1 – Foreword
This report represents the culmination of an intensive spacecraft design course, A&AE
450, undertaken by seniors during a single semester. The students perform a feasibility
study for a specified mission goal, subject to certain constraints.
The entire class works as a single team to achieve this goal. They elect a Project
Manager and an Assistant Project Manager and organize into specialized groups to study
(in this case) attitude control, communications, mission operations, power, propulsion,
and structures and thermal control.
At the end of the semester the students deliver a formal presentation of their results.
Besides this report, the class provides an appendix, which provides detailed analyses of
their methods and trades studies.
The quality of the work in this report is consistent with the high standards of the
aerospace industry. The students who participated in this study have demonstrated that
they have mastered the fundamentals of astronautics, have learned to work efficiently as a
team, and have discovered innovative ways to achieve the goals of this project.
In this particular project, the students were challenged to minimize the absolute cost of
delivering a small payload to the Moon's surface and to minimize the relative cost of
delivering an arbitrary payload. In both cases the payload was required to travel 500
meters across the lunar surface and to transmit images and data back to the Earth. In the
small payload case, the mission corresponds to the Google Lunar X PRIZE. In the case
of the large payload, the idea was to determine the payload mass that provides the lowest
cost per kilogram, i.e., “the biggest bang for the buck,” that would perhaps provide
insight into the size of a delivery system that could support a human outpost on the
Moon. While cost is very difficult to assess (in large part because of the proprietary
nature of the subject), the students managed to give meaningful and reasoned estimates of
the driving economic factors.
I believe this design team rose to the occasion to produce an important feasibility study.
The leadership of the Project Manager and Assistant Project Manager as well as the
outstanding cooperation of the team members were key elements in the success of their
project. They have every right to feel proud of their accomplishment and I am proud of
them.
Professor James M. Longuski
Purdue University
April 6, 2009
Introduction - Background
Section 2.1, Page 4
2 – Introduction
2.1 – Background
The Moon has been a point of interest to humans since the beginning of recorded time. It
has been the subject of many works of art and literature and the inspiration for countless
others. While the Moon has always appeared tantalizingly close, advancements in science
and technology made only in the past fifty years have enabled lunar exploration.
Thrust forward by the space race between the United States and the Soviet Union in the
late 1950’s and 1960’s, robotic exploration of the Moon started off with a bang (or,
rather, a resounding “thud”). Both countries encountered recurring failures during their
early attempts at exploration, and they quickly understood the full complexity of lunar
missions. This sentiment resonated in President John F. Kennedy’s speech at Rice
University on September 12, 1962, stating that “we choose to go to the Moon in this
decade and do the other things, not because they are easy, but because they are hard.”
Kennedy’s ambitious goal of landing a man on the moon before 1970 established the
Apollo Program.
After the Apollo program achieved its primary objective in 1969, the program retired in
1972, initiating a long lull of worldwide lunar exploration. However, recent exploration
initiatives, emerging space programs, and potential lunar resources have placed the Moon
back in the “hot seat” of exploration.
To this day, lunar exploration has remained strictly the domain of large government-run
organizations due to prohibitive mission costs. The Google Lunar X PRIZE (GLXP)
aims to change this fact. A “$30 million international competition to safely land a robot
on the surface of the Moon, travel 500 meters over the lunar surface, and send images and
data back to the Earth,” the GLXP requires that “teams must be at least 90% privately
funded” (Google, 2009). By providing a monetary prize, the X PRIZE Foundation aims
Author: Solomon Westerman
Introduction - Background
Section 2.1, Page 5
to enable commercial access to the Moon, bringing the Moon back to the forefront of
inspiration for individuals and corporations alike.
Author: Solomon Westerman
Introduction – What’s in this Report?
Section 2.2, Page 6
2.2 – What’s in this Report?
Project Xpedition is the product of Purdue University’s Aeronautical and Astronautical
Engineering Department Senior Spacecraft Design class in the spring of 2009. This
report represents countless hours of work by 30 students over the course of a semester.
This report is divided into two sections: a Main Body and an Appendix. The Main Body
consists of a mission overview and important specifications for the team’s vehicles. All
analysis and detail of the solution path are contained in the Appendix. The codes
referenced in this paper are available on the team’s website.
Author: Solomon Westerman
Introduction - Acknowledgements
Section 2.3, Page 7
2.3 – Acknowledgements
The design team would like to thank the following individuals for sharing their time,
expertise, and advice. Their continued support enabled us to create a quality design
fitting of the School of Aeronautics and Astronautics.
AAE450 Instructional Team
James M. Longuski, Professor – Instructor
Alfred Lynam, Graduate Student – Teaching Assistant
Joseph Gangestad, Graduate Student – Project Advisor
We would also like to recognize guest lecturers, industry contacts, and other professors
for their contributions to the project.
Their guidance was invaluable in the design
process.
Guest Lecturers
David Filmer, Professor – Link Budget Analysis
Robert Manning, Graduate Student – Thermal Control
Dr. Boris Yendler, Lockheed Martin – Satellite Thermal Management
Industry and Academic Contacts
Edward Bushway III – Moog Inc.
Daniel Grebow – Graduate Student
Stephen Heister – Professor
Ivana Hrbud – Professor
Martin Ozimek – Graduate Student
Christopher Patterson – Graduate Student
Bruce Pote – Busek Company
Peter Van Beek – Maxon Motors
Author: Solomon Westerman
Introduction - Acknowledgements
Section 2.3, Page 8
Additionally, we would like to thank LunaTrex, an official Google Lunar X PRIZE team
for the insight and guidance they shared with our team. We wish LunaTrex the best of
luck in the competition!
LunaTrex Lecturers
Pete Bitar – Team Leader
Mary Cafasso – Technical Team Leader
Joseph Gangestad – CIO
Author: Solomon Westerman
Introduction – Acronym List
Section 2.4, Page 9
2.4 – Acronym List
This acronym list provides a comprehensive list of acronyms used in this paper in
alphabetical order. We hope this will prove a useful tool for you while reading this
report.
ACS – Attitude Control System
AOL – Active Oxygen Loss
CAD – Computer Aided Design
CCAFS – Cape Canaveral Air Force Station
CMOS – Complementary Metal Oxide Semiconductor
COTS – Commercial Off-The-Shelf
CPU – Central Processing Unit
CuInSe – Copper-indium-selenide
DC/DC – Direct Current/Direct Current
EP – Electric Propulsion
FCS – Flow Control System
FEA – Finite Element Analysis
FEM – Finite Element Model
FS – Factor of Safety
GaAs – Gallium-Arsenide
GaInP2 – Gallium-indium-phosphorus
Ge – Germanium
GLXP – Google Lunar X PRIZE
HD – High Definition
HET – Hall Effect Thruster
HETL – High Energy Tangent Landing
HRE – Hybrid Rocket Engine
LEO – Low Earth Orbit
LL – Lunar Lander
Author: Solomon Westerman
Introduction – Acronym List
Section 2.4, Page 10
LLLE – Lunar Lander Locomotion Engine
LLME – Lunar Lander Main Engine
LLO – Low Lunar Orbit
MLI – Multi-Layer Insulation
MOSFET – Metal Oxide Semiconductor Field Effect Transistor
MS – Margin of Safety
NAND – Inverted AND logic
OTV – Orbital Transfer Vehicle
PAF – Payload Attach Faring
PARI – Pisgah Astronomical Research Institute
PCDU – Power Conditioning and Distribution Unit
PFA – Perfluoroalkoxyethylene
PPU – Power Processing Unit
PTFE – Polytetrafluoroethylene
RFHRE – Radial Flow Hybrid Rocket Engine
RTG – Radio-isotope Thermo-electric Generator
SNAP – System Nuclear Auxiliary Power
SRM – Solid Rocket Motor
TLI – Trans-lunar Injection (also Lunar Transfer)
TOF – Time of Flight
UHF – Ultra High Frequency
VIRAC – Ventspils International Radio Astronomy Centre
Author: Solomon Westerman
Project Overview – Design Requirements
Section 3.1, Page 11
3 – Project Overview
3.1 – Design Requirements
The goal of Project Xpedition is threefold:
Project Xpedition Goals
1) Minimize the cost of a mission satisfying GLXP requirements – while
carrying a ballast of 100g.
2) Minimize the cost of a mission satisfying GLXP requirements – while
carrying a ballast of 10kg.
3) Minimize the cost per kilogram to the lunar surface satisfying slight
modifications to the GLXP requirements discussed below.
The first and second goals of Project Xpedition are similar to the goals of teams currently
participating in the GLXP. The ballast in either of these cases could be representative of
a small scientific or commercial payload, as commercial interest in a low cost payload
delivery service to the lunar surface is promising. Because the mission requirements for
the first and second goals are directly related to the GLXP, these two cases are referred
together as “GLXP-Sized” missions.
The third goal of Project Xpedition is distinctly different from the first and second goals.
The global minimum of cost per kilogram to the lunar surface satisfying slight
modifications to the GLXP rules is more applicable to a lunar base re-supply mission
than the GLXP contest itself. Because this design goal is open-ended on the range of
payload deliverable to the lunar surface, this goal is referred to as the “Large” mission.
We provide a summary of the GLXP requirements below. For a detailed listing of GLXP
requirements, please refer to Section A-9 in the Appendix.
Author: Solomon Westerman
Project Overview – Design Requirements
Section 3.1, Page 12
Base GLXP Requirements
1) Safely land a robotic vehicle on the lunar surface
2) The vehicle or secondary vehicle deployed by the vehicle must move a
distance of 500 meters on the surface of the Moon in a deliberate manner.
3) Transmit an “Arrival Mooncast” and a “Mission Complete Mooncast” to
Earth.
The arrival mooncast and mission complete mooncast are specific sets of data to be sent
back to Earth in order to verify successful completion of the GLXP requirements.
Additional mission success requirements are imposed on each mission.
These
requirements are provided by the instructor and are not part of the GLXP.
Mission Success Requirements
1) All three payload cases should have a 90% probability of success.
2) GLXP-Sized mission requirements must be satisfied by the GLXP deadline on
December 31, 2012.
We approach the probability of success requirement through different methods for the
GLXP-sized and Large payload missions. A cost-effective GLXP-Sized mission was
deemed impractical to satisfy the success rate with a single mission, so we plan multiple
launches to meet the overall success rate requirements.
Selected mission configurations for all three missions lead to flight times of
approximately one year. If a re-supply mission to a lunar base fails, a potential gap of
more than two years’ resources is possible until the next mission can reach the base. For
this reason, the reliability requirement of the third goal is more stringent than the GLXPsized missions: a mission success rate of 90% is required for a single mission.
Author: Solomon Westerman
Project Overview – Design Requirements
Section 3.1, Page 13
In addition to satisfying the base GLXP and mission success requirements, the instructor
placed further requirements on the mission designs:
Additional Requirements
1) The vehicle or any secondary vehicle must survive one lunar day and one
lunar night. After the lunar night, a vehicle must transmit eight minutes of
video to Earth. This optional GLXP bonus requirement is considered to carry
the full bonus monetary prize in addition to the main GLXP prize for the
GLXP sized missions.
2) The small “XPF Payload” described in the GLXP guidelines is ignored in
favor of ballast described in the Project Xpedition Goals.
3) A trajectory correction maneuver of 50 m/s provided by the main propulsion
system is required during trans-lunar injection.
4) The communications system during trans-lunar injection is required to be in
contact 90% of the transfer time.
5) The landing site must be near a historical landmark.
6) If using wheels to locomote on the lunar surface, the vehicle is required to
have the ability to climb a 45-degree incline, on a solid surface.
We make slight modifications to the base GLXP requirements and Additional
Requirements for the large payload case:
Large Payload Requirements
1) Any vehicle used to satisfy the large-payload requirements may “hover” at a
maximum of 100 meters above the lunar surface and translate 500 meters
without touching down first.
2) Any vehicle used to translate 500 meters must do so in over one minute.
Author: Solomon Westerman
Project Overview – Design Requirements
Section 3.1, Page 14
We add these requirements with a lunar re-supply mission in mind. The first requirement
eliminates the requirement of touching down on the surface before locomotion, deemed
unnecessary for lunar re-supply. However, the 500-meter displacement requirement is
still in effect in order to allow for re-supply of a lunar base with small errors in landing
location by the re-supply vehicle. The second requirement reduces the risk of impact to
the lunar base by constraining the average approach speed of the re-supply vehicle.
Author: Solomon Westerman
Project Overview – Interpretation of Design Requirements
Section 3.2, Page 15
3.2 – Interpretation of Design Requirements
Project Xpedition is a feasibility study aimed at providing an initial design concept as a
stepping stone to further development of low cost lunar missions. This limited scope
allows us to avoid obstacles encountered in detailed mission design. We acknowledge,
but ignore the following issues:
1) Politics
Our project involves physics, not politics.
International treaties and export
control are ignored. In reality, red tape from a number of sources will cause a lag
in development that may push mission timelines outside acceptable limits.
2) Launch vehicle piggybacking
We assume any unused capacity in the launch vehicle can be sold back to the
launch vehicle company, as the Project Manager does not anticipate having
sufficient funds in his checking account to purchase the unused capacity. This is
a reasonable assumption, as launch vehicles typically carry multiple, smaller
secondary satellites in addition to a primary satellite. We further assume that our
selected launch vehicle will deliver our payload to a final delivery orbital altitude
of our choosing (within capability).
3) Launch timing
Launch vehicles are typically booked years in advance and have limited flexibility
in launch dates. In our analysis, we assume launch vehicles can launch at any
time from their primary facility.
Author: Solomon Westerman
Project Overview – Design Process
Section 3.3, Page 16
3.3 – Design Process
We strive to maintain transparency in the design process in order to provide insight to our
selected mission configurations. Although we believe our selected mission configuration
is appropriate for our design process, a different design process is not guaranteed to
provide the same results.
The Project Manager and Assistant Project Manager are responsible for the successful
completion of the project. The Project Manager is not assigned a technical group but
instead acts as the primary systems engineer for the team. The Project Manager is
responsible for the structure of the design process as well as the final selection of the
mission configuration, total cost, and risk analysis. The Assistant Project Manager, in
addition to being assigned to a technical group, assists in the structure and management
of the design process.
Our design process was clearly divided into three steps: preliminary design, alternative
design and evaluation, and final design and evaluation. All three payload cases were
analyzed simultaneously.
Preliminary Design – Weeks 1 - 3
The team was initially divided into six disciplines: Attitude, Communications, Mission
Operations, Power, Propulsion, and Structures/Thermal. Each of these technical groups
roughly corresponds to specialty areas within the School of Aeronautics and
Astronautics. Each group member has special interest and extra coursework in their
group. Technical groups contain between four and seven members including a group
leader who was responsible for attending extra meetings and delegating work
appropriately within the group.
Design is inherently an iterative process that requires a seed to grow. The team began the
design with this seed – X0. This seed described a conventional lunar mission, based on
Author: Solomon Westerman
Project Overview – Design Process
Section 3.3, Page 17
proven, historical methods to deliver payload to the lunar surface. While not optimized
and certainly not cost-effective, the X0 design provided a baseline for comparison to all
alternatives suggested later in the design process. This allowed each group to break
down systems on X0 and attempt to improve upon it, as detailed in the alternative design
and evaluation process.
Alternative Design and Evaluation – Weeks 3 - 7
In the third week, each team member was assigned a “phase group” in addition to a
technical group. Ignoring the Earth launch, three phases completely describe our lunar
mission: lunar transfer, lunar descent, and locomotion. Phase groups consist of one or
more members from each technical group and proved to be more effective and agile than
technical groups when evaluating alternative mission architectures. Phase groups avoid
some of the iterative nature of design: each phase group is relatively independent of each
other and usually only needed high level parameters from other phases to complete
analysis. As an additional benefit, phase groups were not locked in to a specific vehicle
type to complete a phase and can work with other phase groups to develop an efficient
architecture.
The primary action of the alternative design process was to develop and analyze
alternative methods within each phase. If an alternative was clearly superior to the X0
design, that alternative replaced the old design and became the new baseline. Each
alternative design was taken to a different level of maturity depending on its level of
merit within the mission as a whole.
At the end of the seventh week, the architectures for lunar transfer and lunar descent were
firm, with clear winners chosen from alternatives. Locomotion, however, went through a
rigorous process involving trades between mass, power, cost, and reliability to ensure the
lowest overall mission cost. The selection of the mission configuration for all phases at
Author: Solomon Westerman
Project Overview – Design Process
Section 3.3, Page 18
the end of the seventh week allowed two weeks to finish final analysis on the selected
mission configuration for each payload case.
Final Design and Evaluation – Weeks 8 – 9
The purpose of final design and evaluation was to refine the models used in the
alternative design and evaluation process to provide reasonable accuracy to the selected
mission configuration. Changing mission configurations in this step was only done under
extenuating circumstances. The team formed “integration managers” to handle the fast
turnaround on numbers between phase groups during this process. Three integration
managers, each assigned to a payload case, managed the mass, power, volume, and cost
budgets for their payload case. By the end of the ninth week, finalization of the mission
configuration for each payload case was complete.
Author: Solomon Westerman
Project Overview – Risk and Cost Analysis
Section 3.4, Page 19
3.4 - Risk and Cost Analysis
The concepts of risk and cost are interrelated and are thus combined in this section of the
paper. While true costs remain elusive and a detailed risk analysis is outside the scope of
this project, we believe the models used in this project are valid for a feasibility study.
The purpose of this section is to highlight the methodologies used in risk and cost
analysis in order to give further insight into our selected mission configurations.
Mission risk and cost are broken down by vehicle. Costs for each vehicle are broken into
the following categories:
1) Purchase – the cost of purchasing a component from another company.
2) Integration – the cost associated with manufacturing the vehicle.
3) Research and Development – the cost of developing or refining a technology inhouse.
Overhead is a category of cost that we include which is not associated with a particular
vehicle. Overhead represents the recurring cost of maintaining a company. This cost
includes salary for engineers working on the software and design levels, software license
costs, and communication rental expenses.
We assume that all purchases are made in 2009. Integration, research and development
costs are spread out evenly over three years and overhead over four years. We purchase
the launch vehicle in 2011, the same year as launch. All costs include a flat 3.8%
inflation rate used to represent all past and future cash flows in terms of 2009 dollars.
Our costing method tends to underestimate the total mission costs. Although an accurate
estimate of total mission cost requires more analysis, this does not affect the technical
feasibility of this project.
Author: Solomon Westerman
Project Overview – Risk and Cost Analysis
Section 3.4, Page 20
Risk analysis is also broken down by vehicle. The risk analysis technique used for the
Earth launch phase is distinctly different from the technique used for the other three
phases. Estimating the success rate of a launch vehicle is not trivial – the problem is
complicated further if the vehicle does not have extensive heritage. In our analysis, we
assume the launch vehicle probability of success is a function of historical reliability data
and the number of stages in the launch vehicle (a measure of complexity of the system.)
In the GLXP-sized payload cases we selected the Dnepr launch vehicle, a modified
version of an old intercontinental ballistic missile. The success rate of the launch vehicle
depends on the number of stages on the launch vehicle and historical launch data. For the
large-sized payload case, the selected launch vehicle has yet to be flown. The reliability
of this launch vehicle is assumed to be similar to the success rate of similar-sized launch
vehicles from the same country of origin.
Risk analysis for each vehicle that we design follows a different methodology from
launch vehicle risk analysis. We identify major failure modes for the vehicle and assign
an appropriate probability of success for each failure mode. Moving parts and systems
developed in-house are typically assigned a lower base reliability than non-moving or
purchased systems. If necessary, we augment this base reliability with money from
research and development to a maximum reliability cap. The probability of success for
each of the failure modes are multiplied together to calculate the success rate of each
vehicle. Each vehicle’s probability of success is then multiplied together to produce the
overall mission success rate.
