Early on it was decided that the most important design driver should be safety. After safety the team was
concerned with producing a specific impulse of around 760 seconds, the capability of the engine to start without
external aid (called a bootstrap start), and the engine should be capable of producing around 337kN of thrust while
minimizing weight [2].
Two sets of engine tests were conducted throughout the lifetime of the NERVA program. First was the NRX
(Nuclear Reactor Experimental) series of tests conducted in February 1966. This series of tests was concerned with
proving the capability of a bootstrap start, investigating the stability of the system during a wide variety of operational
modes, and observing the endurance of reactor components. These objectives were all achieved. The reactor was
started a number of times under varying conditions and was shown to be highly controllable and predictable. Overall
the NRX series was highly successful and led to the Ground Experimental Engine (XE) test series [2]. The XE engine
was designed to be flight ready and represents the most complete systems ever constructed. Testing of the XE series
went much the same as the NRX series. A total of 28 bootstrap starts were accomplished and the engine ran for a total
of 115 minutes with no sign of failure. The tests were a complete success [26].
Current nuclear thermal propulsion concepts are based extensively on the NERVA program’s findings but
are focused on smaller and more fuel efficient systems. Advances in nuclear fuels have led to the ability to produce
higher chamber temperatures which increases fuel efficiency [3].

24
3.2.1.3. Escort System Specifics
Centurion will be using a system proposed by Pratt and Whitney called the Escort system. Escort uses three
bimodal nuclear thermal propulsion units to provide thrust and power to the vehicle. Each unit is designed with its
own closed Brayton cycle generator for power production as well as shielding and all necessary turbomachinery.
Escort also comes equipped with radiators capable of dissipating the large thermal load generated by the fission
reactor. In the reactor the fission of 235
U is used to generate thermal energy. The uranium is suspended in a tungsten
cermet (W-UO2) [27]. Tungsten is a dense element capable of withstanding high temperatures and can mitigate the
effects of gamma radiation. In addition tungsten is highly resistive to corrosion due to hydrogen, which increases the
durability and lifetime of the reactor.
In the foldout on page 26, there is a basic representation of the Escort system. Beginning at the liquid
hydrogen tank in the upper left corner the propellant is drawn out of the tank by means of a turbopump driven by an
expander cycle. To accomplish expansion and drive the pump liquid hydrogen is first injected into the nozzle and runs
up the nozzle and into the control drum. This cools the nozzle and control drum as well as drives the pump. After the
expansion the propellant is injected into reactor where it is heated and expanded further to produce thrust.
Separately, the closed Brayton cycle generator uses a fluid mixture of helium and xenon to extract heat from
the reactor to generate electric power. A series of valves allows the HeXe mixture to flow from its storage tank into
the reactor where it is heated and then forced through a turbine to generate power. Once through the turbine, the fluid
enters a radiator where the majority of the remaining heat from the reactor is dissipated.
Inside the reactor are the hexagonal tungsten uranium dioxide fuel elements. A cross section of these fuel
elements can be seen in the fold out. The yellow circles on the cross section represent the paths that the hydrogen
propellant takes through the reactor. In the center of the image the green circles represent the coaxial flow paths of the
helium xenon generator fluid. And the red area is the W-UO2 cermet fission material.
In the bottom right of the fold out is a model of the reactor and generator system. Of particular importance
in this image is the external shield on top of the reactor. This shield is made of lead and protects any potential crew
from the harmful effects of gamma rays produced in the reactor.