In some cases, the reliability of a particular system is too low to satisfy the mission
requirements, even with research and development.
Our approach to satisfy the
requirements is to add a redundant system to the vehicle. This greatly increases the
success rate of the particular sub-system, as we assume the cause of a failure in one
system will not reduce the redundant system’s success rate. This assumption leads to
Author: Solomon Westerman
Project Overview – Risk and Cost Analysis
Section 3.4, Page 21
optimistic success rates as the cause of failure of one system is potentially coupled with
the redundant system.
Author: Solomon Westerman
Project Overview – Results
Section 4, Page 22
4 – Results
Our feasibility study found that it is practical for a commercial company (with a limited
funding base) to deliver small payloads to the lunar surface. However, a commercial
company is not likely to profit directly from the Google Lunar X PRIZE purse money.
Table 4-1 summarizes total mission costs (with the arbitrary payload case included as a
comparison to the GLXP-sized missions).
Table 4-1 Mission Results
Mission
GLXP
GLXP
Large
Payload
100 g
10 kg
1743 kg
Total Cost
$27.1M
$29.6M
$222.6M
Cost/Payload
$271M/kg
$3.0M/kg
$128k/kg
Profit
$(4.8M)
$(7.3M)
---
Footnotes:
Payload – Payload mass deliverable to lunar surface
Total Cost – Mission cost, 2009 dollars
Cost/Payload – Cost per kilogram of payload
Profit – Net cash flow with GLXP purse
The most cost-effective mission configuration is relatively independent of payload
delivered to the surface. In order to minimize total mission cost, a company should use a
solar-electric propulsion system while the spacecraft travels from the Earth to the Moon.
In order to deliver payloads from lunar orbit to the lunar surface, a conventional, softlanding vehicle powered by a single hybrid engine is a cost-effective solution. Our
analysis indicates that this mission configuration results in low mission costs for payloads
ranging from 1 gram to approximately two metric tons. Although our analysis did not
converge on a minimum mission cost for payloads larger than two metric tons, the team
believes this mission configuration is an excellent springboard for further analysis of
large payload cases.
Author: Solomon Westerman
Project Overview – Results
Section 4, Page 23
Our analysis indicates four feasible methods to travel short distances on the lunar surface
for a large range of payload sizes. Our cost model indicates the technique of locomotion
has a relatively low effect on mission cost. While we are confident in our selected
methods of locomotion, a different method of selection may lead to a different type of
locomotion. We split our locomotion configuration results into two categories: missions
satisfying the GLXP requirements and missions satisfying the modified GLXP
requirements as outlined in the Design Requirements section.
If a company is attempting to satisfy the modified GLXP requirements, the most costeffective design for translating across the surface to deliver payload to a lunar base is to
throttle the main landing engine to a low thrust level and translate above the lunar surface
after the main descent burn is completed. This method is an inexpensive and reliable
technique to displace short distances on the lunar surface. Our analysis indicates this
“hover” method is the global minimum in total mission cost for all payload cases.
However, this method is only incorporated in the arbitrary payload mission, as this
mission is not required to satisfy the exact GLXP requirements.
In order to satisfy the GLXP requirements and win the purse money, a different method
of locomotion is ideal and is dependent on the ballast mass prescribed in the mission
requirements. For a GLXP-sized mission with 100g ballast, the optimal method of
locomotion across the lunar surface is a small spherical ball with a movable, offset center
of mass.
We use two identical spherical balls to maintain an acceptable level of
reliability. This type of locomotion is not ideal for the larger ballast case of 10kg, where
a separate chemical propulsion system becomes the preferred method of locomotion
across the lunar surface. We use a single hybrid engine with a separate combustion
chamber from the main landing engine in order to satisfy the GLXP locomotion
requirements. Like the 100g case, two independent systems of locomotion are required
to maintain an acceptable level of reliability.
Author: Solomon Westerman
Mission Configuration – 100g Payload
Section 5, Page 24
5 – Mission Configuration 100g Payload
Author: Solomon Westerman
Mission Configuration – 100g Payload – System Overview
Section 5.1, Page 25
5.1 – System Overview
The mission configuration for the 100g payload consists of the orbital transfer vehicle
(OTV), Lunar Lander, and two redundant Space Balls that move across the lunar surface.
We design the system to move 500m across the lunar surface, survive the lunar night,
take and send pictures and video, and visit a heritage site on the moon. Figure 5.1-1
outlines the configuration of the entire system, and Fig. 5.1-2 shows the relative sizes
between the Space Balls and the Lunar Lander.
Fig. 5.1-1 View of the entire system. The OTV is bottom portion
and the Lander is the upper portion.
(Korey LeMond)
Author: Brittany Waletzko
Mission Configuration – 100g Payload – System Overview
Section 5.1, Page 26
Fig. 5.1-2 Space Balls and Lunar Lander on the Moon.
(Ryan Lehto)
Table 5.1-1 summarizes the mission timeline for the 100g payload. The landing site is in
Mare Cognitum, near the Apollo 12 landing site.
Author: Brittany Waletzko
Mission Configuration – 100g Payload – System Overview
Section 5.1, Page 27
Table 5.1-1 Mission Timeline
Elapsed Time
(ddd:hh:mm)
-365:00:00
000:00:00
000:00:03
000:00:04
000:00:04
000:00:05
000:00:06
000:00:06
000:00:14
Event
Vehicle
Launch
Arrive in LLO
In lower orbit
Rotate and Land
Systems check
Deployment from Lander
Orientation
Travel 500m
Braking maneuver, dust removal
Take picture of Lander, Begin
000:00:15
transmission to Lander
000:00:23
End photo transmission
Transmit arrival Mooncast (near
real-time video, photos, HD video,
000:00:23
XPF set asides, data uplink set) to
Earth
Transmit Mission Complete
001:33:56
Mooncast (near real time video,
photos, HD video)
Finished transmitting, prepare for
002:08:04
night
009:00:00
Standby for lunar night
025:00:00
Power up after night
026:00:00
Transmit telemetry and photo
026:00:14
Mission Complete
Elapsed Time given in days, hours, and minutes
Launch Vehicle/OTV
OTV
Lander
Lander
Space Ball
Space Ball
Space Ball
Space Ball
Space Ball
Space Ball
Space Ball
Lander
Lander
Lander
Lander
Lander
Lander
Table 5.1-2 summarizes system masses at key mission checkpoints.
Table 5.1-2 100g Payload System Masses
Parameter
Injected Mass to Low Earth Orbit
Injected Mass to Low Lunar Orbit
Mass on Lunar Surface
Payload Delivered
Mass
436
156
79
100
Units
kg
kg
kg
g
The following sections present greater detail on each vehicle subsystem.
Author: Brittany Waletzko
Mission Configuration – 100g Payload – Launch Vehicle
Section 5.2.1, Page 28
5.2 – Launch Vehicle
5.2.1 – Launch Vehicle Selection
We select the Dnepr-1 rocket to place the 100 g payload into a Low Earth Orbit (LEO).
The rocket is a converted intercontinental ballistic missile operated by the International
Space Company Kosmotras (a joint venture among Russia, Ukraine, and Kazakhstan).
The Dnepr-1 was chosen based on its low launch cost and adequate reliability. The
Dnepr-1 employs previous technologies and stages from its missile counterpart which
contributes to its low launch cost. To date, the Dnepr-1 has eleven successful missions
and only one failure. Figure 5.2.1-1 shows the launch profile of the Denpr-1.
Fig. 5.2.1-1 Dnepr-1 Launch Profile.
(Space Launch System Dnepr User’s Guide)
Author: Zarinah Blockton
Mission Configuration – 100g Payload – Launch Vehicle
Section 5.2.1, Page 29
Table 5.2.1-1 Characteristics of the Dnepr-1 Launch Vehicle
Variable
Height
Diameter
Lift-off Mass
Mass to 400km orbit
Number of Stages
Orbit Inclination
Cost per kilogram to LEO
Value
34.3
3
208,900
3,400
3
50.5
4,800
Units
m
m
kg
kg
-degrees
$/kg
The Dnepr-1 operates from one of the largest spaceports in the world, Baikonur
Cosmodrome in Kazakhstan. The Baikonur Cosmodrome is located a sufficient distance
from densely populated areas; this will ensure safety during stage drops. Although
launching from a site this far north of the equator decreases launch vehicle capability, the
launch price and performance in LEO still make the Dnepr-1 the most attractive option
for our mission configuration. Figure 5.2.1-2 shows the location of our launch site.
Fig. 5.2.1-2 Map of Launch Site.
(Google Maps)
Author: Zarinah Blockton
Mission Configuration – 100g Payload – Launch Vehicle
Section 5.2.2, Page 30
5.2.2 –Earth Parking Orbit Selection
Atmospheric Drag
We display the effects of atmospheric drag from altitudes of 200 km to 500 km in Table
5.2.2-1. At 200 km the drag force is still quite significant, and would cause the parking
orbit to decay quickly. Based on the data in Table 5.2.2-1, a parking orbit of 400 km is
the optimal drag solution. At 400 km the drag force is less than 5% of the thrust (thrustto-drag is greater than 20).
The industry standard for minimum thrust-to-drag ratio in
Earth orbit is approximately 10. A higher orbit will only further increase our launch costs,
which are the largest mission cost driver.
The 400 km altitude is also within the
capabilities of our selected launch vehicle for the 100g payload, the Dnepr.
Table 5.2.2 - 1 Atmospheric Drag Effects at Various Altitudes
:
Altitude (km)
200
300
400
500
Drag (mN)
76.4
17.8
4.10
0.96
Author: Andrew Damon
Thrust/Drag
1.44
6.18
26.83
114.6
Mission Configuration – 100g Payload – Launch Vehicle
Section 5.2.2, Page 31
Cost Comparison for Parking Orbit
The altitude of the parking orbit affects both launch and lunar transfer costs. Higher
parking orbits reduce the cost of lunar transfer by decreasing mass and power
requirements for the OTV; however, the launch cost increases because it is a higher
energy orbit. We find that launch cost has the greatest impact on overall mission cost, so
we select a parking orbit that minimizes launch cost.
We see in Fig. 5.2.2-1 that
performing translunar injection from the lowest parking orbit possible results in
minimum cost.
Launch and Lunar Transfer Cost(Million $)
18
16
14
12
10
8
6
4
400
500
600
700
800
900
1000
Parking Orbit Altitude (km)
1100
1200
Fig. 5.2.2-1 Mission Cost with Varying Parking Orbit Altitude.
(Levi Brown)
Because we must launch into a minimum 400 km orbit to overcome drag effects, we
select a 400 km circular parking orbit.
Author: Levi Brown
Mission Configuration – 100g Payload – Launch Vehicle
Section 5.2.3, Page 32
5.2.3 - Attitude Determination in LEO
To maintain attitude control in low Earth orbit (LEO), we employ both a sun sensor and
star sensor. When the payload faring detaches from the launch vehicle, these sensors
come online and determine the attitude of the orbit transfer vehicle. The sun sensor is a
product of Valley Forge Composites and is accurate to approximately one degree. Sun
sensors are designed for attitude determination during de-spin maneuvers, such as the
OTV leaving the launch vehicle, as well as for directing solar arrays while in transit to
the Moon. The star sensor, used in parallel with the sun sensor, has much higher
accuracy and is used for attitude determination throughout the mission. This particular
sensor performs to an accuracy of 15 arcseconds (one fourth of a degree) by identifying
the location of stars instead of the sun. The combination of these subsystems can satisfy
our requirements for determining and monitoring attitude while the OTV remains in
LEO.
These systems have very high performance and use only a modest amount of spacecraft
resources.
The sun sensor we selected has a mass of 0.35 kg and at peak power
consumption uses only 2.5 Watts. The star sensor is slightly more robust with a mass just
less than 3.2 kg, volume of 16,400 cubic centimeters, and a maximum power
consumption of 10.2 Watts. These units together have a total mass of approximately 3.55
kg and determine the attitude for the entire mission (Valley Forge, 2009). Furthermore,
the sensors can be purchased from Valley Forge Composite Technologies in a package
including reaction wheels for approximately $400,000.
We must maintain continuous contact with ground stations after the OTV deploys; an
accuracy in attitude determination of approximately one degree ensures we can
accomplish this goal. This constraint that a ground station must be in view at all times
would place a hefty burden on attitude determination and attitude control except that we
employ a moveable antenna.
For this reason, we assume that the upper bound on
required accuracy (the higher the accuracy, the lower the value in degrees) is about half a
Author: Kristopher Ezra
Mission Configuration – 100g Payload – Launch Vehicle
Section 5.2.3, Page 33
degree. This accuracy is definitely achievable because the star sensor alone has an
accuracy of one fourth of a degree. In typical cases, it is reasonable to think that in LEO
even the sun sensor could be accurate enough to place a ground station in view of the
antenna.
Unlike other subsystems onboard the OTV, the attitude determination systems do not
scale or vary with changing mission requirements. As we will discuss later in this
document (Sections 6.2.3 and 7.2.3), there is no change to the attitude determination
methods regardless of new payload masses, new landing/locomotion equipment, or new
mission architectures. The star sensor and the sun sensor will be employed without
significant changes (excluding perhaps position within the OTV) in each of the three
payload cases.
Author: Kristopher Ezra
Mission Configuration – 100g Payload – Lunar Transfer
Section 5.3.1, Page 34
5.3- Lunar Transfer
5.3.1 – Structure
Limit Loads
Launch loads or limit loads from different phases of the launch are obtained from the
Dnepr User’s Manual. We choose these loads to evaluate the capability of our spacecraft
design. Table 5.3.1-1 shows these launch loads for the selected launch vehicle. Loads
are expressed in terms of gravitational acceleration (g’s), and produce forces in all
structural elements of our spacecraft. These forces are broken into axial and lateral
components throughout launch. We select the highest levels of acceleration to apply to
our spacecraft design (axial and lateral simultaneously). We base our entire structural
analysis on these selected loads. This assumption is valid because the most severe
dynamic and static loads are expected to occur through the launch phase of the mission
for our Orbital Transfer Vehicle (OTV).
Table 5.3.1-1 Dnepr launch limit loads
Axial
Acceleration
Lateral
Acceleration
1st Stage Burn: Maximum Lateral Acceleration
3.0 ± 0.5
0.5 ± 0.5
2nd Stage Burn: Maximum Longitudinal Acceleration
7.8 ± 0.5
0.2
Event
Table based from Dnepr User’s Guide
The loads specified are valid only if the spacecraft meets certain stiffness, or natural
frequency requirements. These requirements are summarized in Table 5.3.1-2 for the
Dnepr launch vehicle. These requirements are based on the launch vehicle’s dynamic
characteristics and are used to prevent an amplification of the loading.
Table 5.3.1-2 Dnepr spacecraft stiffness requirements
Thrust (Hz)
20
Lateral (Hz)
10
Table based from Denpr User’s Guide
Author: Tim Rebold
Mission Configuration – 100g Payload – Lunar Transfer
Section 5.3.1, Page 35
Margin of Safety
We apply a factor of safety (FS) when we conduct our analysis. This is done to ensure
that uncertainty will not make our analyses invalid. Uncertainty can come in the form of
imperfections in the structure, material property variations (strength), manufacturing
tolerance errors in structural members, errors in predicting loads in the mission, and other
factors. The yield margin of safety (MS) which we use in all of our analyses is shown as:
𝑀𝑆 =
𝐴𝑙𝑙𝑜𝑤𝑎𝑏𝑙𝑒 𝑦𝑖𝑒𝑙𝑑 𝑙𝑜𝑎𝑑 (𝑜𝑟 𝑠𝑡𝑟𝑒𝑠𝑠)
−1
𝐷𝑒𝑠𝑖𝑔𝑛 𝑦𝑖𝑒𝑙𝑑 𝑙𝑜𝑎𝑑 (𝑜𝑟 𝑠𝑡𝑟𝑒𝑠𝑠)
(5.3.1-1)
The design yield load takes the predicted loads or stresses and multiplies it by the FS
chosen for the mission. If the margin of safety is negative it means that we have
exceeded the allowable loading or stress of the material or member. This means we must
redesign or find a way to lower the expected loads. We use a FS of 1.5 for our
preliminary analyses to be conservative.
Configuration
Our OTV is 1.29 m tall including the Lunar Lander (LL) integration skirt, and has a 1.8
m diameter. Figure 5.3.1-1 displays the primary components and dimensions of the OTV
in a schematic. The OTV is cylindrical with its payload (Lunar Lander-cone shaped)
resting on top. An integration skirt joins the two. We select the overall dimensions of the
OTV based on sizing constraints. Once we choose the basic layout, we can select the
primary structural components that will make our spacecraft. In the following sections
we describe the primary structural components of our spacecraft.
Stiffeners
We use stiffeners as the primary support structure for the spacecraft. The stiffeners are
also referred to as stringers or C-Channels (due to their cross section shape). We choose
stiffeners to provide the main load path for all the forces seen by the spacecraft
throughout launch.
The stiffeners are capable of withstanding the bending and
Author: Tim Rebold
Mission Configuration – 100g Payload – Lunar Transfer
Section 5.3.1, Page 36
compressive forces transferred throughout the entire structure. Six stiffeners provide
enough stiffness and strength as well as distribute the loads uniformly into the launch
vehicle interface at the aft end of the spacecraft
Fig. 5.3.1-1 OTV configuration and design showing primary components and dimensions.
(Tim Rebold)
Propulsion Module
The propulsion system of the OTV is supported with a truss frame structure. The
propulsion system consists of the Xenon propellant tank, feed system, Hall thruster, and
thermal protection equipment. Section A-5.3.1 provides more details on how we chose a
layout and sized the frame.
Four members are the fewest needed to support our
propulsion system.
Electronics Module
The electronics module (E-MOD) houses all the electronics needed for the mission. The
structural support for the E-MOD consists of a floor support beam design and skin which
Author: Tim Rebold
Mission Configuration – 100g Payload – Lunar Transfer
Section 5.3.1, Page 37
covers the entire assembly. The skin serves as a barrier between the Sun’s radiation and
equipment. The skin also offers micrometeorite protection during Lunar Transfer. We
also choose the skin overlay on the floor to act as a radiator to conduct unwanted heat
away from the electronics. The amount of space needed to fit all of the equipment
determines the allowable dimensions. Since the floor skin overlay is so thin (0.5 mm), it
is not expected that they will carry any loads. All equipment is mounted above the six
beam supports. The geometry and layout of this design is shown in Section A-5.3.1.
Skin
We place 0.5 mm skin (shear paneling) around the entire OTV to mainly provide
shielding from the Sun’s radiation and to offer micrometeorite protection. We choose the
thickness to be as thin as possible to reduce weight yet provide adequate support for the
above reasons mentioned. From a structural standpoint, the skin will offer substantial
torsional rigidity for the entire OTV, and handle most of the shear loading seen in the
OTV.
Integration
Integration structure includes the Payload Attach Fitting (PAF) and LL skirt. The latter
joins the OTV to the LL, while the former joins the OTV to the launch vehicle inside the
fairing. There are six locations on both the base and top of the LL skirt where the OTV
and LL are bolted into. These six locations correspond to the location of each stiffener.
Similarly, the PAF attaches the OTV to the launch vehicle inside the fairing. The Dnepr
User’s Guide shows a standard launch vehicle interface. This interface determines the
base dimensions of our PAF. Geometry can be seen in Section A-5.3.1.
Sizing
We size structural members in the OTV by predicting the expected loads transferred
through the spacecraft. We reduce the weight of the spacecraft by minimizing the cross
section dimensions of each member in the OTV. Modes of failure need to be well
Author: Tim Rebold
Mission Configuration – 100g Payload – Lunar Transfer
Section 5.3.1, Page 38
understood to identify the most likely cause for failure in a given member. The most
likely cause for failure will determine the criteria for our design. We can apply basic
techniques and equations to optimize the capability of each member, while making it as
light as possible. The most common failure modes that are used to determine the size of
structural members in our OTV design are discussed in detail in Section A-5.3.1.