25
3.2.1.4. Nuclear Reactor Safety
Nuclear thermal propulsion is the enabling technology that will lead to more ambitious missions to more
distant places in the solar system. Unfortunately, the public opinion of nuclear technology is that it is dangerous and
we should stay away from it. But the truth is that when handled properly thermonuclear rocket systems are no more
dangerous than conventional chemical propulsion systems. That being said, there are numerous precautions that must
be followed to ensure safety while handling nuclear devices.
A main point of concern is mitigating the effects of the nuclear radiation being emitted from the reactor core.
There are three types of radiation that must be dealt with; alpha particles, beta particles, and gamma rays. Both alpha
and beta particles are not damaging and can be easily stopped with a thin sheet of aluminum [28]. Gamma rays are
high energy (~1 MeV) photons that are emitted as a byproduct of fission. The gamma rays emitted from fission of
235
U have energy of about 13.3 MeV. Considering the reactor runs at a peak power of around 500MW, the level of
gamma radiation from the reactor is potentially dangerous [29]. However there are numerous ways in which these
affects will be mitigated and kept to safe levels. First, there is a 3cm thick lead radiation shield positioned above the
reactor. Second, when docking with the payload or while nearby any life forms, the reactors will be run at a
dramatically decreased rate such that a safe distance from the reactor will be greater than or equal to 25 meters.
Other safety protocols are concerned with the timing of the operation of the nuclear fission reactors. When
launching Centurion, the reactors will not be run in a critical state before leaving the atmosphere. This will ensure that
if any problems occur they will not endanger any population on Earth. In addition, the reactors will not be allowed to
return to Earth after being run at a critical state [65]. For this reason, at the end of the ten missions Centurion will be
placed at EML4 indefinitely.
3.2.1.5. Main Propellant/Tankage
A number of inert gasses could be used in conjunction with the nuclear thermal propulsion system to
produce thrust. However, when equation (3-1) is considered, it is clear that Hydrogen is the ideal choice because it
has the lowest molecular weight M [65].

27
𝑰 𝒔𝒑 = 𝑨𝑪 𝒇√𝑻 𝒄 𝑴⁄
(3-1)
The parameters A and Cf are constants and properties of the fuel and nozzle respectively. The important part of the
equation is under the radical. Tc is the temperature of combustion in the “combustion” chamber and M is the molecular
weight of the fuel. In this case the combustion chamber is where the inert hydrogen gas is heated to produce thrust,
the hydrogen is not combusted. The lower the molecular weight of the gas being expelled from the nozzle, the higher
the chamber temperature then the specific impulse will be greater. For this reason, and considering Hydrogen gas has
the lowest molecular weight of all gasses, Hydrogen was chosen as the fuel for the main propulsion system of
Centurion. To calculate the mass of fuel needed to make a round trip to the L1 and L2 the following equations were
used:
𝑴 𝒑 = 𝑴 𝒑𝟏 + 𝑴 𝒑𝟐 + 𝑴 𝒑𝟑 + ⋯ + 𝑴 𝒑𝟏𝟏 (3-2)
where
𝑴 𝒑𝒊 = ∑ 𝑴𝒊(𝒆
∆𝑽 𝒊
𝒄 − 𝟏) (3-3)
Mp is the total mass of the propellant and each Mpi is the mass of the propellant needed to achieve each major burn. Mi
is the combined mass of the structure, payload, and fuel, ∆Vi is the change in velocity needed to make a major
maneuver, and c is the exit velocity of the propellant. A tabulated version of this calculation is provided in Table 3.2.
The four major orbit transfer burns occur at stages 1, 3, 9, and 11. Theses maneuvers represent 99% of total fuel
consumption for a trip to L2. The remaining losses are due to halo orbit correction burns and boil off.
Table 3.2. Fuel Consumption to and from L2 [Isp = 911s]
Stage Description ∆V [m/s]
Time
Elapsed
Mi (Includes
Payload) [kg]
Mpi Propellant
Spent [kg]
1 Departing LEO 3095 10 Minutes 87,745 25,682
2 Transit to L2 0 5.3 Days 62,063 42
3 Arrive at L2 1097 3 Minutes 62,021 7,165
4 Halo orbit 1 0 15 days 32,177 83
5 Halo Correction 1 18 6 Seconds 32,094 64
6 Halo Orbit 2 0 15 days 32,030 83
7 Halo Correction 2 23 6 Seconds 31,947 82
8 Halo Orbit 3 0 15 days 31,866 83
9 Depart L2 974 2 Minutes 38,588 3,986
10 Transit to LEO 0 6 days 34,602 25
11 Arrive at LEO 3099 4 Minutes 34,577 10,133
Final Mass = 24,444 ∑ 𝑀 𝑝𝑖 = 47,428