Author: Tim Rebold
Mission Configuration – 100g Payload – Lunar Transfer
Section 5.3.1, Page 39
Material Selection
The OTV is generally comprised of three different materials. All structural frame
members are fabricated from aluminum due to the fact that they must be designed to
survive the launch loads that are applied to the vehicle. Aluminum has a higher yield
strength, and all structural frame members are therefore stronger. The shear paneling that
encompasses the vehicle was fabricated out of a lower strength magnesium alloyed with
small concentrations of aluminum and Zinc. The yield strength of the magnesium was
approximately two thirds of that of aluminum, but the density was also two thirds the
density of aluminum. Since shear panels in general carry little in the way of loads, instead
acting as a pathway for shear flow, we chose a lower strength lower density material in
the form of magnesium. Magnesium also is used as the material for the floor of the
electronics module and the floor of the OTV itself. The floor panel of the OTV carries
little in the way of loads due to the fact that the propulsion system is attached to the truss
frames. However, some form of meteorite protection is necessary, and magnesium
suffices. The third material present in the OTV is that of AIS 1015 low carbon steel. This
material is used in all pins, fasteners, and other connection devices, as these devices are
the only load path between structures and therefore see all the launch loads at a given
time.
Table 5.3.1-3 Material Properties
Material
Density
(kg/m3)
Al 6065
2700
Mg AZ31
1770
AIS 1015 Low 7800
Carbon Steel
Yield
Strength (Pa)
1.03E8
1.5E8
2.55E8
Author: Korey Lemond
Young’s
Modulus (Pa)
6.8E10
4.4E10
2.05E11
Mission Configuration – 100g Payload – Lunar Transfer
Section 5.3.1, Page 40
FEA Analysis
As the configuration and sizing of the Orbital Transfer Vehicle (OTV) reaches maturity
we perform a detailed structural analysis of our model.
Fig. 5.3.1-2 Finite element model (FEM) of the OTV.
(Tim Rebold)
The software we use to perform our finite element analysis is IDEAS 12. We use this
package to perform a basic static load analysis to determine critical stresses and
displacements in structural members. We also perform a modal analysis to determine if
our spacecraft meets the stiffness requirements provided in the Dnepr User’s Guide.
Figure 5.3.1-2 shows a complete finite element model (FEM) of our OTV spacecraft
Lander Skirt Analysis
The first analysis we perform is on the Lunar Lander (LL) skirt. By performing a basic
static test we can see the capability of the skirt. Figure 5.3.1-3 shows the LL skirt
geometry.
Author: Tim Rebold
Mission Configuration – 100g Payload – Lunar Transfer
Section 5.3.1, Page 41
Fig. 5.3.1-3 Lander skirt geometry.
(Tim Rebold)
Results
After performing the analysis we find that stress is not a concern. By picking the
locations of where the LL was integrated to the OTV we effectively limited stress and
displacement in the skirt. We find that a buckling analysis determines how thin we can
make our skirt walls (or webs). We settle on a final wall thickness by assuring ourselves
that the skirt does not buckle under the predicted loading.
OTV Analysis
We analyze the entire OTV next. We perform a system by system analysis to simplify
our results. We conduct a basic static test, and modal analysis to find the natural
frequencies of our spacecraft while mounted to the launch vehicle interface.
Results
The first system we observe is the propulsion module frame. The max Von Mises stress
is 324 N/mm2 and max displacement is 1.36 cm. This corresponds to a margin of safety
of -0.17. The negative margin of safety means that yielding occurs in that area of peak
stress. Material yield is unacceptable by our mission requirements. The member failing
is not seeing unreasonably high stresses, so a slight change in design should eliminate the
problem. Figure 5.3.1-4 shows the results for the propulsion frame. The margins of
safety for all other systems are passing.
Author: Tim Rebold
Mission Configuration – 100g Payload – Lunar Transfer
Section 5.3.1, Page 42
Fig. 5.3.1-4 Propulsion frame stress contour results indicate material failure.
(Tim Rebold)
We next perform a linear buckling analysis. The buckling load factor from this analysis
is 0.19, which means that a buckling failure occurs at the current level of applied loads.
We observe that a stiffener is first to buckle from the applied loads. We need to redesign
so this does not occur. We last perform a modal analysis. We observe that the first
lateral mode is well below the requirement. This means that the OTV needs to be stiffer
in the lateral direction.
Design Modifications
Next, we make modifications to our spacecraft until all requirements are satisfied.
OTV Modified Analysis Results
Again, the first system we observe is the propulsion module frame. The max Von Mises
stress is 80 N/mm2 and max displacement is 2.14 mm. This corresponds to a margin of
Author: Tim Rebold
Mission Configuration – 100g Payload – Lunar Transfer
Section 5.3.1, Page 43
safety of 2.14. The positive margin of safety means that this system meets mission
requirements. A buckling load factor of 1.42 and lateral mode of 10.6 Hz are also
determined from the buckling and modal analyses. Therefore, our OTV design meets all
structural requirements. Table 5.3.1-4 shows the structural component masses as a result
of these analyses.
Table 5.3.1-4 Component mass totals for OTV design
Mass
Components
(kg)
PAF*
47.04
E-MOD floor beams & overlay
5.25
Shear / Skin Panels
15
Propulsion Support Frame
3.01
Stringers / Stiffeners
12.48
Lander Skirt
12.19
Fasteners (welds, rivets, bolts, adhesives) ** 2.01
TOTAL
49.94
* Not included in final OTV mass
** Estimate
Author: Tim Rebold
Mission Configuration – 100g Payload – Lunar Transfer
Section 5.3.2, Page 44
5.3.2 – Mission Operations
Trajectory
Per the analysis described in Section A-5.3.2, we design the trajectory for the 100 g
payload to have the configuration detailed in Table 5.3.2-1.
Table 5.3.2-1 100 g Payload Trajectory Configuration
Parameter
Payload Mass
Thrust
Mass Flow Rate
Earth Phase Angle
Moon Phase Angle
Parking Orbit Altitude
Capture Orbit Altitude
Initial Mass
Flight Time
Value
156.2
78.5
4.1
108
288
400
25
436.0
365
Unit
kg
mN
mg/s
deg
deg
kg
kg
kg
days
This model inherently contains a mismatch in position and velocity at the intersection of
the spiral out and spiral in curves (See Fig. 5.3.2-1). We calculate the propellant mass
required to produce the ΔV mismatch operating the main engine. We add this propellant
mass to the OTV to account for the error in position and velocity. We assume that
performing small maneuvers throughout the trajectory eliminates the fairly large bias in
position and velocity at the intersection point. Table 5.3.2-2 contains these mismatch
values.
Table 5.3.2-2 100 g Payload Trajectory Configuration
Parameter
Position
Velocity
Propellant Mass
Mismatch
5676
417.6
9.4
Figure 5.3.2-1 illustrates the resultant trajectory.
Author: Levi Brown
Unit
kg
km/s
kg
Mission Configuration – 100g Payload – Lunar Transfer
5
Section 5.3.2, Page 45
Spiral Out and In of Spacecraft
x 10
2
Spiral Out from Earth
Spiral In to Moon
1.5
1
y
hat
(km)
0.5
0
-0.5
-1
-1.5
F -2
i
g-2.5
.
-2
-1
0
1
x hat(km)
5
.
Fig. 5. 3.2-1 100 g Payload Trajectory.
(Levi Brown)
Author: Levi Brown
2
3
4
5
x 10
Mission Configuration – 100g Payload – Lunar Transfer
Section 5.3.2, Page 46
Lunar Capture Orbit
The lunar capture altitude is 25 km, which optimizes mass savings while maintaining
orbital stability. A small altitude produces the most efficient lunar descent trajectory and
propellant mass savings. However, the lunar terrain and the instability of these smaller
orbits restrict the altitude to the minimum value of 25 km. Once we enter this orbit the
Lunar Lander separates from the OTV and begins its descent to the Moon’s surface.
Author: Kara Akgulian
Mission Configuration – 100g Payload – Lunar Transfer
Section 5.3.2, Page 47
Trajectory Correction Maneuver
One requirement for this project includes the capability to perform a 50 m/s burn for
course correction.
Due to factors beyond the scope of this project, such as Sun
perturbations, the spacecraft will deviate from our trajectory design during actual flight.
To account for this bias, we carry additional propellant for the main engine.
The
propellant is available for making small corrections in flight as necessary.
We calculate the propellant necessary for a 50 m/s burn and record the results in Table
5.3.2-3.
Table 5.3.2-3 Correction Maneuver Configuration
Parameter
Isp
mo
Propellant for Correction
Value
1952
436.0
1.1
Author: Levi Brown
Unit
sec
kg
kg
Mission Configuration – 100g Payload – Lunar Transfer
Section 5.3.3, Page 48
5.3.3 – Propulsion on the Orbital Transfer Vehicle
Propulsion Subsystem Overview
We select a low-power Hall Effect Thruster (HET) for primary thrust onboard the Orbital
Transfer Vehicle. HETs are a type of electrostatic propulsion (EP), and there are two
defining features to keep in mind with these devices: They have very low thrust (tens of
milliNewtons), and very high specific impulse (thousands of seconds). Our HET
propulsion system contains four critical components, including the Power Processing
Unit (PPU), the propellant tank, the propellant Flow Control System (FCS), and the
thruster itself. We place each component within the spacecraft as shown in Fig. 5.3.3-1.
2m
Fig. 5.3.3-1 Placement of the key propulsion system components in the overall vehicle.
(Brad Appel)
The EP system propels the spacecraft from Low Earth Orbit to Lower Lunar Orbit with a
slow, continuous thrust. Our payload is the wet mass of the Lunar Lander in addition to
the Space Balls. Table 5.3.3-1 lists the required EP system totals for the 100g mission.
Table 5.3.3-1 Electric Propulsion System Totals
Variable
Wet Mass
Dry Mass
Required Power
Burn time
Purchase Cost
Value
169
30
1,540
365
$1.03 million
Author: Brad Appel
Units
kg
kg
Watts
days
2009 dollars
Mission Configuration – 100g Payload – Lunar Transfer
Section 5.3.3, Page 49
The Hall Thruster
A Hall Effect Thruster provides three fundamental advantages for our mission. First, the
system has a very high payload mass fraction: 38% for our 100g mission. Second, HETs
physically require a small propulsion system - the thruster itself is no bigger than a
computer monitor, and its associated power conditioning equipment is about the size of a
PC tower. This compact feature drives down the volume as well as the mass of the
spacecraft. And third, the Hall Thruster system is inherently simpler than many other
propulsion options. This fact goes a long way in making the system cheaper to integrate
and more reliable overall.
Specifically, we use the BHT-1500 from the Busek Company for the main engine. There
are actually only a few commercial sources for Hall Thrusters, and from what data are
available, the BHT-1500 offers the best performance for our mission. Table 5.3.3-2
summarizes the performance and physical specifications of the thruster (Azziz , 2005).
The performance numbers (mass flow rate and specific impulse) are optimized by the
team in order to provide the cheapest overall system. For the best combination of high
power-to-thrust ratio and easy handling properties, we select Xenon as the propellant.
Table 5.3.3-2 Specifications for the Hall Thruster
Variable
Thrust
Specific Impulse
Mass Flow Rate
Power Input
Efficiency
Input Voltage
Mass
Value
78.5
1950
4.1
1526
0.53
350
5.7
Units
mN
s
mg/s
W
-VDC
kg
We place the HET protruding through the OTV floor paneling at the aft end of the
spacecraft. The thruster produces 700 Watts of excess heat, which we radiate away using
this Magnesium floor panel.
Author: Brad Appel
Mission Configuration – 100g Payload – Lunar Transfer
Section 5.3.3, Page 50
Power Processing Unit
The Hall Thruster comes with very specific and demanding power requirements. We
require a Power Processing Unit (PPU) to transform, monitor, and condition the power
input to the HET. Because the development of a new flight HET is usually accompanied
by the customization of a PPU, we logically select the Power Processing Unit from
Busek. Some important features of the PPU are listed in Table 5.3.3-3 (Osuga, 2007),
(Kay, 2001).
Table 5.3.3-3 Specifications for the Power Processing Unit
Variable
Operating Voltage
Input Power
Efficiency
Mass
Value
28
1352
93%
10
Units
VDC
Watts
-kg
The PPU serves other system functionalities as well. It collects all telemetry from the
HET, such as temperature and voltage data, and communicates with the spacecraft main
computer via a serial transmission cable. The PPU is also equipped to monitor and
command a mass flow controller, should the mission require it. The PPU has three critical
interfaces: Main power from PDCU to PPU, Commands and Telemetry from the central
computer to the PPU, and power input to the HET. These are all soft harness connections
and do not require a particular positioning for the PPU.
Within the spacecraft, we place the PPU forward of the rest of the propulsion system,
inside the OTV electronics compartment. This placement provides the most convenience
for thermal control and alignment of the spacecraft center of mass.
Xenon Propellant Storage and Flow Control System
We store the Xenon propellant in a supercritical state in order to minimize the volume it
takes up (supercritical means the gas is compressed to the brink of becoming a liquid).
This storage condition has two consequences: the tank must be strong enough to
withstand high pressure, and the tank must be temperature regulated to keep the Xenon
Author: Brad Appel
Mission Configuration – 100g Payload – Lunar Transfer
Section 5.3.3, Page 51
supercritical. We use a metal-lined, carbon composite overwrapped tank from Arde Inc.
The tank is spherical with a diameter of 60 cm and a dry mass of 12 kg. We initially
pressurize the tank to 15.2 MPa (2200 psia); however the pressure steadily drops
throughout the mission. For thermal control, a 5-Watt resistance heating wire is
embedded in the tank with thermostat regulation. We wrap the tank with Multi-Layer
Insulation in order to mitigate heat loss by radiation.
The tank feeds our Flow Control System (FCS), which conditions the Xenon flow to the
mass flow rate required by the Hall Thruster. Our FCS accomplishes four essential tasks:
First, a solenoid valve isolates the high pressure Xenon in the storage tank. Upon this
valve opening, the Xenon flows to a pressure regulator, where it is stepped down from
several thousand psi to around 40 psi (Moog Corporation, 2009). After passing through a
high-purity filter, the propellant flow is split into the three required feed lines for the
HET. The last stop for the Xenon before entering the thruster is through a flow restrictor,
which we calibrate to deliver the desired mass flow rate (Mott Corporation, 2009). These
flow restrictors are sensitive to temperature changes, and so we setup a heating coil and
thermostat system for the FCS as well.
Figure 5.3.3-2 is a schematic of the propulsion system on the OTV. A parts list is
included in Table 5.3.3-4.
Author: Brad Appel
Mission Configuration – 100g Payload – Lunar Transfer
Section 5.3.3, Page 52
Fig. 5.3.3-2 Electric Propulsion System Schematic.
(Brad Appel)
Table 5.3.3-4 Parts List for the EP System.
Label #
1
2
3
4
5
6
7
8
9
10
11
Component
Tank Heating Wire
Xenon Storage Tank
Tank Multi-Layer Insulation
Solenoid Latch Valve
Pressure Regulator
High Purity Filter
Sintered Flow Restrictors
FCS Heating Wire
Thruster Radiation Panel
Hall-Effect Thruster
Power Processing Unit
Author: Brad Appel
Mission Configuration – 100g Payload – Lunar Transfer
Section 5.3.4, Page 53
5.3.4 – Attitude
Attitude Control
Spacecraft attitude control is necessary to maintain a specific trajectory within the
translunar phase of our mission. We are controlling the Orbital Transfer Vehicle (OTV)
by a set of three systems: sensors, reaction wheels, and thrusters. These systems comprise
our attitude control system (ACS) and combine to provide the necessary attitude
corrections to accomplish our mission.
Our inertial and relative positions must be known before any attitude correction can be
done. We use a set of sensors to obtain and translate this knowledge. The star sensor,
made by Valley Forge Composite Technologies (VFCT), takes pictures of visible stars
and compares them to a database of star systems. Since stars are considered inertially
stable, the orientation of the OTV relative to space can be determined.
A Sun sensor, also made by VFCT uses pictures of the Sun to obtain the position of the
OTV relative to the Sun; therefore, determining the position of the spacecraft relative to
the Earth and Moon. As sensing is independent to the size of a spacecraft, we use these
same sensor systems for all missions. These systems, however, only provide a picture of
the OTV in space, and provide no control to the spacecraft.
Reaction wheels comprise one layer of attitude control within the OTV. A reaction
wheel uses electric motors that spin weighted plates to create torque. This torque is used
to make small corrections in the attitude of the OTV.
Our friends at VFCT also
manufacture reaction wheels. The VF MR 4.0 has enough capacity to overcome all
torques caused by environmental perturbations, thrust misalignment, and drift error. The
torque experienced has a cumulative effect and eventually will need to be de-spun to
recharge the torque-creating power of the reaction wheel.
Author: Brian Erson
Mission Configuration – 100g Payload – Lunar Transfer
Section 5.3.4, Page 54
Comprising the final layer of control, thrusters provide the force necessary to de-spin
reaction wheels and make independent attitude corrections.
Several systems were
analyzed according to their cost, inert mass, and propellant mass required. We choose a
hydrogen peroxide (𝐻2 𝑂2 ) monopropellant thruster to provide the counter torque needed
to de-spin the reaction wheels. The small thrusters are configured in a 4-axis redundant
system to provide reliable 3-axis control.
Our three attitude subsystems: sensors, reaction wheels, and thrusters combine to create a
complete system accomplishing the mission of controlling the OTV during the translunar
phase. Total attitude system mass and cost is small compare to mission entirety. VFCT is
providing a cost decrease pending the use of advertisement within our mission. Our OTV
provides more than enough power for mission success. System volume is small enough to
be easily integrated into the current OTV platform. Table 5.3.4-1 provides an overview
of our ACS specifications for the 100g payload.
Table 5.3.4-1 OTV ACS Budget for 100g payload
Device
VF STC 1 (star sensor)
VF SNS (sun sensor)
VF MR 19.6
𝐻2 𝑂2 thruster
𝐻2 𝑂2 propellant
Inert Mass
Totals
Mass (kg)
6.4
0.7
10.4
0.36
0.02
1.9
19.8
Cost ($) Power Required(W) Volume (𝒎𝟑 )
133,333
20.4
0.0163
133,333
5
0.0163
133,333
76
0.009375
1,500
-0.000009
100
--1,000
-0.0000696
403,000
101.4
0.0420536
Author: Brian Erson
Mission Configuration – 100g Payload – Lunar Transfer
Section 5.3.4, Page 55
Attitude Propellant Use
Four factors are used in determining the type of propellant within the ACS: mass
implications, cost of total system, current use of propellant on OTV, and ease of
integration.
As seen in Table 5.3.4-1, we choose 𝐻2 𝑂2 for its low weight and cost. The integration
requires a small increase in inert mass, but not enough to eliminate 𝐻2 𝑂2 from
contention. The propellant is not currently used on the OTV, but the preceding factors
outweigh the possibility of using a different propellant. 𝐻2 𝑂2 is a stable, nontoxic
substance that has long term storability. The performance characteristics outlined in
Section A-5.3.4 are adequate to provide enough thrust to fulfill all mission requirements.