28
To store the liquid hydrogen propellant, Centurion will make use of a custom fabricated 700 cubic meter
aluminum tank with active and passive thermal control. Hydrogen must be stored at or below 20 Kelvin in order to be
a liquid. Such a low temperature is possible to be maintained but requires robust thermal control. The Escort system
is designed to operate alongside a zero boil off cryogenic storage system. However, in order to better model a real
system Centurion was designed to compensate for boil off rate of 1% loss per month. For a trip to EML2 this comes
to a total loss of just over 316 kg of propellant.
Attitude Control Propulsion System
The attitude control propulsion system will be responsible for providing small amounts of ΔV needed to
adjust the orientation of Centurion. Thrusters will be used in conjunction with control moment gyroscopes to
accomplish the goals of the attitude control system. The propulsive portion of the attitude control system will mainly
be used for slewing maneuvers while docking and refueling.
Table 3.3. Potential ACS propulsion technologies [16] [17] [30] [31] [32]
Type Engine Fuel Isp
[s]
Thrust
[N]
Propellant
Mass[kg]
Ion Aerojet NEXT Xenon 4100 0.235
Busek BHT-20k Xenon 2320 0.807
Cold Gas MOOG 58-118 Unknown 72 3.5 560
AMPAC SVT01 Xenon 45 0.05 900
Monopropellant AMPAC
MONARC -90
Hydrazine 235 90 170
Aerojet MR-107N Hydrazine 232 109-296 180
Bipropellant
EADS 10N
NTO, MON-1, MON-
3 and MMH
291 10 140
Aerojet R-1E MMH/NTO 280 111 144
Table 3.3 shows technologies considered for use in the attitude control propulsion system. Ion thrusters were
considered for their high specific impulse. However, they were not chosen based on their low thrust (<1N). Cold gas
thrusters are the simplest type of thruster and are thus the most reliable. Unfortunately they are not very efficient and
for this reason were not chosen. Monopropellant and bipropellant chemical thrusters are very comparable in their
reliability but bipropellants are more efficient. For this reason the Aerojet R-1E thrusters were chosen.

29
3.2.2.1. Aerojet R-1E
The Aerojet R-1E is a versatile and reliable thruster.
Previously used on the space shuttle this thruster is dependable
and flight proven many times over. Each thruster has a mass of
just 2 kg and can produce a steady state thrust of 111N. Every
thruster is capable of firing 330,000 times with no limitations on
the duration of the burn [30]. By using these engines for the
attitude control propulsion system in conjunction with control
moment gyroscopes a two point failure system has been created. This ensures that Centurion will always be able to
adjust its attitude.
3.2.2.2. ACS Propellant/Tankage
As a bipropellant thruster the Aerojet R-1E requires a mixture of Mono Methyl Hydrazine (MMH) and
Nitrogen Tetra Oxide (NTO), where MMH is the fuel and NTO is the oxidizer. These fuels mix optimally at a mass
mixture ratio (O:F) of 1:6. To calculate the mass of propellant required for the attitude control propulsion system a
ΔV of 10 m/s was used. This represents the total amount of ΔV required for the entire lifetime of the OTV.
The fuel tanks of the attitude control propulsion system were selected based on three criteria; the volume of
the tank, mass of the tank and how many tanks would be required to house the propellants, as shown in Table 3.4.
Table 3.4. Potential ACS propellant tanks [33] [34] [35]
Tank Propellant Volume (L) Mass (kg) Tanks Required
MOOG GEO Sat. Hydrazine 220 27 4
ATK 80505-1 Any 134 16 4
Astrium OST 31/0 MON/MMH 235 16 4
The volume required for the MMH is 84 Liters and the volume required to house the NTO is 82 Liters. With these
constraints in mind ATK’s 80505-1 tank was selected for use. This tank is made of 6AL-4V titanium with a rubber
diaphragm at the mid sphere location.
Figure 3-2. Aerojet R-1E thrusters [30]