Space Environment Perturbations
In this section we perform a “worst case” needs assessment for attitude control due to
environmental perturbing forces during the translunar phase. In order to compute the
maximum expected torques, a few assumptions are made. The largest assumption in this
analysis is that all forces add linearly and are of their greatest magnitude. Additionally,
both a right circular cylinder and a cube are used to approximate the shape of the OTV so
as to calculate environmental forces on it. Many variables, such as the spacecraft's
electric charge and reflective properties, are historically based on missions including
Spirit-1 and Lunar Surveyor (Berge, 2009). The sources of environmental disturbances
include electromagnetic radiation from the sun, reflected radiation from the Earth, Earth's
thermal radiation, an estimated amount of spacecraft radiation, solar wind, gravity
torques, meteoroids, Earth's magnetic field, and a nonpropulsive mass expulsion from the
spacecraft (Longuski, 1982)(deWeck, 2001). We then compute the corresponding torques
based on the moment arm for each force on the OTV (see environmental.m).
Even with these conservative estimates, the sum of all the torques is only on the order of
1.5x10-3 N-m for the cylinder and 1.5x10-5 N-m for the cube. If we constantly use an
Author: Brittany Waletzko
Mission Configuration – 100g Payload – Lunar Transfer
Section 5.3.4, Page 56
attitude control thruster to correct this torque, the cube only requires roughly 0.3 kg of
propellant (see environmentalpropmass.m). This compares to a historical average of
roughly twelve kilograms of attitude control propellant for spacecraft (deWeck, 2001).
As such, these torque corrections are small enough that they can be compensated for with
a reaction wheel.
Author: Brittany Waletzko
Mission Configuration – 100g Payload – Lunar Transfer
Section 5.3.5, Page 57
5.3.5 – Communication
Communication Hardware/Configuration
We find that during the Lunar Transfer it is necessary to maintain communication
between Earth and the OTV for 90% of the time. To make this happen we must consider
the Lunar Transfer trajectory and the orientation of the vehicle to insure the antenna is
pointed in a direction capable of sending and receiving a signal. We conclude that to meet
these requirements one system is all that we need. Instead of placing a system on the
OTV, which would be abandoned during Lunar Decent, we place one system on the
Lander. This one system fulfills all communication requirements for Lunar Transfer,
Lunar Decent, and for operations while on the surface of the Moon. More information
regarding the selected system and configuration can be found in the lunar decent Section
5.4.5.
Author: Joshua Elmshaeuser
Mission Configuration – 100g Payload – Lunar Transfer
Section 5.3.5, Page 58
Communication Ground Stations
In order to ensure our spacecraft behaves as intended we need to monitor it during the
Lunar transfer. In order to monitor our spacecraft we will make use of ground stations on
Earth. These ground stations and our spacecraft will communicate with each other using
radio antennae operating at 2.2 GHz in the S-Band range. Figure 5.3.5-1 which shows the
location of our ground stations as if the Earth is viewed looking at the North Pole as well
as the coverage “net” created.
Fig. 5.3.5-1 Our four ground stations. 100% tracking capability outside straight lines.
(Trenten Muller)
We use four ground stations. They are:
1) Mt. Pleasant Radio Observatory. Hobart, Tasmania, Australia. A 26 meter dish.
2) Hartebeesthoek Radio Astronomy Observatory (HRAO). Johannesburg, South
Africa. A 26 meter dish.
3) Pisgah Astronomical Research Institute (PARI). Rosman, North Carolina. USA.
One of the 26 meter dishes.
4) James Clark Maxwell Telescope. Mauna Kea Observatory, Hawaii, USA. A 15
meter dish.
Author: Trenten Muller
Mission Configuration – 100g Payload – Lunar Transfer
Section 5.3.5, Page 59
The latitude and longitude for these four ground stations are listed in Table 5.3.5-1.
Table 5.3.5-1 Geographic Locations of Ground Stations
Ground Station
Mt. Pleasant
HRAO
PARI
Maxwell
Latitude (o)
42.81 S
25.55 S
35.20 N
19.82 N
Longitude (o)
147.44 E
27.68 E
82.87 W
155.48 W
The four ground stations that we select will be used for all payload cases and will be used
during all mission phases. All have horizon to horizon scanning, meaning that they can
scan the entire sky at their locations, excluding geographical constraints. The James Clark
Maxwell Telescope will only be used to bridge the large gap between Mt. Pleasant and
PARI while within 8,400 km. The tracking ability of these stations will be above 90%.
Based on an e-mail conversation with Professor John M. Dickey (2009) of The
University of Tasmania, home to the Mt. Pleasant Observatory, we have estimated the
cost to operate these stations at $1 million for one year of usage.
Professor Dickey
provided a quote of $1 million for the use of the Mt. Pleasant Observatory for one year of
continuous use. Because we will only use one station at a time we will be able to spread
the $1 million quote amongst all of the stations as we use them individually throughout
the year.
Author: Trenten Muller
Mission Configuration – 100g Payload – Lunar Transfer
Section 5.3.5, Page 60
Communications Link Budget
We budget 0.5 kbps (kilobytes per second) for uplink and downlink communication with
the Orbital Transfer Vehicle (OTV) during the long trip to lunar orbit. Our
communication link consists of telemetry transmission from the OTV, as well as
commands from Earth to the OTV. The telemetry from our OTV consists of vital data
concerning the spacecraft’s operation, including information about attitude, power
consumption, and temperature. The commands that our OTV receives from Earth contain
data pertaining to mission operations procedures.
The 0.5 kbps we allocate for the communication link during the Lunar Transfer phase of
the mission is well below the bandwidth capabilities of our chosen communication
equipment. We choose not to employ this equipment to its full capability in an effort to
reduce power consumption during this lengthy voyage to lunar orbit. This electricity
conservation is vital to our mission, because of the power consumption of our electric
propulsion system and the premium of available power.
Communications Antenna Pivot
To communicate with ground stations during the Lunar Transfer phase of the mission our
spacecraft employs two patch antennae which are located on the Lunar Lander. We
choose to use the communication equipment on the Lunar Lander in order to eliminate
the need for redundant systems on the OTV. The elimination of unnecessary redundant
systems provides for a large savings in mass and cost.
As our spacecraft travels to lunar orbit its attitude in space will not always be conducive
to transmitting and receiving data with Earth ground stations. The high gain patch
antennae located on the Lunar Lander has a limited beam width, and needs to be pointed
towards Earth. As maneuvering the entire spacecraft for communication would be very
costly to the mission, we choose to simply gimbal the antennae. To do this gimbal
Author: Michael Christopher
Mission Configuration – 100g Payload – Lunar Transfer
Section 5.3.5, Page 61
activity we have designed a patch antenna pivot. This pivot design is pictured in Fig.
5.3.5-2.
Fig. 5.3.5-2 Patch antenna pivot.
(Michael Christopher)
As see in Fig. 5.3.5-2 the pivot consists of two plates, a base plate and an antenna mount
plate, two connecting arms, and two stepper motors. The base plate is mounted directly
onto the outer surface of our Lunar Lander. We choose to include two precession stepper
motors to first increase range of motion, as seen in Fig. 5.3.5-3 and Fig. 5.3.5-4. Second,
we choose to include two motors for redundancy. If a single motor fails, the other motor
will still be able to operate the pivot with a slightly limited range of motion.
Author: Michael Christopher
Mission Configuration – 100g Payload – Lunar Transfer
Section 5.3.5, Page 62
Fig. 5.3.5-3 Pivot in "normal" position.
(Michael Christopher)
Fig. 5.3.5-4 Pivot in 180 degrees from "normal" position.
(Michael Christopher)
The pivot integrated onto our Lunar Lander is displayed in Fig. 5.3.5-5; there is a
duplicate antenna/pivot assembly on the other side of the Lander. We include two of
these assemblies for yet more needed redundancy. With all of this redundancy we can
operate a communication link with three of the four motors out of operation, and one of
the two antennae out of operation. Without this redundancy, the communication system
would be a single point mission failure.
Author: Michael Christopher
Mission Configuration – 100g Payload – Lunar Transfer
Section 5.3.5, Page 63
Our choice not to include a communication system on the OTV, and use the one on the
Lunar Lander, presented a reliability challenge. This challenge was overcome by
including redundancy both in the pivot design, and in the placement of two identical
systems. At the extremely low cost of approximately $83.50 and low mass of about 0.2
kg, including two systems is wise design choice.
Fig. 5.3.5-5 Pivot integrated onto the Lunar Lander.
(Josh Elmshaeuser)
Author: Michael Christopher
Mission Configuration – 100g Payload – Lunar Transfer
Section 5.3.6, Page 64
5.3.6 – Thermal Control
Electronics Thermal Control
Three main components of the OTV need to be thermally controlled. These consist of the
electronics, the xenon tank, and the electric thruster. Because of the different thermal
requirements of each of these systems, each has a separate thermal control system. The
electronics comprise of the PCDU, the battery, the PPU, and the DC/DC converters.
Together, these components put out a total of 217W of heat. To cool these components,
we mount them on an aluminum plate, which removes the heat through conduction. We
note that the plate also serves as a structural support for the components. Attached under
the plate is a series of capillary heat pipes, with ammonia as the working fluid. When the
ammonia heats up and boils, the ammonia vapor travels to two radiators that are exposed
to space. After the heat radiates to space, the ammonia cools and condenses and, through
capillary action, travels back to the heat source under the electronics. We see the mass
and volume breakdown of the thermal control components for the electronics in Table
5.3.6-1.
Xenon Tank Thermal Control
To keep the xenon within its permissible storage temperatures, we embed a 5-Watt
resistance heating wire in the tank with thermostat regulation. We also wrap the tank
with Multi-Layer Insulation in order to mitigate heat loss by radiation.
Electric Thruster Thermal Control
The electric thruster is 55% efficient at converting its input power to thrust. The thruster,
therefore, generates approximately 639W of heat that we need to dissipate to keep the
thruster under its maximum operating temperature of 473K (200ºC). To maximize the
emissivity of the thruster, we paint the thruster with Z93 white paint. This increases the
amount of heat that the thruster is capable of radiating to space. At the base of the OTV,
a magnesium alloy shroud is installed around the thruster. The surface area of this shroud
is more than adequate to sufficiently radiate the thruster’s waste power to space.
Author: Ian Meginnis
Mission Configuration – 100g Payload – Lunar Transfer
Section 5.3.6, Page 65
We see the dimensions of the OTV’s thermal control system in Table 5.3.6-1. Because
we account for the aluminum shroud in the structure group’s mass budget, we do not
consider the shroud in this table.
Table 5.3.6-1 100g Payload OTV Thermal Control Dimensions
System
Component
Mass (kg)
Dimensions
Electronics
Ammonia
Heat Pipes
~0
2.22
Radiators
1.21
Wire Heater
Aluminum Shroud
1
N/A
I.D. = 3.35cm; O.D. = 3.65cm;
Length = 5m
Total Cross-Sectional Area =
0.895m2
(8 fins @ 0.112m2 each)
N/A
Xenon Tank
Electric Thruster
Author: Ian Meginnis
Mission Configuration – 100g Payload – Lunar Transfer
Section 5.3.7, Page 66
5.3.7 – Power
Introduction
The operational life of the Orbital Transfer Vehicle (OTV) extends to at least one year.
We employ a light-weight and reliable power generation system to provide sufficient
power for this time span. The OTV power is provided by two circular Ultraflex solar
arrays and a secondary (rechargeable) lithium-ion battery. During periods in sunlight, the
arrays provide adequate power to meet the OTV’s requirements. We do not assume,
however, that the spacecraft remains in sunlight throughout the transfer.
rechargeable battery provides power during periods of darkness.
The
Figure 5.3.7-1
illustrates a CATIA computer model of one of the partially deployed solar arrays.
Fig. 5.3.7-1 Top and side views of OTV Ultra-flex solar array CATIA model; partially deployed.
(Ian Meginnis)
Author: Ian Meginnis
Mission Configuration – 100g Payload – Lunar Transfer
Section 5.3.7, Page 67
Power Budget
We see the power budget for the OTV during the translunar phase broken down by group
in Table 5.3.7-1. The largest contributing factor to the OTV power budget arises from
the electric thruster. This device consumes over 75% of the entire OTV power budget.
To ensure all of the power needs are adequately met, we increase the total power
production by 5%. This increase accounts for fluctuations in power requirements by each
group that occur throughout the mission. The specific components for the individual
systems are mentioned in Section A-5.3.7 of this report.
Table 5.3.7-1 100g Payload OTV Power Budget by Group
Group
Power
Units
Propulsion
Communication
Attitude
Power
Lunar Lander
(during translunar)
TOTAL
1529
0
101.4
120
Watts
Watts
Watts
Watts
105
Watts
1959
Watts
Solar Array Sizing
With the power budget, we determine the size of the solar arrays. The arrays are sized to
meet the power requirements of the OTV during sunlight. Able Engineering produces the
two Ultraflex-175 solar arrays, which are comprised of triple-junction gallium-arsenide
solar cells. The power density for the arrays dictates the required mass and area. We see
the mass, stowage volume, and deployed area for the two solar arrays in Table 5.3.7-2.
Table 5.3.7-2 100g Payload OTV Solar Array Dimensions (total)
Parameter
Value
Units
Mass
Stowage Volume
Deployed Area
13.06
0.0664
6.54
kg
m3
m2
Author: Ian Meginnis
Mission Configuration – 100g Payload – Lunar Transfer
Section 5.3.7, Page 68
During launch, we store the circular solar arrays in two rectangular boxes on opposing
sides of the OTV. After the OTV reaches Earth parking orbit, we unfurl and deploy the
arrays.
We implement a sun sensor and two motors to ensure the arrays provide
maximum power. The motors provide two degrees of freedom for each array.
The high cost of the solar arrays directly affects the orbit trajectory and engine selection
for the OTV. We optimize the vehicle design to reduce overall mission cost. The cost
per watt of the triple-junction solar arrays is approximately $1000, which results in a total
cost of $1.96 million.
Battery Sizing
With the power budget, we also determine the size of the battery. The battery is sized to
provide power to the OTV during periods of darkness and prior to solar array
deployment. The rechargeable lithium-ion battery, manufactured by Yardney, has a
certain power density that determines the battery dimensions. We see these dimensions
in Table 5.3.7-3.
Table 5.3.7-3 100g Payload OTV Battery Dimensions
Parameter
Mass (includes housing)
Volume
Total Energy
Value
12.17
0.0043
1534
Units
kg
m3
W-hr
Power Subsystem Components
The power system for the OTV includes all devices necessary to generate, manage, and
distribute power to the individual OTV components. These devices include:

Solar Arrays

Batteries

Power Conditioning and Distribution Unit (PCDU)
Author: Ian Meginnis
Mission Configuration – 100g Payload – Lunar Transfer
Section 5.3.7, Page 69

Power Processing Unit (PPU)

Direct Current / Direct Current (DC/DC) Converters
Not to
scale
Fig. 5.3.7-2 OTV Subsystem Power Description.
(Ian Meginnis)
The solar arrays generate approximately 1.96kW with a direct-current output voltage of
~200V.
This power first travels to the PCDU.
Broad Reach Engineering custom
fabricates the PCDU. The PCDU conditions and regulates power from the solar arrays
and distributes this power to the battery, PPU, and DC/DC converters.
The OTV’s battery is available as a commercial off-the-shelf product.
The battery
consists of ten 3.6V, 43 A-hr lithium-ion battery cells. These cells group together to
produce the required 1.53kW-hr of energy that we need for the battery. If the battery is
not fully charged, we route excess power from the solar arrays to the batteries via the
PCDU. During periods of darkness, the PCDU detects a drop in power from the solar
arrays and draws power from the batteries to support the OTV components.
Author: Ian Meginnis
Mission Configuration – 100g Payload – Lunar Transfer
Section 5.3.7, Page 70
The PPU also receives power from the PCDU. The PPU increases the voltage from 100V
to the 340V the electric thruster requires. In order to power the OTV’s individual
components, we use a series of DC/DC converters. These devices decrease the voltage
that comes from the PCDU to whatever voltage is required by the individual components.
Because each component has a unique input voltage requirement, a separate DC/DC
converter exists for each component. A total of six DC/DC converters are used to
support the individual OTV components. Modular Devices Incorporated produces the
converters, which are available as off-the-shelf products.
Table 5.3.7-4 100g Payload OTV Power Subsystem Dimensions
Component
Solar Arrays
Battery
PCDU
DC/DC Converters
Mass (kg)
13.06
12.17
8.88
1
Volume (m3)
0.0664 (stowed)
0.0043
0.0157
0.000218
Author: Ian Meginnis
Mission Configuration – 100g Payload – Lunar Descent
Section 5.4.1, Page 71
5.4 – Lunar Descent
5.4.1 - Structures
Lander Structure
The most important consideration in designing the structure of our Lunar Lander for the
100g payload is the mass of the frame. We identify two key drivers for the mass of this
frame. These drivers are total volume and total mass of the Lunar Lander at lunar
touchdown.
The shape of the Lunar Lander for the 100g payload case is a conic frustum. A conic
frustum geometry can be pictured as a large cone with the top part chopped off as
depicted in Fig. 5.4.1-1. There are two reasons for using a conic frustum frame design.
First, a conic frustum houses the subcomponents of the Lunar Lander (such as space
balls, H202 tank, and main engine) while minimizing volume. Minimization of volume
also lowers the mass for thermal control. The various frame component masses can also
be recalculated easily for a conic frustum design without a Finite Element Analysis
(FEA) computational package. We provide the overall dimensions of the frame needed to
house all of the Lunar Lander subcomponents in Table 5.4.1-1; the total volume of the
Lunar Lander is 1.05 m3.
Table 5.4.1-1 Basic Frame Dimensions for 100g payload Lunar Lander
Variable
Height
Bottom Diameter
Top Diameter
Length of Legs
Value
1.0
1.3
1.0
0.607
Units
meters
meters
meters
meters
Figure 5.4.1-1 illustrates the main parts of the frame. The frame floor consists of a
circular outer ring and circular inner ring connected by four rectangular floor support
beams. An additional circular ring makes up the top of the Lunar Lander. The top
circular ring connects to the frame floor by four hollow circular side support beams.
Author: Ryan Nelson
Mission Configuration – 100g Payload – Lunar Descent
Section 5.4.1, Page 72
Fig. 5.4.1-1 Listing of Basic Lunar Lander Frame Components.
(Ryan Nelson)
These side support beams support the various loads subjected to the Lunar Lander
throughout the mission while maintaining a low thickness. The low thickness of the side
supports enables the storage of the four Lunar Lander legs within these side supports
during Earth launch and Lunar Transfer. A 0.5mm magnesium skin placed around the
entire Lunar Lander frame protects against micrometeorites and provides thermal
protection. The sum of all these frame components yields a total mass for the Lunar
Lander frame of 11.44 kg.
The total mass of the Lunar Lander at lunar touchdown determines the thickness, size and
mass of the individual frame components. We use a safety factor of 1.5 for all frame
components in the structural sizing code, Lander_Frame_ball.m. With the dry Lunar
Lander mass at touchdown and the Earth g’s experienced at lunar touchdown, the forces
seen on the Lunar Lander frame are calculated. The size and mass of the different frame
components are then calculated based on the types of loading that cause initial failure to
get the overall frame mass.
Author: Ryan Nelson
Mission Configuration – 100g Payload – Lunar Descent
Section 5.4.1, Page 73
Material Selection
The lander is made of the same materials that the OTV is fabricated from. The floor and
shear panels are fabricated from magnesium AZ31, while all other structural members are
made of aluminum 6065. All bolts, rivets, fasteners, and connectors are made of AIS
1015 low carbon steel.
Author: Korey Lemond
Mission Configuration – 100g Payload – Lunar Descent
Section 5.4.1, Page 74
CAD
We see in the following figures a CAD model created in the program CATIA. This model
gives us a visual representation of what our Lander looks like, and gives us a general idea
of where certain systems are located. The Lander assembly is designed with a color
coding system in mind to easily identify certain subsystems.
1) Red
–
Structures
2) Orange
–
Attitude
3) Yellow
–
Power
4) Green
–
Communication
5) Blue
–
Propulsion
6) Tan
–
Space Ball Housing
Author: Joshua Elmshaeuser
Mission Configuration – 100g Payload – Lunar Descent
Fig. 5.4.1-2 External View of the Lander with legs deployed.