30
4. Structural Definition
4.1 Design Approach
According to the mission requirements, Centurion should have payload capacity of 50,000 lbs to LEO and
service lifetime of 5 years or 10 missions. Our vehicle will stay in orbit and dock with manned Orion capsule or cargo
payloads.
The structure of Centurion consists of a systems module, fuel tank, and propulsion module. The systems
module houses various equipment and sensors from ADCS, Communication, Thermal & Power, and Docking
subsystems. It is responsible for controlling the operation of the entire Orbital Transfer Vehicle as well as
communicating with the ground station. The middle section of the OTV is allocated as the fuel tank for both the main
propulsion system and the attitude control propulsion system. Due to the cryogenic nature of liquid hydrogen fuel,
specifically designed thermal shielding is installed around the fuel tank and the structural wall to minimize the effect
of boil-off. The bottom of the OTV is the bimodal thermal nuclear propulsion system utilizing a nuclear reactor and
three thrusters. Most radiators from the Thermal subsystem will also be installed in this section to effectively manage
the thermal performance of the entire OTV.
4.2 Concept Development
Material Selection
The structure of Centurion is divided into three major categories: the truss structure, the fuel tank wall, and
the outer casing. The material selection process for each part are discussed in the following sections.
After comparing the properties listed in Table 4.1, aluminum alloys were considered for the outer casing as
well as inner truss structure of Centurion while composite was considered for the cryogenic fuel tank. Aluminum
honeycomb offers unparalleled stiffness and one of the highest strength-to-weight ratios of any structural core
materials currently available. When treated with chemical conversion coating, the aluminum honeycomb becomes
resistant to corrosion and moisture. It would be the ideal material for the outer casing of Centurion. Aluminum-lithium
alloys, despite its often toxic and dangerous manufacturing process, are great in weight reduction and possess excellent
tensile strength and cryogenic strength. Since the thrusters of Centurion would be burning liquid hydrogen, aluminum-
lithium alloys would be great material for providing overall thermal shielding to the OTV. Propellant tanks have been
traditionally fabricated out of metals. Switching from metallic to composite propellant tank construction dramatically
increases the performance capabilities of the OTV through a significant reduction in weight.

31
Table 4.1. Comparison of common material for space vehicles [36]
Material Advantages Disadvantages
Composites - Low density
- Good strength in tension in appropriate
direction
- Can be tailored for high stiffness,
strength, and low coefficient of thermal
expansion
- Insufficient in compression and tension in
incorrect direction
- Brittle
- Costly to machine in small numbers
- Behaves poorly in environments with high
levels of radiation
Beryllium - High stiffness per density ratio
- Strength close to that of steel
- Alloys of Beryllium are extremely stiff
and lightweight
- Retains its properties up to 1000 degree
Fahrenheit
- Low ductility, fracture toughness, and
impact resistance
- Cannot be primary structural material
- Difficult to fabricate: costly
- Extremely toxic to humans
- Susceptible to surface damage during
machining due to brittleness
Titanium - High strength
- High stiffness to density ratio
- Low weight
- Low coefficient of thermal expansion
- Can replace Al in higher temperature
environments up to 1200 degree
Fahrenheit
- Suitable for cryogenic applications
- Hard and costly to machine
- Low fracture toughness
- Not as light or durable as Al
- Can become brittle at low temperatures or
when placed under repeated loads
- Touching fluids/lubricants can degrade
- Poor resistance to wear
Magnesium - Low density, lighter than Aluminum
- Useful for lower strength, lightweight
applications up to 400 degree Fahrenheit
- Prone to corrosion, needs protective
coatings
- Low yield strength
- Cannot be used in primary structure or areas
subject to wear, abrasion/erosion, or in
contact with moisture
Steel - High strength
- Treatment gives good range of strength,
hardness, and ductility
- High density
- Difficult to machine
- Most alloys are magnetic
Aluminum &
its alloys
- Low density, high strength per weight
ratio
- Easy to manufacture/machine
- If anodized, low surface
absorption/emission properties
- Good in compression
- High coefficient of thermal expansion
- Low hardness
- Cannot be used above 400 degree
Fahrenheit
The PAMG-XR1 5056 Aerospace Grade Aluminum Honeycomb Core from Plascore Inc. were chosen as
the material for the outer casing of Centurion. PAMG-XR1 5056 honeycomb is made from 5056 aluminum alloy
foil and meets all the technical requirements of AMS C7438 Rev A.

32
Figure 4-1. PAMG-XR1 5056 Aluminum honeycomb
Besides the excellent strength, its density of 129.75 𝑘𝑔/𝑚3
would shed significant weight from the outer
casing of Centurion.
The cryogenic fuel tank of the OTV would be constructed from the CYCOM® 5320-1 toughened epoxy resin
prepreg system from Cytek Industries Inc. This epoxy resin system is chosen by NASA because of their high
performance composite cryotank due to its low-cost, lightweight, and superior strength. At 1310𝑘𝑔/𝑚3
, its density is
only a fraction of metallic materials traditionally used for cryotank construction.
Figure 4-2. CYCOM 5320-1 toughened epoxy resin prepreg system [37]

33
Mass Estimation
Mass estimates of all components from all the subsystems were obtained and tabulated. The total dry mass
of the OTV is approximately 15950 kg. The launch mass was estimated by adding the fuel mass as well as the required
payload to LEO. At launch, the fully fueled OTV carrying max payload has a total mass of 88225 kg.
Table 4.2. Mass estimates of Centurion and its components
Subsystem Component Quantity Mass estimate/kg
Structures &
Communication
fuel tank 1 6000
wall + shielding 1 2000
antenna+supporting truss 2 120
ADCS
attitude thruster 4 32
CMG 4 138
CPU 1 3
Proximity Sensor 1 4
star tracker 2 4
attitude thruster fuel tank 4 120
sun sensor 2 2
Thermal & Power
battery 3 120
Radiator + supporting coolant
equipment
65 372
solar panel 2 20
Docking NASA docking system 1 340
Propulsion
main thrusters+nuclear
reactors
3 6675
Total Dry Mass
15950
Fuel Mass
49595
Payload Mass
22680
Total Mass
88225

34
Vehicle Internal Volume
According to the dimensions indicated on Figure 1-2, the outer diameter and total length of Centurion are 7
m and 26.24 m respectively. This ensures that Centurion could fit into the fairing of different launch vehicles, like the
Falcon 9. Due to the amount of liquid hydrogen fuel required for mission to L2, the bulk of the internal volume of the
OTV would be used for cryogenic propellant storage. With propellant tank wall and thermal and radiation shieldings,
the internal volume of the cryogenic propellant tank is 625 𝑚3
, which is more than enough for the entire mission from
LEO to L2.
Vehicle Structural Design
As shown in Figure 1-3, the OTV was divided into three segments, namely systems module, cryogenic
propellant tank, and propulsion system. The structural design of each segment was an iterative process since numerous
revisions were made during the entire project to meet evolving requirements from the other subsystems.
4.2.4.1. Systems Module
Despite the relatively small size of the systems module, it provides structural support to power systems, flight
computers, attitude control thrusters, and various other delicate equipment. At the same time, the systems module will
dock with payload modules from our clients. As a result, the top priority in the structural design process is ensuring
its structural integrity as well as normal operations of the equipment it protects. Also, it is desirable to design the
structure of the systems module with a large margin of safety to prevent possible failures. For Centurion, as shown
in Figure 4-3 below, the inner truss connects the NASA docking system to the cryogenic propellant tank. It also
provides mounting planes as well as structural support to the IMU, CMG, batteries, CPU, sun sensors, star tracker
cameras, radiators, solar panels and the antenna. The top plane and bottom corners of the inner truss are welded onto
the NASA docking system and the main propellant tank respectively. A pair of sun sensors and a pair of star tracker
cameras are bolted onto the top plate of the inner truss, with each pair arranged on opposite sides of the circumference.
Inside the truss, delicate components like the CPU, IMU, and CMG are securely attached to additional mounting
surfaces. On opposite sides of the truss, there are two aluminum truss arms, each of which carries a parabolic reflector
antenna, a hinged radiator, and a solar panel. The truss arm would extend outside the outer casing of the systems
module so that the radiators, solar panels, and antennas could have an unobstructed field of view in space. There are
also two sets of attitude control thrusters in the systems module. Two bipropellant fuel tanks are fixed to the inner
truss while connected to two attitude thrusters via fuel lines.