(Joshua Elmshaeuser)
1 meter
H2O2 tank
Author: Joshua Elmshaeuser
Section 5.4.1, Page 75
Mission Configuration – 100g Payload – Lunar Descent
Section 5.4.1, Page 76
Fig. 5.4.1-2 Close up of the Landers internal components with legs stowed.
(Joshua Elmshaeuser)
Author: Joshua Elmshaeuser
Mission Configuration – 100g Payload – Lunar Descent
Section 5.4.1, Page 77
1.3 meters
Fig. 5.4.1-3 View of the Landers engine systems.
(Joshua Elmshaeuser)
Author: Joshua Elmshaeuser
Mission Configuration – 100g Payload – Lunar Descent
Section 5.4.1, Page 78
Fig. 5.4.1-4 Visualization of the Lander on the Moon.
Earthrise background is public domain on NASA’s Earth Observatory website:
http://earthobservatory.nasa.gov/IOTD/view.php?id=4882
(Joshua Elmshaeuser)
Author: Joshua Elmshaeuser
Mission Configuration – 100g Payload – Lunar Descent
Section 5.4.2, Page 79
5.4.2 – Mission Operations
Introduction
The purpose of this section of is to provide a complete description of the lunar descent
trajectory and landing site. We provide an overview of the descent phase beginning from
lunar parking orbit, as well as a description of the landing site. See Section 5.3.2 for a
description of the flight trajectory up to lunar parking orbit injection.
Descent Overview
At this stage, the Lunar Lander has separated from the orbital transfer vehicle and is in a
25 km circular orbit around the Moon. This orbit will be referred to as the lunar parking
orbit. The spacecraft remains in this orbit while descent system checks are performed.
After completing these checks, a command is sent to the Lunar Lander to begin descent.
A small burn is performed by the attitude thrusters, which places the Lunar Lander in an
elliptical trajectory with a perilune altitude of 13.4 km. This elliptical orbit is referred to
as the lunar descent transfer orbit. When the Lunar Lander reaches perilune of this
elliptical orbit, we begin final descent with Lunar Lander main engine ignition. Final
descent lasts 207 seconds culminating in touchdown on the surface of the Moon. Figure
5.4.2-1 gives a pictorial overview of the descent.
Lunar Parking Orbit
Moon
Lunar Descent Transfer Orbit
Final Descent
Fig. 5.4.2-1. An overview of the entire lunar descent. Note: Not to scale
(John Aitchison)
Author: John Aitchison
Mission Configuration – 100g Payload – Lunar Descent
Section 5.4.2, Page 80
Final descent is a continuous burn with four distinct phases: radial, rotation, vertical, and
landing coast. Figure 5.4.2-2 gives a pictorial overview of the descent. Please note that
the shape of this trajectory is not to scale; for to scale depictions of the trajectories see
section A-5.4.2.
2
1
3
4
1.) Begin radial burn
2.) End radial burn, begin rotation
3.) End rotation, begin vertical burn
4.) End vertical burn, coast to landing
Fig. 5.4.2-2 An overview of the final descent. After radial burn, the Lunar Lander throttles down
and rotates to vertical, then throttles up to maximum and begins vertical burn and coasts to a
landing. Note: Not to scale
(John Aitchison)
Tables 5.4.2-1 and 5.4.2-2 provide high level overviews of the descent.
Table 5.4.2-1 Final Descent Action Breakdown
Action
Radial Burn
Rotation
Vertical
Coast to Land
Time (s)
000 - 176
176 - 181
181 - 196
196 - 207
Thrust (N)
1150
115
1150
130
Author: John Aitchison
Mission Configuration – 100g Payload – Lunar Descent
Section 5.4.2, Page 81
Table 5.4.2-2 Descent Propellant Masses
Action
Time (s)
De-Circularize
Radial Burn
Rotation
Vertical
Coast to Land
Unusable
Contingency
Total
000 - 176
176 - 181
181 - 196
196 - 207
Propellant
Mass (kg)
0.4
67.5
0.2
5.8
0.5
2.3
1.5
78.2
Footnotes: This is the total propellant mass in
LPO.
Landing Site Approach
As the Lunar Lander approaches the landing site on final descent, it is critical that
adequate clearance between the spacecraft and mountainous terrain of the moon be
maintained. At all times, a clearance of over 6 km is maintained between the Lunar
Lander and the spacecraft, as depicted in Fig. 5.4.2-3.
Fig. 5.4.2-3 The green line is the descent trajectory, the blue is representative mountainous terrain
on the Moon. At all times, more than 6 km of clearance is maintained. Note: Not to scale
(John Aitchison)
Author: John Aitchison
Mission Configuration – 100g Payload – Lunar Descent
Section 5.4.2, Page 82
Landing Location
We choose Mare Cognitum (4°S 23°W) as our landing site. Figure 5.4.2-4 shows the
Moon and the location of Mare Cognitum. The landing location is relatively flat for
approximately 20 km in all directions.
Immediately beyond this area there are
mountainous regions with altitudes of 2780 m. At a distance of 60 km from the central
landing location the altitudes reach a maximum of 4300 m.
Mare Cognitum
Fig. 5.4.2-4 Landing site Mare Cognitum.
(www.science.nasa.gov)
Located about 20 km northwest from the landing site is the original landing of the Apollo
12 mission. Our vehicle is equipped with a camera to capture an image and video of the
landing location, where artifacts from the mission were left behind. Recording this
historical site will fulfill the Google Lunar X PRIZE Heritage Site Bonus. Figure 5.4.2-5
shows the anticipated landing site, the landing site of Apollo 12 and the surrounding
terrain.
Author: Kara Akgulian
Mission Configuration – 100g Payload – Lunar Descent
Fig. 5.4.2-5 Topographical map of the landing site.
(NASA/USGS/LPI/ASU)
Author: Kara Akgulian
Section 5.4.2, Page 83
Mission Configuration – 100g Payload – Lunar Descent
Section 5.4.3, Page 84
5.4.3 – Propulsion
Overview
We chose a radial flow hybrid engine as the Lunar Lander main engine. The oxidizer is
98% hydrogen peroxide (H2O2) and the solid fuel grains are Polyethylene (PE). As
shown in Fig. 5.4.3-1, we inject H2O2 around the perimeter of the combustion chamber
between the PE fuel grain plates. The combustion gases exit the combustion chamber
through the center hole in the lower fuel grain plates and into the nozzle. Catalyzing
chemicals integrated into the PE fuel grains causes rapid decomposition of the H2O2 and
allows raw H2O2 to be injected directly into the combustion chamber.
Fig. 5.4.3-1 Radial flow hybrid engine dimensions and configuration
(Thaddaeus Halsmer)
The dimensions of the combustion chamber are dictated by the regression behavior of the
fuel grain plates, total burn time and the desired oxidizer to fuel mixture ratio. We size
the chamber based on these properties and the desired thrust of the engine. The resulting
dimensions are shown in Fig. 5.4.3-1 along with the exit diameter of the nozzle.
Author: Thaddaeus Halsmer
Mission Configuration – 100g Payload – Lunar Descent
Section 5.4.3, Page 85
Performance/Operating Parameters
The Lunar Lander propulsion system is capable of being throttled to 10% of its maximum
thrust by adjusting the mass flow rate of the H2O2. Engine throttling allows us to meet
the retro-burn thrust and relatively low touchdown thrust requirements with a single
engine. Table 5.4.3-1 summarizes the primary performance specifications for the Lunar
Lander propulsion system.
Table 5.4.3-1 Hybrid Engine Performance Specifications
Parameter
Thrust Max / Min
Isp, vacuum average
Chamber Pressure Max / Min
Nozzle Area ratio Ae/At
Burn Time
Specification
1100 / 110
320
2.1 / 0.21
100
198.6
Units
[N]
[s]
[MPa]
––
[s]
Propellant Feed System
We selected a stored gas–pressurization engine feed system with helium being used as
the pressurizing gas. Figure 5.4.3-2 is the feed–system fluid schematic showing the
architecture of the fluid systems. Also shown is the helium purge system that clears the
oxidizer feed lines of any foreign debris before engine firing and flushes out residual
oxidizer upon shutdown. Table 5.4.3-2 presents the description of the corresponding
fluid control components.
Table 5.4.3-2 Fluid control components
Item
SV01
SV02
MOV
CK01
CK02
HV01
HV02
REG
F01
Description
Ullage Solenoid Valve
Purge Solenoid Valve
Oxidizer flow control valve
Check valve
Check valve
Manual vent/fill valve
Manual drain/fill valve
Ullage pressure regulator
Oxidizer filter/screen
Author: Thaddaeus Halsmer
Mission Configuration – 100g Payload – Lunar Descent
Section 5.4.3, Page 86
Fig. 5.4.3–2 Lunar Descent Propulsion Fluid Schematic.
(Thaddaeus Halsmer)
Power Requirements (during operation)
The number of electronic fluid control components in use at any given time dictates
electrical power requirement of the propulsion system during operation. Based on the
known power requirements of the individual components, the power required for the
propulsion system will not exceed 124 watts during its use.
Author: Thaddaeus Halsmer
Mission Configuration – 100g Payload – Lunar Descent
Section 5.4.3, Page 87
H2O2 Storage
The hydrogen peroxide is stored in a spherical aluminum tank. We use Aluminum 1060
due to its ability to reduce the Active Oxygen Loss (AOL) from the H 2O2 during the
mission. For a sealed Aluminum 1060 container the AOL for 90% H2O2 for 1 year is
0.1% of the initial hydrogen peroxide amount (Ventura, 2005). Hence, the amount of
H2O2 decomposed over the duration of 1 year for our mission is 0.078 kg. As compared
to the 78 kg of total propellant carried, this value is negligible.
The hydrogen peroxide tank is sized based on the tank operating pressure and the amount
of H2O2 used during the mission. Table 5.4.3-3 describes the pressure, size and mass of
the H2O2 tank. We use a safety factor of 1.5 in determining the following parameters.
Table 5.4.3-3 Pressure, Size and Mass of the H2O2 tanks
Parameter
Operating Pressure
Value
3.07
Units
MPa
Volume
Thickness
Diameter
Mass
0.0533
1.075
0.467
2.06
m3
mm
m
kg
The total volume of the oxidizer tank is the sum of the volume of the H2O2 required for
the mission (0.0512 m3: including the usable propellant volume, boil-off volume and the
volume of unusable propellant) and a 4% ullage volume (0.00205 m3). Ullage is the
additional volume added to the tank to give the propellant space to expand in case of
increasing temperature and varying pressures inside the tank.
Pressure Feed System
Helium is used as the pressurant gas for the pressure-fed system. We chose Titanium as
the material for the He tanks. Table 5.4.3-4 shows the detailed specifications of the
Author: Saad Tanvir
Mission Configuration – 100g Payload – Lunar Descent
Section 5.4.3, Page 88
pressurant tank as well as the mass of the Helium gas required to complete the landing
mission.
Table 5.4.3-4 Pressurant System parameters
Parameter
Operating Pressure
Value
21
Units
MPa
Volume
Thickness
Diameter
Mass of Tank
Mass of Helium
0.0257
5.76
0.366
0.87
0.89
m3
mm
m
kg
kg
Propulsion System Inert Mass
The inert mass of the propulsion system is largely dependent on the amount of propellant
used during lunar descent. With propellant masses obtained from the trajectory analysis
done by mission operations, we size the chamber, the oxidizer (H2O2) tank and the
pressurant system. We use historical as well empirical relations to size the rest of the
propulsion system. The total propulsion system inert mass for the hybrid landing engine
is approximately 30 kg. Table 5.4.3-5 gives a detailed breakdown of the propulsion
system inert masses.
Table 5.4.3-5 Propulsion system inert mass breakdown
Parameter
Value
Oxidizer (H2O2) Tank
Pressurant Gas (He)
Pressurant Tank (He)
Nozzle
Combustion Chamber
Injector
Skirts and Bosses
Feed System
Valves
Structural Supports
Propulsion System Inert Mass
Author: Saad Tanvir
2.06
0.89
0.87
12.28
3.15
0.95
0.32
0.32
6.94
2.15
29.91
Units
kg
kg
kg
kg
kg
kg
kg
kg
kg
kg
kg
Mission Configuration – 100g Payload – Lunar Descent
Section 5.4.3, Page 89
Thermal Analysis
The operating temperature of the propulsion system during the lunar transfer phase is
decided based on the safe operating temperature range of the H2O2 and the polyethylene.
We chose an operating temperature of 10 oC (283 K). The tanks are insulated with MultiLayer Insulation (MLI). We require 35W of constant power to maintain this temperature
throughout Lunar Transfer.
During lunar descent, the temperature of the propulsion system drops. The temperature
drop in the hydrogen peroxide tank is 6.0 K (from 283K to 277K). Similar temperature
drop from the valves and the feed lines are 2.0K and 2.1K respectively. Because the drop
in temperature of these components during lunar descent is small, its affect on engine
performance is negligible. Hence, there is no power required for thermal control on the
propulsion system during lunar descent.
Author: Saad Tanvir
Mission Configuration – 100g Payload – Lunar Descent
Section 5.4.4, Page 90
5.4.4 – Attitude Control
During lunar descent, we use the hybrid propulsion descent engine to slow the Lunar
Lander to a safe landing speed. Ideally the thrust points directly through the vehicle’s
center of mass. In practice, the thrust is offset from the center of mass due to (for
example) manufacturing limitations and center of mass changes as propellant is burned.
In addition to the thrust-offset torque acting on the vehicle, there are also much smaller
environmental perturbing forces that act on the Lunar Lander. We discuss the effect of
these perturbing forces in greater detail in Section 5.3.4 (which pertains to attitude
control during Lunar Transfer). During lunar descent, the environmental forces are so
much smaller than the thrust-offset force that we neglect them in our analysis.
The sensors of our attitude control system are described in Section 5.3.4. We place these
sensors on the lander so that they may be used in both Lunar Transfer and lunar descent.
A laser altimeter is also placed on the Lunar Lander to find the distance to the surface
during descent.
The attitude control actuators on the lander consist of twelve hydrogen peroxide thrusters.
We choose hydrogen peroxide thrusters because they draw their propellant from the
descent-engine propellant tank and thus save on complexity and cost. These thrusters
(manufactured by General Kinetics) are each capable of producing 13.3 N of thrust. The
twelve thrusters are arrayed around the vehicle as shown in Fig. 5.4.4-1.
This
configuration provides attitude control around all three of the vehicle’s axes.
Additionally, the attitude control thrusters provide the small initial burn beginning the
descent phase.
Author: Christine Troy
Mission Configuration – 100g Payload – Lunar Descent
Section 5.4.4, Page 91
Lander Top View
Lander Side view
Fig. 5.4.4-1 Lander thruster configuration schematic for 100g payload case.
(Christine Troy)
The properties of the vehicle and the thrusters govern the amount of attitude control
propellant required. The two most crucial factors in the amount of propellant required
are the distance from the vehicle’s center of mass to the thruster and the specific impulse
of the thruster. We select the vehicle configuration and the attitude thrusters and then we
can calculate how much propellant we need. Table 5.4.4-1 summarizes the relevant
properties of the lunar descent attitude control system, including the amount of propellant
needed for attitude control.
Table 5.4.4-2 Properties of lunar descent attitude control system for 100g payload
Variable
Torque about each axis
Thrust available for lunar descent
Attitude control propellant mass
Total mass of attitude thrusters
Value
25.2
56
3.21
1.08
Author: Christine Troy
Units
Nm
N
kg
kg
Mission Configuration – 100g Payload – Lunar Descent
Section 5.4.5, Page 92
5.4.5 – Communication
Communication Hardware/Equipment
Two communication systems are installed onboard the Lander in order to communicate
with Earth and our Space Balls. One system communicates with Earth. The other system
communicates directly with the Space Balls. These systems operate at a data rate of 51.2
kbps at a frequency of 2.2 GHz. The total mass, power usage, and purchase price is listed
in Table 5.4.5-1.
Table 5.4.5-1 Communication Equipment Totals
Mass (kg)
Power Usage (W)
Purchase Price (2009 $)
2.48
60.53
315,668
The communication systems consist of:
1) Two antennae to communicate with Earth.
2) Two antennae mount pivots.
3) A transmitter to transmit to Earth.
4) A receiver to receive signals from Earth
5) An antenna to communicate with the Space Balls.
6) A transceiver to transmit to and receive signals from the Space Balls.
7) A video camcorder to record video and take pictures
8) A Computer board to control these systems
Author: Trenten Muller
Mission Configuration – 100g Payload – Lunar Descent
Section 5.4.5, Page 93
Communication Power Requirements
The Lander sends information back to an Earth station at a data rate of 51.2 kbps. The
signal from the Lander requires a minimum of 33 Watts to get enough information
packets to the ground station. The frequency of the signal is 2.2 GHz, and allows
constant communication from the lunar surface.
Author: John Dixon
Mission Configuration – 100g Payload – Lunar Descent
Section 5.4.5, Page 94
Communication Timeline (and completion of GLXP requirements)
During the descent from lunar orbit to the surface of the moon our Lunar Lander is still
engaged in the 0.5 kbps (kilobytes per second) data link. This link is a continuation of the
data link used during the Lunar Transfer phase of the mission. This link still includes
telemetry from the spacecraft to Earth, and commands from Earth to the spacecraft.
Once on the lunar surface we begin our required transmissions to complete the Google
Lunar X PRIZE (GLXP) mission requirements. These files that we transmit to meet
mission requirements are outlined in Table 5.4.5-2. Photo files are produced from the
onboard digital camera. X PRIZE Foundation (XPF) set asides are files that are provided
by XPF and are stored on the craft’s computer memory before launch. These XPF set
asides include the first e-mail and text messages to be sent from the Moon. The data
uplink set is provided by the XPF, transmitted to the Moon, and then downloaded back to
the Earth. After the Locomotion phase of the mission begins, more mission requirements
will be transmitted. These transmissions are discussed in the Locomotion section of this
report, Section 5.5.3.
Table 5.4.5-2 Lunar Arrival Mooncast
Item
Photos
XPFb Set Asides
Data Uplink Set
Data Uplink Set
Link Directiona
Down
Down
Up
Down
Size [MB]
5
10
10
10
Footnotes:
a – Down is from Moon to Earth; up is from Earth to Moon.
b – X PRIZE Foundation.
Author: Michael Christopher
Transmission Time [hr]
0.24
0.47
0.47
0.47
Mission Configuration – 100g Payload – Lunar Descent
Section 5.4.6, Page 95
5.4.6 – Thermal Control
Lunar Day
We need to maintain operable temperature of the Lunar Lander from Lunar Transfer
through the lunar night. During Lunar Transfer, the side of the Lunar Lander that faces
the sun is at 394 Kelvin while the side out of the sun can be 2 Kelvin (Scott, 2009). The
temperature gain from the sun needs to be transported through the Lunar Lander to the
shaded side. During a portion of Lunar Transfer, part of the mission is spent where the
earth completely shades the sun from the Lunar Lander. During this time the Lunar
Lander’s internal temperature decreases. The Lunar Lander requires different types of
thermal control to maintain the operable temperature.
We use a multilayer insulation (MLI) blanket to reduce the amount of heat absorbed by
the Lunar Lander and to keep the heat in during the lunar night. As the surface area of the
Lunar Lander increases, the heat absorbed during the day increases and the heat out
during the night increases. For the 100g payload case, the MLI blanket will cover the
entire outside of the Lunar Lander, the Space Ball compartments, and the propulsion
system. The propulsion system and the Space Ball compartments will need the MLI
blanket because these systems should not undergo temperature differences.