35
Figure 4-3. Systems Module with outer casing removed
Not shown in Figure 4-3 is the outer casing of systems module. The outer casing consists of two layers, one being
made of aluminum honeycomb for structural support, and the other being an aluminum lithium alloy for thermal
shielding.
In order to reject excess internal heat generated by the onboard equipment, the Alpha Deployable Radiator
(ADP) manufactured by Swales Aerospace was hinged onto the truss arm. ADP is designed to be attached to the
spacecraft through spherical-bearing hinges, pyrotechnic, or paraffin release actuators and snubbers.

36
Figure 4-4. Hinged Radiator for Systems Module [38]
4.2.4.2. Cryogenic Propellant Tank
By using liquid hydrogen as the propellant for the main propulsion system, thermal shielding becomes
equally important as structural support. As a cryogenic fluid, liquid hydrogen must be stored at approximately -423
F° and properly shielded to prevent a phenomenon known as boil-off.
As shown in Figure 4-5, the propellant tank consists of two half spheres and one cylinder. This configuration
is generally preferred to use for pressurized structures. Meanwhile, it also provides a total internal volume of 625 𝑚3
which is the amount of liquid hydrogen fuel required for the mission.

37
Figure 4-5. Cryogenic Propellant Tank
4.2.4.3. Propulsion Systems
The propulsion system is found at the very bottom of the OTV. The bimodal thermal nuclear propulsion
system would be housed and shielded in this segment of the structure. As shown in Figure 4-6, three reactor-thruster
sets are separated, by a layer of radiation shielding, from the upper section of Centurion. This ensures the rest of the
OTV would not be adversely affected by the nuclear radiation from the thermal nuclear reactors.
Figure 4-6. Propulsion System with outer casing removed

38
The extreme heat generated by the reactors and thrusters is another major concern during the design process
of the propulsion system structure. In order to effectively dissipate the heat to space, a deployable radiator system
developed by Lockheed Martin was chosen for this purpose. It uses active, mechanically pumped liquid ammonia
loops to transport heat out to space. With a total surface area of 65 m2
, the pair of double-sided radiators could be
folded for launch and then unfolded into their extended form for final deployment once in space.
Figure 4-7. Deployable radiator [39]
4.2.4.4. Thermal Shielding
Table 4.3. Temperature limits for common materials [40]

39
The highest temperatures affecting structural design typically arise from atmospheric entry or robust
propulsion systems. These conditions require the use of special materials, tailored insulation, or both [41]. The
Space Shuttle uses tile insulation on its exposed aerodynamic surfaces. Most of these areas have normal aluminum
skin-stringer or honeycomb panels beneath, though the most critical locations (e.g., stagnation points) use titanium.
4.2.4.5. Radiation Shielding
Electromagnetic and particle radiation, such as protons and electrons from radiation belts, solar emissions,
and cosmic radiation, can remove structural material. The amount is usually no more than 1 mg/𝑐𝑚2
, which has no
serious effect on the design of most structures. Thin films, however, such as a solar sails, must account for this
degradation. Radiation also reduces the ductility of most materials. This must be anticipated for long-duration or high-
exposure missions since the design life time of the Centurion is expected to be 5 years or 10 missions. Another major
source of radiation is the thermal nuclear reactor onboard. Since relevant shielding will come with the thermal nuclear
propulsion system, we assume the internal radiation to be at the minimal level and would not pose any threat to
onboard equipment as well as payload capsule.
Structural Testing
Besides being the most massive structural component onboard Centurion, the cryogenic propellant tank
serves as the key component connecting the systems module with the main propulsion system. It must be able to
withstand various loads during launch, orbital transfers, and other maneuvers. A brief structural analysis of the
propellant tank was performed in PTC Creo to simulate a simplified model to equivalent loads.
The worst loading on the propellant tank would occur during burn times as the tank was subjected to massive
compressive loads. The fully fueled propellant tank was subjected to internal pressure, top and bottom structural loads,
and gravitational loads estimated at 6 Gs according to the user’s manual from SpaceX [42]. As shown in Figure 4-8,
a maximum stress of approximately 33 MPa was found along the cylindrical segment of the wall. Comparing this
value with the critical value given in the datasheet of CYCOM 5230-1, it is apparent that the propellant tank is
structurally safe.

40
Figure 4-8. Stress Analysis of cryogenic propellant tank
4.3 Critical Design Issues
Since aluminum honey combs will be the primary structure of Centurion, consideration must be made to
ensure they will not fail due to thermal cycling. As Centurion will experience repeated thermal loading and unloading
during its mission to L2, potential damages caused by thermal cycling must be studied and tested to ensure the
structural integrity of its primary structure. Also, because the thermal-nuclear propulsion system is still in
development, its reliability and radiation effects on the overall structural integrity of Centurion is difficult to gauge.
5. Communication and Systems
5.1 Frequency Band Selection
The Near Earth Network (NEN) provides several bands for uplink and downlink including S-band, X-band,
and Ka-band. These bands correspond to a certain operating frequency range. A higher frequency corresponds to
higher data rates, however with this increase comes increases in power requirements and pointing accuracy. Table 6.1
summarizes the bands available for use with the NEN.

41
Table 5.1. NEN Frequency Band Characteristics [2]
Band Frequency (GHz)
S-band 2-3
X-band 7-11
Ka- band 18-30
As Centurion is not transmitting data that would require very high bandwidth, very high frequency bands are
not required. Figure 5-1 depicts the relationship between atmospheric attenuation of a signal and signal frequency.
There is very little attenuation in the lower frequencies that S and X support, however for higher bands such as Ka
there begins to be a much higher level of attenuation. The higher the attenuation of a signal the less reliably the signal
goes through. As Centurion does not require the high speed data transmissions that Ka-band allows, it is much safer
to utilize the two lower frequency bands of S and X. Both of these bands are thoroughly supported on the NEN.
Figure 5-1. Atmospheric attenuation as a function of frequency [43]
5.2 Radiometric Tracking
The NEN provides several services for position and velocity determination that will be useful for Centurion.
These services include Doppler, range, and angle tracking. These services will enable accurate orbit determination
which is vital for orbital transfer maneuvers as well as for docking maneuvers. A spacecraft’s range is measured by
round trip travel time of a sequence of sinusoidal tones originating at one of many different ground-based stations.
The trip time is then divided by the speed of light to calculate the position of the spacecraft. As this signal travels to
and from the spacecraft, its frequency is slightly modified by Doppler shift. Comparing the modified signal to the