We use a passive cooling device to reject heat in the Lunar Lander. This thermal control
system is similar to the thermal control for the orbital transfer vehicle (OTV). This
system includes an aluminum heat plate, a heat pipe and a radiator. The aluminum plate
is used to conduct the heat from the communication equipment. The communication
equipment sits on this aluminum plate. The area of the heat plate is the sum of the bottom
surface area of the communication equipment. This system can be seen in Fig. 5.4.6-1.
The ammonia heat pipe runs underneath the heat plate. This pipe is five feet long and
connects the aluminum plate to the radiators. Increasing temperature in the Lunar Lander
vaporizes ammonia in the heat pipes. Ammonia vapor flows to radiators via capillary
Author: Kelly Leffel
Mission Configuration – 100g Payload – Lunar Descent
Section 5.4.6, Page 96
action, rejecting heat from the Lunar Lander system. The ammonia will liquefy and be
brought back through the ammonia pipe by capillary action. This action is continuous
during the mission.
Fig. 5.4.6-1 Schematic of Passive Thermal Cooling System.
(Josh Elmshaeuser)
The heat pipe connects to the radiator on the Lunar Lander. The radiator on the side that
faces the sun allows the ammonia to flow through the radiator with little or no
temperature increase. The radiators are painted white to reflect this heat. The radiator in
the shade of the Lunar Lander rejects the heat flowing through the heat pipe. Figure
5.4.6-2 shows this heat exchange. The radiator’s area is large enough to expel the heat the
Lunar Lander gains at a peak heat time. The peak heat time is on the lunar surface when
Author: Kelly Leffel
Mission Configuration – 100g Payload – Lunar Descent
Section 5.4.6, Page 97
the sun covers the entire Lunar Lander and all the communication equipment is running.
The radiator consists of aluminum and sits at an angle where the radiator will be able to
expel the heat from the bottom when the sun is directly overhead.
Fig. 5.4.6-2 Heat transfer in the Lunar Lander - Heat enters on the side with the sun and
leaves on the shaded side both through the Lunar Lander and out the radiators.
(Kelly Leffel)
We operate two valves, one on each heat pipe system, to control the temperature of the
Lunar Lander. These valves are the yellow section on Fig. 5.4.6-2. As the temperature of
the Lunar Lander decreases due to the lack of sun, either in the earth’s shade or the lunar
night, the valves will close. This process eliminates the rejection of heat when the Lunar
Lander is at the minimum operable temperature.
We use heaters to control the temperature of some components of the Lunar Lander
during Lunar Transfer. The propellant and oxidizer tanks along with the ball
compartments use resistance heaters to replace heat lost to space throughout the Lunar
Transfer. The resistance heaters replace the amount of the heat lost by the equipment.
Two additional heaters sit on the heat pipe where the valves are located. These heaters
Author: Kelly Leffel
Mission Configuration – 100g Payload – Lunar Descent
Section 5.4.6, Page 98
provide heat to the ammonia if the temperature of the radiator drops below 195 Kelvin,
the freezing temperature of ammonia (Gilmore, 2002). The total power required by the
thermal control system during Lunar Transfer is 10 Watts.
The mass of the thermal equipment was a significant driver in choosing a thermal control
system. The current system that we use has a total mass of 9.57 kilograms. The
breakdown of the thermal control subsystems are found in Table 5.4.6-1. The total mass
was calculated in the MATLAB program LanderThermalControl.m.
Table 5.4.6-1: Day Thermal Control of Lunar Lander
Components
MLI Blanket
Aluminum Plate
Heat Pipe
Radiators
Ammonia
Heaters
Total
Author: Kelly Leffel
Mass (kg)
2.35
1.40
2.60
2.70
0.021
0.50
9.57
Mission Configuration – 100g Payload – Lunar Descent
Section 5.4.6, Page 99
Surviving the Lunar Night
The ambient temperature of the lunar surface during its fourteen day dark phase
plummets to 140 degrees Kelvin which is far below the tolerance ranges for many of the
Lunar Landers’ components. Heat is lost through conduction by contact with the surface
and through radiation. We minimize thermal losses by using multi-layer insulated
blankets to reduce radiation losses from the protected volume. Total thermal losses for
the 100 gram payload configuration are expected to be about 11 Watts which must be
replaced during the period of darkness.
We use chemical energy to provide heat to the Lander during the lunar night. We choose
hydrazine for its simplicity of use and for its high energy content. Less than four
kilograms of hydrazine are required to survive the lunar night, and is contained in a small
tank wherein a rod of Aluminum/Nickel catalyst is extruded by a bimetallic actuator as
seen in Fig. 5.4.6-3.
Fig. 5.4.6-3 Hydrazine heating system (right) and catalyst actuator (left).
(Tony Cofer)
Author: Tony Cofer and Adham Fakhry
Mission Configuration – 100g Payload – Lunar Descent
Section 5.4.6, Page 100
The thermostatic element must be experimentally calibrated to provide the desired
equilibrium temperatures throughout the thermally controlled volume
We minimize power requirements during lunar night by shutting down all electrical
equipment. A small electrical switch powers the Lunar Lander down when solar power
drops below a predetermined threshold and turns it on again at the lunar dawn.
Author: Tony Cofer and Adham Fakhry
Mission Configuration – 100g Payload – Lunar Descent
Section 5.4.7, Page 101
5.4.7 – Power
Battery
We size the battery for the Lander based on power requirements from the Lander engine,
communication and attitude systems. The Lander engine, the communication gear and the
attitude systems all depend on the battery during the landing phase.
The Lander’s engine requires power to operate the various valves and pumps to make the
engine function, which total 124 Watts and will be operating for 250 seconds. The
communication gear uses a peak total power of 60.53 Watts, to ensure that we can
maintain constant communication between the Earth and the Lander throughout the entire
Landing phase. We provide enough power for the communication gear to transmit for 27
minutes at peak power consumption. The attitude system uses a peak total power of 25.4
Watts to power our system and requires 15 minutes of peak power to make sure the
Lander arrives at its destination. Table 5.4.7-1 highlights the total power requirement for
all systems and their individual components on the 100 g Lander.
Table 5.4.7-1 Power breakdown of system on Lander
Item
Power
Units
Propulsion
124
Watts
Communication
60.53
Watts
Attitude
25.4
Watts
Total
209.93
Watts
Note: Complete description of each device can be found
in section 5.4.2
For a complete background on each device within each system, please refer Section A5.4.7.
We purchase the batteries from the company Yardney Technical Products, Inc., who
provide readymade lithium ion batteries that are used in space vehicles.
Author: Adham Fahkry
Mission Configuration – 100g Payload – Lunar Descent
Section 5.4.7, Page 102
We select a 12 Ampere-hour battery, but the battery only requires 11.69 Ah. We decided
that as a safety precaution the battery should have a little bit extra power, in case of any
problems that might occur after separating from the OTV. Refer to Section A-5.4.7 to see
how we calculated the capacity of the battery in Ah.
In order to avoid loss of charge when not in use, which happens with all lithium ion
batteries, our battery on the Lander will be completely drained during integration with the
entire spacecraft. The team will send a command that will allow the battery to charge
when it approaches the moon, using the OTV solar cells before separation. We compute
the battery size using lander100g.m.
Solar Array Sizing
We use solar cells to provide enough power to the communication system during the twoweek long lunar day. The peak power required is 60.53 W to run all of the
communication equipment, refer to Section A-5.4.7 to see breakdown of power
requirements for each device within the communication equipment.
Based on these specifications, we set up the solar cells in a deployed configuration for the
entire mission. The cells are statically fixed to reduce costs and complexity. The cell has
an area of 0.785 m2. Figure 5.4.7-1 displays the power consumed and the power
generated by the cells through a lunar day.
Author: Adham Fahkry
Mission Configuration – 100g Payload – Lunar Descent
Section 5.4.7, Page 103
Fig. 5.4.7-1 Maximum Potential Power of solar cells during a lunar Day.
The communication gear has enough power to meet its peak power requirements
on the 1.7st day of the lunar day. At maximum power, the solar cells provide
74.5% more power than required at day 7 of the lunar day.
(Adham Fakhry)
The red line represents the power used by the Communication gear, the blue line
represents the power produced by the cells. The solar cells provide more than enough
power required by the communication gear. Also, the excess power ensures the dust
accumulated on the lunar surface or impacts from meteorites, will not severely affect the
power collected by the cells as they provide extra power. We sized the solar cells to meet
the communication system’s power requirements during data transmission from the Earth
to the Moon.
The solar cells are provided by Able Engineering, who will be supplying solar cells for
the Orbital Transfer Vehicle and the Lander. The Lander solar cells will cost $250,000
and the solar cells will also weigh a maximum of 2 kilograms. Solar array sizing is
Author: Adham Fahkry
Mission Configuration – 100g Payload – Lunar Descent
Section 5.4.7, Page 104
automated in lander.m. Refer to Section A-5.4.7 for breakdown of costs and the
constants used to size the solar cells.
DC-DC Converters
The DC-DC converters for the 100 gram case depend on the number of component within
the system and the voltage they each require to operate, the DC-DC converters change
the voltage from the battery and the solar cells to a higher voltage for each individual
device will require. Table 5.4.7-2 below displays the DC-DC converter system for the
100 g payload.
Table 5.4.7-2 DC-DC converter system
Component
Quantity
Units
Mass
Dimensions
Temperature Range
Cost
0.725
0.033 x 0.033 x 0.033
-50 to 150
$ 21,000
kg
m
C
-
Note: Cost is in 2009 US dollars and is the cost for all the units.
All the converters will be placed in a single aluminum box that will be connected to the
PCDU (power conditioning and distribution unit) that will then branch out to each system
and their various components.
Power Conditioning and Distribution Unit (PCDU)
A power conditioning and distribution unit (PCDU) is needed to safely manage the power
distribution throughout the Lander during mission. Refer to Section A-5.4.7 for complete
breakdown of the PCDU. We choose a PCDU from Terma Incorporated to handle a 240
Watts power system that will be connected to our power system components, listed
below, and the PCDU will cost $12,000 and weigh approximately 1.9 kg.
The PCDU will interface with the following sources:
Author: Adham Fahkry
Mission Configuration – 100g Payload – Lunar Descent
Section 5.4.7, Page 105
1. Rechargeable Lithium-ion battery
2. Solar Cells
3. DC-DC converters
Table 5.4.7-3 shows the specifications of the PCDU.
Table 5.4.7-3 Specifications of the PCDU on the Lander
Component
Mass
Dimensions
Temperature Range
Cost
Quantity
1.932
0.033 x 0.033 x 0.033
-20 to 60
$12,000
Notes: Cost is in 2009 US dollars.
Author: Adham Fahkry
Units
kg
m
C
-
Mission Configuration – 100g Payload – Locomotion
Section 5.5.1, Page 106
5.5 – Locomotion
We have reached the moon. Everything goes silent as the engine shuts off. Suddenly, a
panel blows off of the bottom of the Lander and out pops a small round sphere the size of
a basketball. After a few moments, the ball begins rolling away from the Lander; quickly
picking up speed. Traversing the rolling lunar landscape and bouncing up and over small
bits of moon rock, the Space Ball is on its way to completing the mission we have come
so far to achieve.
Some basic characteristics of the Space Ball are listed in Table 5.5-1 but the following
sections describe all of the different systems that make up the locomotion phase of the
100g payload mission in greater detail.
Table 5.5 - 1 Space Ball Specifications
System Category
Mass
Diameter
Cruise Speed
Min Turning Radius
Value
2.44
0.25
1.04
0.0625
Fig. 5.5-1 Artistic Rendition of Space Ball on the Moon.
(Cory Alban)
Author: Cory Alban
Units
kg
m
m/s
m
Mission Configuration – 100g Payload – Locomotion
Section 5.5.1, Page 107
5.5.1 – Structures/Integration
Space Ball Deployment
The Space Ball resides inside a thermally insulated box with protective foam surrounding
the lower half. We choose to deploy the Space Ball using a linear shaped charge. The
linear shaped charge is made of copper lining with C-4 as the explosive material. The
linear shaped charge surrounds the Space Ball on three edges of the bottom panel. The
fourth edge of the deployment door has a hinge. The charge dislodges the bottom of the
casing, which swings open on the hinge. The Space Ball then drops to the surface ready
to begin the mission. Figure 5.5.1-1 shows us a cross sectional view of the Space Ball
deployment system. Table 5.5.1-1 shows the mass of the deployment system.
Thermally
insulated box
Space Ball
SOLIMIDE Foam
Linear shaped charge
Fig. 5.5.1-1 View of the Space Ball inside the thermally insulated box
surround by SOLIMIDE foam for protection from the linear shaped
charges that cut the bottom of the Lunar Lander for deployment.
(Caitlyn McKay)
Table 5.5.1-1 Mass of deployment system per Space Ball
System
Mass (kg) per Space Ball
Charge
0.580
Foam
0.040
Total
0.620
Author: Ryan Lehto
Mission Configuration – 100g Payload – Locomotion
Section 5.5.1, Page 108
CAD
We arrange the components as shown in Fig. 5.5.1-2. The outer shell seals the Space
Ball and the components inside. A main axle fixed to the shell houses the main drive
motor as well as a stepper motor for turning. In addition, the pendulum mass and the dust
mitigation motor mount to the drive axle. The 100 g payload, CPU, transceiver, camera,
and battery mount to the pendulum swing arms.
Outer Shell
Dust
Removing
Vibrating
Motor
Main Axle &
Motor Housing
CPU
Transceiver
Camera
100g Payload
Fig. 5.5.1-2 Space Ball internal components.
(Ryan Lehto)
Author: Ryan Lehto
0.25
m
Mission Configuration – 100g Payload – Locomotion
Section 5.5.1, Page 109
Fig. 5.5.1-3 The Space Ball is
roughly the same size as a
basketball.
(Ryan Lehto)
An advantage to the Space Ball system is its small size. The Space Ball is similar in size
to a basketball (about 0.25 m diameter) see Fig. 5.5.1-3.
Author: Ryan Lehto
Mission Configuration – 100g Payload – Locomotion
Section 5.5.1, Page 110
Structural Analysis
We choose Lexan, a clear polycarbonate material, for the shell of the Space Balls. The
clear outer shell allows us to mount our camera inside of the ball rather than having to
mount it to the exterior. Keeping everything inside of the ball reduces system complexity
and the number of failure modes. Every fragile component in the Space Ball is protected
by the sturdy Lexan shell. Due to the high impact resistance of Lexan, the Space Ball can
survive an impact with a rock at full cruising speed without cracking the shell or breaking
sensitive components. Pressurized nitrogen inside the Space Ball adds additional stress
to the shell.
The aluminum drive axle runs through the center of the ball and carries a bending
moment from the equipment mounted on the torsion arm. The axle has additional torsion
stress from the drive system.
Tables 5.5.1-2 and 5.5.1-3 describe both the physical structure of the Space Balls and
their load bearing capabilities.
Table 5.5.1-2 Space Ball Load Bearing Capabilities
Failure Mode
Load
Can structure handle load? Factor of Safety
Torsion Stress
0.03N-m
Yes
59
Bending Moment 15.69N-m
Yes
100
Pressure Vessel
101.325kPa
Yes
2.8
Table 5.5.1-3 Space Ball Structure
Structure Component
Size
Shape
Lexan Shell
3.82mm Thick, 25cm Diameter Hollow Sphere
Aluminum Axle
25cm Long, 6cm Diameter
Circular Rod
Aluminum Torsion Arm
12.5cm Long
Rectangular Rod
Author: Kara Akgulian
Mission Configuration – 100g Payload – Locomotion
Section 5.5.1, Page 111
Dust Mitigation
As the Space Ball rolls over the lunar surface, dust will accumulate on the shell of the
Ball. This gathering and sticking of the lunar dust can potentially cause problems. It can
block the view of the camera, prohibiting us from achieving the goals of the Google
Lunar X PRIZE. In addition, buildup of dust increases the risk of it entering the core of
the Space Ball where all the critical mechanical components are. The dust penetration
into the drive system could potentially wedge the mechanisms, causing a failure by
prohibiting them from moving.
Angstrom Aerospace has created inflatable robots to explore distant planets. They are
using an ultrasonic motor to vibrate off the dust (Marks, 2008). We have adapted their
method of dust removal by implementing a linear ultrasonic Piezo drive Motor to vibrate
the dust off the surface of the Ball. The model selected is M-674,164 PILine designed by
Physik Instrumente (PI). At only 38 mm2 and weighing 0.1 kg this option made it the
most efficient solution for dust removal. The motor creates a frequency of 155 kHz,
which vibrates the Space Ball. This frequency is more than sufficient to shake off the
dust.
The ultrasonic motor consists of a rod mounted between two piezo linear motors. The
rod of the motor is attached to the center shaft of the Space Ball so that both motors
oscillate back and forth at a high frequency. This intense motion causes the ball to
vibrate. The motor will not be continuously running it will only be used when dust
begins to accumulate.
When the motor is needed the ball will not be in motion.
Author: Cory Alban
Mission Configuration – 100g Payload – Locomotion
Section 5.5.2, Page 112
5.5.2 – Propulsion
System Design
We implement a center of mass shifting system for propelling the Space Balls. The
Space Ball propulsion system works by swinging a mass, much like a pendulum to shift
the ball’s center of mass. The system includes two electric motors, a continuous D/C
motor mounted on a main driveshaft, and a D/C stepper motor positioned perpendicular
to the main drive motor for turning. The motor system all together has a mass of 0.172
kg. A benefit of the space ball propulsion system is the use of the essential components
as part of the drive system. The pendulum mass includes the 100 g payload, CPU control
board, batteries, and the communication equipment.
Main Drive Motor
We use a 6 mm diameter D/C motor paired with a 6 mm planetary gear head as the main
drive system to propel the Space Ball for the mission. The gear multiplies the motor’s
torque and slows down the motor’s output shaft speed. The motor controls the
forward/back movement by lifting the hanging mass and thus shifting the Space Ball’s
center of mass see Fig. 5.5.2-1. The motor gear combination creates a output torque of
0.03 N-m. The torque is more than capable of accelerating the space ball at 0.0043 m/s2
with an average velocity of 1.04 m/s. The Space Ball travels 500 meters in 8 minutes.
Fig. 5.5.2-1 The forward and backward motion of the
Space ball as controlled by the main drive motor.
(Ryan Lehto)
Author: Ryan Lehto
Mission Configuration – 100g Payload – Locomotion
Section 5.5.2, Page 113
We control left to right movements with the stepper motor. The stepper motor steps and
holds the drive system pendulum mass in 15° increments up to 30° in the desired turning
direction. The left/right movement of the pendulum tilts main drive axle and as result
turns the Space Ball see Fig. 5.5.2-2. The space ball is able to turn in a radius of 0.0625
m.
Fig. 5.5.2-2 The pendulum moves left or right to turn the Space Ball.
(Ryan Lehto)
The main drive motor needs 0.5435 W to propel the ball. Thus, for the entire mission the
space ball requires 0.0725 W-hr to travel 500 m.
Author: Ryan Lehto
Mission Configuration – 100g Payload – Locomotion
Section 5.5.3, Page 114
5.5.3 – Communications
Communication Hardware/Configuration
The Lander relays communications from the Space Balls to Earth and vice versa. For
communications equipment on the Space Balls the limiting factors are size, and mass. We
require the smallest possible components that still have the capability to transmit back to
the Lander at a distance of 500 meters. During hardware selection we consider size,
power and price. The Space Balls contain the following four pieces of communication
equipment.
1) UHF Antenna
2) Transponder
3) CPU
4) Camera
In compliance with the GLXP guidelines any vehicle placed on the moon must take one
self-portrait during the mission, so it is necessary that we place a camera in the Space
Balls. We purchase the camera from Ahlberg Electronics. The camera is designed to be
radiation tolerant enough to handle nuclear power facilities and is suitable for subjection
to space conditions.