42
original signal allows the velocity of the spacecraft to be determined [44]. Angle tracking uses a similar method to
ranging, however, it requires two ground antennas rather than just one. Each satellite calculates range from the
spacecraft by using the distance between the antennas and the angle in the sky [45].
Table 5.2. Near Earth Network Tracking Characteristics [46]
Characteristics Value
Ranging Accuracy 10 Meters (1 sigma)
Doppler Accuracy 1 millimeter per second (1 sigma), 5 second
integration time
Angle Accuracy 0.1 Degrees
Maximum Velocity 2.0 Degrees/second (az and el)
Table 5.2 shows the accuracy capabilities of the NEN. The values that are particularly useful are the ranging
and Doppler accuracies. These values are important for orbital transfers because they allow Centurion to perform
accurate ΔV maneuvers and orbital transfer maneuvers. While the ranging capabilities are not accurate enough for
docking maneuvers they are accurate enough to get in close enough proximity to the fuel depot and the payloads for
the proximity sensors to allow even finer accuracies of range and velocity.
Comparing the NEN(Near Earth Network) with the DSN(Deep Space Network), Figure 5-2 shows that more
vehicles in lunar orbit and L1/l2 orbits use the DSN compared to NEN. This is therefore an opportunity for Centurion
to take advantage of the lack of trraffic in the NEN.
Figure 5-2. Number of missions using NEN vs. DSN [47]

43
NEN can operate optimally at the L1 and L2 ranges, just as well as DSN. G/T is a measure for the antenna gain
accounting for differences in noise measurements at different distances from Earth. G/T is frequently used as a
measure of performance. Figure 5-3 shows that NEN, represented by the red plot, can provide up to 25 dB/Kelvin in
the S-Band. This large range means that NEN can be used to provide communication for vehicles in L1 and L2 orbits
without a significant degradation in the quality of signal.
Figure 5-3. NEN performance compared to DSN using the S-Band [47]
EIRP(Equivalent Isotropically Radiated Power) is another measure of the performance of a communication
network. In Figure 5-4, DSN provides a higher EIRP compared to NEN in the S-band region, which means that DSN
signals are stronger than NEN. However, the graph shows that NEN signals are strong enough to be used at L1 and
L2 orbits, with NEN having a maximum EIRP of about 85 dBm/W.

44
Figure 5-4. EIRP of NEN compared to DSN in the S-Band [47]
NEN has G/T and EIRP values that allow it to be used effectively in communication in vehicles in the L1 and L2
regions, such as Centurion. The largest advantage of using the NEN lies in the smaller number of vehicles currently
using NEN, which provides a larger of communication frequencies that can be used.
5.3 Antenna Selection
As listed in Table 5.3, parabolic reflector is the ideal choice of antenna configuration for Centurion as it
provides the highest gain with reasonable max gain. The high gain antenna is able to transmit data to Earth on two
frequency channels, on at roughly 8.4 gigahertz and the other at around 2.3 gigahertz. The 8.4 GHz channel is the X-
band that sends scientific and engineering data whereas the 2.3 GHz channel is the S-band that relays status of
Centurion to Earth.
Table 5.3. Types of antenna for space communication [11]
Antenna Type Typical Max Gain/ dBI Mass /kg
Parabolic Reflector 15-65 10-30
Helix 5-20 10-15
Horn 5-20 1-2
Array 5-20 20-40

45
Figure 5-5. High Gain Antenna with parabolic reflector
6. Attitude Determination and Control Systems
6.1 Design Approach
The attitude determination and control system (ADCS) capabilities for such a versatile mission must be
comprehensive. The systems of Centurion must be able to perform with high accuracy in low earth orbit (LEO), while
in transit to Lagrange points 1 and 2 (L1 and L2, respectively), while station keeping at those points, and while docking
with the fuel depot and payload module. In addition to performance requirements, these systems must also be power
efficient, cost efficient, and have a lifespan suitable for this mission. To achieve the goals of the ADCS subsystems a
variety of sensors and actuators were considered. Centurion’s attitude sensors must be able to collect accurate attitude
data under any flight condition and at any position along its route. Its actuators must be able to orient the spacecraft
accurately for docking procedures and when the spacecraft is at its peak mass. Furthermore, the control system design

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