We will employ monopole UHF antennae and UHF transceivers which provide a high
data transmission rate over short distance. The UHF antennae allow quick, cheap, and
power efficient communication over the short distance that the Space Balls travel. The
Space Quest Company provides these required systems at the cheapest price.
We use a CubeSat Kit Flight Module as the CPU for the Space Balls. The CPU processes
all the information sent from the Lander to the Space Ball, as well as the pictures taken
by the camera on board.
Author: Joshua Elmshaeuser
Mission Configuration – 100g Payload –Locomotion
Section 5.5.3, Page 115
Communication Power Requirements
The communications equipment on the SpaceBalls was chosen after the power and
communications requirements were known. The SpaceBalls will send information back
to the Lander at a data rate of 51.2 kbps. The signal from the SpaceBalls requires a
minimum of 3 Watts to get enough information packets to the Lunar Lander. The
frequency of the signal is 2.2 GHz, and will allow constant communication from the
SpaceBalls furthest planned distance from the Lunar Lander.
Author: John Dixon
Mission Configuration – 100g Payload –Locomotion
Section 5.5.3, Page 116
Communication Timeline (and completion of GLXP requirements)
After the Locomotion phase on the lunar surface we transmit proof of our locomotion to
fulfill Google Lunar X PRIZE (GLXP) mission requirements. These files consist of 8
minutes of real time video, 8 minutes of high definition video, and 5 megabytes (MB) of
photographs. The Locomotion Mooncast is outlined in Table 5.5.3-1.
Table 5.5.3-1 Locomotion Mooncast
Item
8 min Near Real Time Video
8 min High Definition Video
Photos
Link Directiona
Down
Down
Down
Size [MB]
75
900
5
Transmission Time [hr]
3.53
42.37
0.24
Footnotes:
a – Down is from Moon to Earth; Up is from Earth to Moon
After our survival of the lunar night, for our survival bonus, we have yet again another
Mooncast. This Mooncast is named the Survival Mooncast, and it is outlined in Table
5.5.3-2. All of the same data is transmitted except for the high definition video.
Table 5.5.3-2 Survival Mooncast
Item
8 min Near Real Time Video
Photos
Link Directiona
Down
Down
Size [MB]
75
5
Footnotes:
a – Down is from Moon to Earth; Up is from Earth to Moon
Author: Michael Christopher
Transmission Time [hr]
3.53
0.24
Mission Configuration – 100g Payload - Locomotion
Section 5.5.4, Page 117
5.5.4 – Thermal Control
The inside of the Space Balls will be filled with Nitrogen Gas. The Nitrogen Gas’s
thermal properties enable it to absorb most of the heat present inside the Space Balls,
keeping the temperature rise to a minimum. The Aluminum frame inside the Space Balls
absorb the heat in the Nitrogen Gas and dispel it to space. The Nitrogen Gas has a very
low mass, and a high specific heat, making it an efficient, cost-mass effective cooler of
the Space Balls. The Aluminum frame is an essential support structure and is not
considered an additional mass in the thermal management system, despite acting as such.
Author: John Dixon
Mission Configuration – 100g Payload –Locomotion
Section 5.5.5, Page 118
5.5.5 – Power
Power System
We select three Lithium Manganese Dioxide coin batteries to power each Space Ball.
The non-rechargeable primary batteries are provided by Efficient Energy & Managed
Battery. Each of these batteries provide 3 volts and 0.26 ampere-hours. The dimensions
for each of these batteries are 23mm in diameter and 3mm in height shown in Fig. 5.5.51. This battery system produces 0.78watt-hours per battery and a total of 2.34 watt-hours
for all three batteries before the mission leaves Earth. After the twelve month transit to
the moon, the batteries lose 45% of the original charge.
3m
mm
23mm
Fig. 5.5.5-1 Lithium Manganese Battery CR2330 cylindrical
coin dimensions. [not to scale]
(Jeff Knowlton)
We chose these batteries based on a number of reasons, they provide enough power for
the mission and the correct voltage for components, the most predominate factor is that
they have a mass of only four grams which puts the total mass at twelve grams not
including housing. We budget out the power to each of the required systems as shown in
Fig. 5.5.5-2. The three batteries are sized to complete the 500-meter drive and transmit
data back to the Lunar Lander. The battery we are using provides a margin of safety of
1.52 for all of the power needs of the Space Ball.
Author: Jeff Knowlton
Mission Configuration – 100g Payload –Locomotion
Section 5.5.5, Page 119
Battery Distribution
CPU
Reserve
34%
Drive
Motors
6%
Transmission
55%
Camera
5%
Fig. 5.5.5-2 Space ball Battery distribution over
locomotion mission phase.
(Jeff Knowlton)
Author: Jeff Knowlton
Mission Configuration – 100g Payload – Locomotion
Section 5.5.6, Page 120
5.5.6 – Mission Operations
Lunar Surface Terrain
Lunar regolith is a fine and dry powder. The regolith has a small electric charge, which
causes it to stick to almost anything it touches. The adhesive dust coats the shell of the
Space Ball; however, we employ a dust mitigation technique to address this issue.
Table 5.5.6-1 shows the probability of encountering craters at our landing site.
Table 5.5.6-1 Lunar Crater Hazard Analysis
Crater Size
Small (0-.02m)
Medium (0.02m-4m)
Large (4m+)
Encounter
Probability
Very High
High
Very Low
Is Space
Capable?
Yes
Yes
No
Ball
Solution
Roll Over
Avoid or Roll Through
Avoid
The Space Ball is large enough that craters of small size will not cause mobility
problems. The medium sized craters can be avoided by selecting an appropriate mission
path. The Space Ball can roll over and through medium sized craters while traveling at its
cruising speed. Large craters will be avoided.
Table 5.5.6-2 shows the probability of encountering rocks or other debris .
Table 5.5.6-2 Lunar Debris Hazard Analysis
Debris Size
Small (0-.02m)
Medium (0.02m-4m)
Large (4m+)
Encounter
Probability
High
Medium
Very Low
Is Space Ball
Capable?
Yes
Yes
Yes
Solution
Roll Over
Avoid or Dodge
Avoid
The majority of rocks the Space Ball encounters are of equal or smaller diameter as the
ball itself. In the event of a collision with a rock, the Space Ball is designed to withstand
a full speed impact with large objects. Just as a car wheel rolls over pebbles and small
rocks, the Space Ball runs up and over any objects that are smaller than the outer shell.
Turning capability and choosing an appropriate mission path will eliminate chances of
encountering large debris.
Author: Cory Alban
Mission Configuration – 100g Payload – Locomotion
Section 5.5.6, Page 121
Completion of Mission Requirements
We complete the final mission requirements using the Space Ball. The following is a list
of our objectives.
1) The secondary vehicle deployed by the Lunar Lander must move a distance of
500m on the surface of the Moon in a deliberate manner.
2) The vehicle must carry a 100g payload to the surface of the Moon.
3) Transmit an “Arrival Mooncast” and a “Mission Complete Mooncast” to
Earth
To complete objective one, the Space Ball will complete a series of tasks on the lunar
surface. The Space Ball powers on and runs a systems diagnostic from inside of the
Lander to verify that all systems are working properly. Once all systems are up and
running, the ball is deployed onto the lunar surface. Mission control sends instructions to
the ball that indicate what direction to travel from the Lander based on the surrounding
terrain conditions. Photographs taken from the Lander shortly after touchdown and
satellite images of the area will be the primary factors in determining the navigation path.
The ball accelerates along the surface in the direction of travel until reaching 1.04 meters
per second. The Space Ball maintains this speed for eight minutes. At the end of eight
minutes of travel the ball reaches a distance of 500m from the landing vehicle and
achieves the mission objective.
To complete objective two, the thermal control, propulsion system, and structural
components of the Space Ball are designed with the knowledge that this 100g payload
will need to be carried the 500m distance and kept within reasonable thermal limits. By
satisfying objective one, objective two is subsequently met.
To complete objective three, the ball carries a camera and takes the required self portrait
picture and sends it back to the Lander. The ball turns itself ninety degrees to its right
during the braking maneuver at the 500m travel mark. This turning maneuver ensures that
Author: Cory Alban
Mission Configuration – 100g Payload – Locomotion
Section 5.5.6, Page 122
the camera faces back toward the Lander. In the event that moon dust is obscuring the
camera’s view, a vibration motor will run until the dust is shaken off and the view to the
Lander is clear. The ball transmits the image back to the Lander for eight minutes. Once
the Lander receives the entire image, it sends the data along with the other required
pieces of the “mission complete mooncast” back to Earth. The photo verifies that the ball
did travel the full distance and completes objective three.
Table 5.5.6-3 gives a breakdown of each step of the lunar mission along with a time
breakdown for the mission.
Table 5.5.6-3 Space Ball Mission Breakdown
Step Time(minutes) Tasks to be Completed
1
0
Space Ball performs a system diagnosis.
2
1
Deployment from Lander.
3
2
Direction of travel received from mission control.
Space Ball orients to path of travel.
4
2-10
Accelerate to cruising speed of 1.04m/s.
Travel for 8 minutes until 500m objective achieved.
5
11
Braking maneuver with a 90 degree orientation change to point
camera toward Lander. Shake off dust if necessary.
6
12
Snap photo of Lander from ball and begin transmission.
7
20
Finish Photo Transmission.
Author: Cory Alban
Mission Configuration – 100g Payload – Mission Integration
Section 5.6, Page 123
5.6 – Mission Integration
Two types of integration are necessary when attempting to design a space mission, both
system and mission integration. The first, system integration, deals with putting all the
pieces together, essentially the subsystems, into a working vehicle. In our case, this could
be the OTV, the Lander, or the Space Ball. Once all of these are properly integrated, it is
necessary to integrate the set of vehicles into one mission, therefore the term mission
integration. The first step in the process we undertook was to design the subsystems
present in each vehicle. In general each system is designed to minimize its mass
contribution toward the craft’s total mass. Once the pertinent dimensions are obtained,
the problem of system integration then becomes a much more comprehensive issue.
Several priorities must be considered when integrating a system. The most important
consideration is the center of mass of the set of vehicles. During Lunar Transfer, the
electric thruster must be oriented such that it fires through the center of mass. However, if
an interior component is moved ever so slightly, the center of mass of the vehicle moves
off that line of action. Therefore it is important to maintain a center of mass along that
line of action. Also of importance is thermal control of the mission. All components must
be placed in a position where the thermal control system can cool or heat these
components. An iterative process then ensues to ensure that this happens. Finally, of great
importance in a design study such as this is whether or not this vehicle could actually be
manufactured and assembled in the manner in which it is designed. This is an issue that
requires a more intuitive approach. While we are certainly not well versed in
manufacturing principles, we endeavored to make sure that we could actually construct
the structures involved with this mission.
In the process of integrating the OTV, it became clear early on in the design that it would
be beneficial to put all of the electronics equipment in one electronics module. This was
done so all of the equipment giving off heat could be addressed by one common thermal
control system. After the module was designed, the electronics components such as the
Author: Korey LeMond
Mission Configuration – 100g Payload – Mission Integration
Section 5.6, Page 124
DC converters are then all arranged in the module such that they are supported by the
floor beams but still maintain an axial center of mass relative to the spacecraft as a whole.
From here the heat pipes and radiators are attached, as their positions were determined by
the needs of the electronics module to eliminate heat and eject it into space. At this point,
the frame of the OTV, the electronics module, and the thermal control systems were all in
place as we see in Fig. 5.6-1.
Fig. 5.6-1 OTV Frame.
(Korey LeMond)
As no other component needed to be integrated that was not axisymmetric about the line
of action of the electric thruster, the integration process was simply a puzzle of where
these subsystems could be attached on the vehicle. An example of this principle is the
attachment of solar arrays. The array is comprised of two panels that deploy directly
opposite each other relative to the axis of the OTV. Therefore neither will move the
center of mass off the thruster’s firing line, and it matters relatively little where they are
attached.
Author: Korey LeMond
Mission Configuration – 100g Payload – Mission Integration
Section 5.6, Page 125
The Lander also required integration as a vehicle system. This process occurred in much
the same way as the OTV. The frame was designed to take the loads applied at launch,
and then the subsystems were attached with the same priorities, aligning the center of
mass along the engine’s firing line and overall thermal control of the vehicle. The main
difference in the integration of the Lander is the thermal control system. This vehicle was
pressurized to establish a gaseous environment more readily controlled, which in turn
caused extensive problems that will be detailed more in the next few paragraphs under
mission integration.
When considered individually, each of these vehicles, the OTV, Lander, and the two
Space Balls, were all fairly self sufficient entities. This greatly reduces the thought that is
put into the mission integration as a whole. Several problems did exist however. The first
was the fact that power had to be transmitted to the Lander during Lunar Transfer. This
meant that a cable had to run from the OTV to the Lander during that mission phase.
Ordinarily this is not a problem. However, with the additional knowledge that the Lander
was pressurized, creating a connection that will separate when the Lander separates from
the OTV without compromising the integrity of the Lander is a much more difficult
problem. We found that it is possible to use a fastener port that essentially plugs into a
socket on the Lander, thus creating a power line from the power system of the Lander to
that of the OTV. In effect this would look like the OTV was unplugging from an
electrical socket when the vehicles separate, much like it looks when one unplugs a
simple household appliance.
Another issue with mission integration was the fact that the brain of the mission, the
CPU, was located on the Lander. This means that the OTV is effectively ‘dumb’ upon
launch. This was solved by inserting a fiberoptic cable into the power feed line, in
essence killing two birds with one stone.
Author: Korey LeMond
Mission Configuration – 100g Payload – Mission Integration
Section 5.6, Page 126
The final problem with integrating the OTV and Lander was the fact that they are not the
same size diameter. The Lander has a diameter over half a meter smaller than the OTV.
This then called for an adapter skirt seen in Fig. 5.6-2.
Fig. 5.6-2 Adapter Skirt.
(Korey LeMond)
This skirt had to be strong enough to take the compression launch loads, while also stiff
enough to take the bending stresses caused by the Lander wanting to move laterally due
to launch loads on top of the OTV. The main problem then became connecting the
Lander and OTV to this skirt. We choose to use a series of explosive pyrobolts to connect
the Lander and OTV to the skirt. A pyrobolt is a simple explosive composed of
ammonium nitrol and TNT in small amounts that will gently push the Lander and skirt
apart.
The final integration problem is then integrating the system of vehicles to the fairing atop
the launch vehicle. This was accomplished using a circular steel ring riveted into the
OTV floor with pyrobolts that was then attached to the fairing wall. The steel ring is then
fired off the floor of the OTV using the pyrobolts when the vehicle tumbles out of the
fairing in low earth orbit. The final mission configuration can be seen with an appropriate
scale, without the shear paneling, in Fig. 5.6-3.
Author: Korey LeMond
Mission Configuration – 100g Payload – Mission Integration
Fig. 5.6-3 Final Mission Configuration.
(Korey LeMond)
Author: Korey LeMond
Section 5.6, Page 127
Mission Configuration – 100g Payload – Risk Analysis
Section 5.7, Page 128
5.7 – Risk Analysis
The probability of success of the mission is computed by combining the probability of
success for each vehicle in the mission. We add a redundant space ball in order to boost
the probability of success for the mission. Table 5.7-1 describes the probability of
success for each vehicle.
Table 5.7-1 Vehicle Success Rate
Vehicle
Dnepr Launch Vehicle
Orbital Transfer Vehicle
Lunar Lander
Space Balls (redundant)
Mission Success Rate
Success Rate
94%
88%
88%
98%
72%
In order to satisfy the 90% mission success rate requirement, a contingency launch of a
duplicate mission is planned. This duplicate mission will be launched immediately
following any failure of the first mission.
Table 5.7-2 Mission Success Rate
Number of Missions
One
Two
Success Rate
72%
92%
Although the individual mission success rate is not acceptable, planning a contingency
launch artificially increases our overall system reliability to satisfy the mission
requirement. The necessity of a contingency launch is closely connected to total mission
cost – a lower mission success rate allows for a less costly mission. We are taking a
gamble that our first mission will work correctly and win the GLXP purse. If our first
mission fails, it is prudent to fix the problem, build, and launch a duplicate mission to win
the GLXP purse in order to reduce capital loss.
Author: Solomon Westerman
Mission Configuration – 100g Payload – Cost Analysis
Section 5.8, Page 129
5.8 – Cost Analysis
Table 5.8-1 Total Mission Cost
Expense
Dnepr Launch Vehicle
Orbital Transfer Vehicle
Lunar Lander
Space Balls
Overhead
Total Mission Cost
$/kg Payload
Cost ($M)
4.80
8.69
4.93
0.10
8.58
27.1
270.9
Our costing method tends to underestimate the total mission costs. We anticipate actual
mission costs will be higher than our estimate. Table 5.8-2 tabulates net profit if we
consider the GLXP purse money.
Table 5.8-2 GLXP Purse and Relative Cost
Expense
Total Mission Cost
GLXP Purse
Net Profit
Cost ($M)
(27.1)
22.3
(4.8)
We anticipate a net loss in capital in this mission.
Author: Solomon Westerman
Mission Configuration – 10kg Payload
Section 6, Page 130
6 – Mission Configuration 10kg Payload
Author: Solomon Westerman
Mission Configuration – 10kg Payload – System Overview
Section 6.1, Page 131
6.1 – System Overview
The 10kg payload system configuration meets the mission requirements of surviving the
lunar night, taking then sending pictures and videos, finding a heritage site and moving
500 meters. The system consists of an Orbital Transfer Vehicle and a Lunar Lander, Fig.
6.1-1. The heritage site we choose is where Apollo 12 landed. The Lunar Lander
descends onto the surface close enough that pictures are able to capture the site. The
Lunar Lander takes 8 minutes of high-definition footage and a picture of the panoramic
view. The Lunar Lander will then move by “hopping” from the landing site to a safe area
at least 500 meters away as we see in Fig. 6.1-2. After more pictures have been taken,
the Lunar Lander will shut down for one lunar night. After the lunar night the Lunar
Lander turns back on and sends a signal to allow us to know that it survived. The
complete mission timeline, after the 365 days to travel to lower lunar orbit, we see in
Table 6.1-1 and the masses of each stage in Table 6.1-2.
Fig. 6.1-1 Exploded view of OTV and
Lunar Lander.
(Korey LeMond)
Author: Caitlyn McKay
Mission Configuration – 10kg Payload – System Overview
Section 6.1, Page 132
Fig. 6.1-2 Schematic of the Lunar Lander
hopping 500 meters.
(Joshua Elmshauser)
Table 6.1-1 Mission timeline
Elapsed Time
ddd:hh:mm
-365:00:00
0:00:00
0:00:04
0:00:12
0:03:44
0:03:45
0:03:59
0:04:01
0:12:01
2:06:24
2:06:36
15:23:24
Event
Launch
Lunar Lander reaches LLO and separates from OTV
Lands on lunar surface and starts video taping
Finishes taping and begins transmission of video
Completes video transmission and takes panoramic pictures
Finishes panoramic pictures and begins transmission of pictures
Completes picture transmission and begins hop for locomotion
Locomotion phase complete and begins HD video taping
Begins transmission of HD video and takes panoramic pictures
Ends transmission of HD video and begins transmission of pictures
Ends transmission of pictures and shuts down for lunar night
Turns on and sends signal after lunar night.
Footnotes:
Elapsed time is in days:hours:minutes.
The time begins when the Lunar Lander is in lower lunar orbit and is detached from
the OTV.
Author: Caitlyn McKay
Mission Configuration – 10kg Payload – System Overview
Section 6.1, Page 133
Table 6.1-2 Mission Masses
Phase
Mass delivered to LEO
Mass delivered to LLO
Mass delivered to surface
Payload delivered
Mass in kg
583.5
228.3
107.0
10
We designed to place the 10kg payload on the surface with the least amount of mass.
Typically the greater the mass, the more it cost to put an object on the moon. The
following sections give a more detailed look at our design.
Author: Caitlyn McKay
Mission Configuration – 10kg Payload – Launch Vehicle
Section 6.2.1, Page 134
6.2 – Launch Vehicle
6.2.1 – Launch Vehicle/Site
We select the Dnepr-1 rocket to place the 10 kilogram payload into a low Earth Orbit.
The configuration for this case only weighs 127 kilograms more compared to the 100
gram payload mission. This mass difference does not exceed the Dnepr’s capabilities,
thus the same performance standards presented in section 5.2.1 can be applied here.
Author: Zarinah Blockton
Mission Configuration – 10kg Payload – Launch Vehicle
Section 6.2.2, Page 135
6.2.2 – Earth Parking Orbit Selection
Atmospheric Drag
We analyze the effects of atmospheric drag for the 10 kg payload in the same manner as
for the 100 g payload in Section 5.2.2. We reach the same conclusions and select a
minimum parking orbit of 400 km.
Parking Orbit Selection
Similar to Section 5.2.2, we choose the lowest altitude parking orbit possible. We select
a 400 km circular parking orbit as it overcomes drag effects and yields the minimum cost.
Author: Andrew Damon
Mission Configuration – 10kg Payload – Launch Vehicle
Section 6.2.3, Page 136
6.2.3 – Attitude determination in LEO
This mission configuration uses the same attitude determination subsystems and
methodology as outlined in Section 5.2.3 with no changes.
Author: Kristopher Ezra
Mission Configuration – 10kg Payload – Lunar Transfer
Section 6.3.1, Page 137
6.3 – Lunar Transfer
6.3.1 – Structure
Much of the information presented in this section contains differences between Orbital
Transfer Vehicle (OTV) configurations in the 100g and 10kg payload missions. We
leave out all the similar ideas and material presented for the 100g payload mission. For a
better explanation on various topics discussed in this section please refer back to Section
5.3.1. Refer to Section A-6.3.1 for the details and methodology used to determine our
design for the 10kg payload mission.
Limit Loads
Launch loads or limit loads are obtained from the Dnepr User’s Manual, and used to
evaluate the capability of our spacecraft design just as in the 100g payload case.
Similarly, we only concern ourselves with the launch loads for sizing and designing our
spacecraft, because we assume they will provide an upper bound limit for our vehicle.
Table 6.3.1-1 shows these loads for the Dnepr launch vehicle. The loads specified are
valid only if the spacecraft meets certain stiffness, or natural frequency requirements.
These requirements are summarized in Table 6.3.1-2 for the Dnepr launch vehicle.
Table 6.3.1-1 Dnepr launch limit loads
Event
Axial
Acceleration
Lateral
Acceleration
1st Stage Burn: Maximum Lateral Acceleration
3.0 ± 0.5
0.5 ± 0.5
2nd Stage Burn: Maximum Longitudinal Acceleration 7.8 ± 0.5
Footnote: Table based from Dnepr User’s Guide
Table 6.3.1-2 Dnepr spacecraft stiffness requirements
Thrust (Hz)
20
Lateral (Hz)
10
Footnote: Table based from Denpr User’s Guide
Author: Tim Rebold
0.2
Mission Configuration – 10kg Payload – Lunar Transfer
Section 6.3.1, Page 138
Margin of Safety
We apply the same factor of safety (FS) of 1.5 when we conduct our analysis. The yield
margin of safety (MS) which we use in all of our analyses is shown as:
𝑀𝑆 =
𝐴𝑙𝑙𝑜𝑤𝑎𝑏𝑙𝑒 𝑦𝑖𝑒𝑙𝑑 𝑙𝑜𝑎𝑑 (𝑜𝑟 𝑠𝑡𝑟𝑒𝑠𝑠)
−1
𝐷𝑒𝑠𝑖𝑔𝑛 𝑦𝑖𝑒𝑙𝑑 𝑙𝑜𝑎𝑑 (𝑜𝑟 𝑠𝑡𝑟𝑒𝑠𝑠)
(6.3.1-1)
Configuration
The configuration and layout for the OTV is based on the same design as the 100g
Payload mission. Changes in this design include different thicknesses of individual
structural members to handle different loads produced from a different sized payload.
The main difference between the two missions comes from the change in payload size.
During launch the payload will produce forces on the OTV. The payload (Lander
Lander) becomes heavier in the 10kg payload mission but has a lower center of mass. As
a result most structural members in the 10kg payload mission will experience loads
different of what they experienced in the 100g payload mission. Below we discuss the
changes to individual components and systems.
Stiffeners
The stiffeners (C-Channels) remain the same length with changes made to their cross
section. Since the Lunar Lander (LL) has a lower center of mass in the 10kg Payload
mission (while being slightly heavier), the C-Channels have less of a structural
requirement. This is because bending moments are reduced from the lower center of
mass, which will raise the critical buckling load of each individual stiffener. Therefore,
the stiffeners can be designed lighter and meet their structural requirements.
Propulsion Module
The propulsion frame remains the same as in the 100g Payload mission. See Section
5.3.1 for details on the configuration and layout of the support frame.
Author: Tim Rebold
Mission Configuration – 10kg Payload – Lunar Transfer
Section 6.3.1, Page 139
Electronics Module
The electronics module (E-MOD) remains the same as in the 100g payload mission. See
Section 5.3.1 for details on the configuration and layout of the support assembly.
Skin
The skin also remains the same as in the 100g payload mission. See Section 5.3.1 for
details on the configuration and layout of the skin
Integration
Integration structure includes the Payload Attach Fitting (PAF) and LL skirt. A complete
description of their geometry and functions can be found in section 5.3.1. Both change
due to the change in payload mass and center of mass. The thicknesses in the skirt walls
are optimized to reduce weight while meeting their structural requirements. The LL skirt
ends up getting heavier in this design because the walls need to be thicker to prevent
buckling. This is a result of a heavier payload mass. The PAF ends up getting lighter in
this design since its walls are thinner. A lighter PAF results from the natural frequencies
of the spacecraft being higher because of a lower center of mass. We use the PAF to
drive these fundamental frequencies high enough to meet the Dnepr stiffness
requirements. We also figure that we can add as much mass to the PAF as necessary
since it is separated from our spacecraft after the launch is finished. The PAF itself is
strong and stiff enough that buckling and yielding are not a concern. Detailed geometry
of the LL skirt and PAF can be seen in Section A-6.3.1.
Sizing
The basic principles that we use to size the OTV in the 100g payload mission are also
used in the 10kg payload mission. See Section A-5.3.1 for complete details.
Author: Tim Rebold
Mission Configuration – 10kg Payload – Lunar Transfer
Section 6.3.1, Page 140
Material Selection
The 10kg orbital transfer vehicle is made of the same materials that the 100 gram OTV is
fabricated from. The floor and shear panels are fabricated from magnesium AZ31, while
all other structural members are made of aluminum 6065. All bolts, rivets, fasteners, and
connectors are made of AIS 1015 low carbon steel.
Finite Element Analysis (FEA)
Information presented in this section contains the differences in the 10kg Payload Orbital
Transfer Vehicle (OTV) design versus the 100g payload OTV design.
Refer to Section
5.3.1 for a complete explanation of the analysis. Section A-5.3.1 presents all the details
and methods used to conduct our analysis.
FEA Analysis Results
As for the 100g payload OTV design, we perform a similar finite element analysis (FEA)
for our 10kg payload design.
We can perform a similar analysis, because the
configuration of the OTV remains almost identical for both designs. The only substantial
difference from the previous mission is in the payload characteristics. Table 6.3.1-1
shows the structural component masses as a result of these analyses.
Table 6.3.1-1 Component mass totals for OTV design
Components
PAF*
E-MOD floor beams & overlay
Shear / Skin Panels
Propulsion Support Frame
Stringers / Stiffeners
Lander Skirt
Fasteners (welds, rivets, bolts, adhesives)**
TOTAL
Footnotes: Not included in final OTV mass
Estimate
Author: Tim Rebold
Mass (kg)
41.36
5.25
15
3.01
12.12
14.24
2.12
51.74
Mission Configuration – 10 kg Payload – Lunar Transfer
Section 6.3.2, Page 141
6.3.2 – Mission Operations
Trajectory
We design the trajectory for the 10 kg payload similar to the 100 g payload. The 10 kg
payload results in slightly different initial conditions as seen in Table 6.3.2-1.
Table 6.3.2-1 10 kg Payload Trajectory Configuration
Parameter
Payload Mass (kg)
Thrust (mN)
Mass Flow Rate (mg/s)
Earth Phase Angle (deg)
Moon Phase Angle (deg)
Parking Orbit Altitude (km)
Capture Orbit Altitude (km)
Initial Mass (kg)
Flight Time (days)
Value
228.1
104
5.3
83
221
400
25
585.6
365
This model inherently contains a mismatch in position and velocity at the intersection of
the spiral out and spiral in curves (see Fig. 6.3.2-1). We calculate the propellant mass
required to produce the ΔV mismatch operating the main engine. We add this propellant
mass to the OTV to account for the error in position and velocity. We assume that
performing small maneuvers throughout the trajectory eliminates the fairly large bias in
position and velocity at the intersection point. Table 6.3.2-2 contains these mismatch
values.
Table 6.3.2-2 10 kg Payload Trajectory Configuration
Parameter
Position (km)
Velocity (km/s)
Propellant Mass (kg)
Mismatch
687.6
435.1
13.1
Figure 6.3.2-1 illustrates the resultant trajectory.
Author: Levi Brown
Mission Configuration – 10 kg Payload – Lunar Transfer
5
Section 6.3.2, Page 142
Spiral Out and In of Spacecraft
x 10
2
Spiral Out from Earth
Spiral In to Moon
1.5
1
(km)
-0.5
hat
0
y
0.5
-1
-1.5
-2
-2.5
-1
0
1
x hat(km)
2
3
4
5
x 10
Fig. 6.3.2-1 10 kg Payload Trajectory.
(Levi Brown)
Lunar Capture Orbit
The lunar capture orbit is the same as the 100g payload case, see section 5.3.2.
Trajectory Correction Maneuver
Similar to Section 5.3.3, we calculate the propellant necessary to perform a 50 m/s burn
with the main engine. The OTV has larger mass for the 10 kg payload, which results in
more propellant required for the correction as seen in Table 6.3.2-1.
Table 6.3.2-1 Correction Maneuver Configuration
Parameter
Isp (s)
mo (kg)
Propellant for Correction (kg)
Author: Levi Brown
Value
1964
585.6
1.5
Mission Configuration – 10kg Payload – Lunar Transfer
Section 6.3.3, Page 143
6.3.3 – Propulsion on the Orbital Transfer Vehicle
Propulsion Subsystem Overview
Although the OTV sees a 46% increase in payload mass in this mission as compared to
the 100g payload case, this actually has a minimal effect on the Propulsion hardware. The
only differences are the power required and total Xenon propellant required. Table 6.3.31 summarizes the propulsion system parameters for our 10kg mission.
Table 6.3.3-1 Propulsion System Totals
Variable
Wet Mass
Dry Mass
Required Power
Burn time
Thrust
Specific Impulse
Mass flow Rate
Value
215
30
2,043
365
104
1964
5.4
Author: Brad Appel
Units
kg
kg
Watts
days
mN
s
mg/s
Mission Configuration – 10kg Payload – Lunar Transfer
Section 6.3.4, Page 144
6.3.4 – Attitude
Attitude Control
As mentioned in Section 5.3.4, three systems are integrated to perform attitude control of
the OTV.
1. Sensors – The same set of sensors are used for all payloads: a sun sensor and star
sensor, both made by VFCT
2. Reaction Wheels – Due to the increased mass of the OTV, a larger reaction wheel
must be used to maintain attitude stability throughout the trans lunar phase.
VFCT manufactures the VF MR 10.0, a wheel capable of creating adequate torque
for the system.
3. Thrusters – The 𝐻2 𝑂2 thruster hardware needed is the same as the 100g case.
Our three attitude subsystems: sensors, reaction wheels, and thrusters combine to create a
complete system accomplishing the mission of controlling the OTV during the translunar
phase. Below, Table 6.3.4-1 provides an overview of our ACS specifications for the 100g
payload.
Table 6.3.4-1 OTV ACS Budget for 10kg payload
Device
VF STC 1 (star sensor)
VF SNS (sun sensor)
VF MR 10.0
𝐻2 𝑂2 thruster
𝐻2 𝑂2 Propellant
Inert Mass
Totals
Mass (kg)
6.4
0.7
20
0.36
1.26
3.02
31.74
Cost ($)
133,333
133,333
133,333
1,500
100
1,000
403,000
Author: Brian Erson
Power Required(W)
20.4
5
120
---145.4
Mission Configuration – 10kg Payload – Lunar Transfer
Section 6.3.4, Page 145
Attitude Propellant Use
As mentioned in Section 5.3.4, 𝐻2 𝑂2 is an adequate propellant to fulfill all mission
requirements.
Space Environment Perturbations
The analysis for the 10kg payload is identical to the analysis for the 100g payload
detailed in Section 5.3.4.
Author: Brian Erson
Mission Configuration – 10g Payload – Lunar Transfer
Section 6.3.5, Page 146
6.3.5 – Communication
Communication Hardware/Configuration
For the 10kg payload we will be using the same systems as for the 100g payload. For
more information refer to Section 5.3.5 and Section 5.4.5.
Communication Link Budget
Please see Section 5.3.5 for the communication link budget during the lunar transfer. The
communication requirements and equipment are unchanged with the change in vehicle
and payload mass.
Communication Ground Stations
For the 10kg payload the ground station usage is the same as the 100g payload and can be
found in Section 5.3.5.
Communication Antenna Pivot
The patch antenna pivot discussed in Section 5.3.5 will be used again for the 10kg
payload case, and will be essentially unchanged.
Author: Michael Christopher
Mission Configuration – 10kg Payload – Lunar Transfer
Section 6.3.6, Page 147
6.3.6 – Thermal Control
Electronics Thermal Control
Similar to the 100g payload case, we thermally control the electronics, the xenon tank,
and the electric thruster for the 10kg payload OTV. The electronic components put out a
total of 283W of heat. To cool these components, we employ the same thermal control
system in the 100g case. In Table 6.3.6-1, we see the mass and volume breakdown of the
thermal control components for the electronics.
Xenon Tank Thermal Control
For the 10kg payload case, we take advantage of a simple wire heater and multi-layer
insulation to keep the xenon within its storage temperatures.
Electric Thruster Thermal Control
The electric thruster generates approximately 849W of heat that we need to dissipate to
keep the thruster under its maximum operating temperature of 473K (200ºC). We see the
dimensions of the OTV’s thermal control system in Table 6.3.6-1.
Table 6.3.6-1 10kg Payload OTV Thermal Control Dimensions
System
Electronics
Component
Ammonia
Heat Pipes
Radiators
Xenon Tank
Electric Thruster
Wire Heater
Aluminum Shroud
Mass (kg) Dimensions
~0
2.53
ID = 3.83cm; OD = 4.13cm;
Length = 5m
1.576
Total Cross-Sectional Area =
1.168m2
(8 fins @ 0.146m2 each)
1
N/A
N/A
Author: Ian Meginnis
Mission Configuration – 10kg Payload – Lunar Transfer
Section 6.3.7, Page 148
6.3.7 – Power
Similar to the 100g payload case, the OTV for the 10kg case is powered by two circular
Ultraflex solar arrays and a secondary (rechargeable) lithium-ion battery. Due to the
increased power requirements, however, the dimensions of both the solar arrays and the
battery change. We see a CATIA computer model of one of the partially deployed solar
arrays in Fig. 6.3.7-1.
Fig. 6.3.7-1: Top and side views of OTV Ultra-flex solar array CATIA model; partially deployed.
(Ian Meginnis)
Power Budget
We see the adjusted power budget for the OTV during the translunar phase in Table
6.3.7-1.
The largest contributing factor to the OTV budget still arises from the
propulsion group’s electric thruster. This device consumes over 75% of the entire OTV
power budget. To adequately meet all of the power needs, we increase the total power
production by 5%.
Author: Ian Meginnis
Mission Configuration – 10kg Payload – Lunar Transfer
Section 6.3.7, Page 149
Table 6.3.7-1 10kg Payload OTV Power Budget by Group
Group
Propulsion
Communication
Attitude
Power
Lunar Lander
(during Lunar Transfer)
TOTAL
Power
2029
0
145.4
120
Units
Watts
Watts
Watts
Watts
105
Watts
2534
Watts
Solar Array Sizing
With the power budget, we determine the sizes of the solar arrays. Since the solar arrays
for the 10kg payload case are merely scaled up, we employ the sizing methods from the
100g payload case. We see the mass, stowage volume, and deployed area for the two
solar arrays in Table 6.3.7-2.
Table 6.3.7-2 10kg Payload OTV Solar Array Dimensions (total)
Parameter
Mass
Stowage Volume
Deployed Area
Value
16.89
0.0858
8.45
Units
kg
m3
m2
The manner in which the solar arrays are stored and deployed is identical to the 100g
payload case. To track the sun during the translunar phase, we include two motors and a
sun sensor.
Similar to the 100g payload case, one of the largest driving factors in determining the
OTV’s mission configuration for the 10kg case is the high cost of the solar arrays. The
solar arrays cost a total of $2.53 million.
Author: Ian Meginnis
Mission Configuration – 10kg Payload – Lunar Transfer
Section 6.3.7, Page 150
Battery Sizing
The lithium-ion battery cells for the 100g payload case are also employed for the 10kg
case. We see the new dimensions of the battery in Table 6.3.7-3. Although the OTV’s
battery still comprises of the same types of lithium-ion cells, we have 13 cells instead of
ten cells. This results from the increased battery energy capacity.
Table 6.3.7-3 10kg Payload OTV Battery Dimensions
Parameter
Mass (includes housing)
Volume
Total Energy
Value
15.93
0.0056
2008
Units
kg
m3
W-hr
The components for the OTV’s power system for the 100g payload case are also used for
the 10kg case.
We reference Fig. 5.3.7-2 in Section 5.3.7 for a diagram of the OTV
power system components. Table 6.3.7-4 shows the dimension of the 10kg payload
power system components.
Table 6.3.7-4 10kg Payload OTV Power Subsystem Dimensions
Component
Solar Arrays
Battery
PCDU
DC/DC Converters
Mass (kg)
16.89
15.93
11.49
1
Volume (m3)
0.0858 (stowed)
0.0056
0.0203
0.000218
Author: Ian Meginnis
Mission Configuration – 10g Payload – Lunar Descent
Section 6.4.1, Page 151
6.4 – Lunar Descent
6.4.1 – Structures
CAD
We see in the following figures a CAD model created in the program CATIA. This model
gives us a visual representation of what our Lander looks like, and gives us a general idea
of where certain systems are located. The Lander assembly is designed with a color
coding system in mind to easily identify certain subsystems.
7) Red
–
Structures
8) Orange
–
Attitude
9) Yellow
–
Power
10) Green
–
Communication
11) Blue
–
Propulsion
Author: Joshua Elmshaeuser
Mission Configuration – 10g Payload – Lunar Descent
Fig. 6.4.1-1 External View of the Lunar Lander with legs deployed.
(Joshua Elmshaeuser)
Author: Joshua Elmshaeuser
Section 6.4.1, Page 152
Mission Configuration – 10g Payload – Lunar Descent
Section 6.4.1, Page 153
Fig. 6.4.1-2 Close up of the Lunar Lander’s internal components with legs stowed.
(Joshua Elmshaeuser)
Author: Joshua Elmshaeuser
Mission Configuration – 10g Payload – Lunar Descent
Fig. 6.4.1-3 View of the Lunar Lander’s engine systems.
(Joshua Elmshaeuser)
Author: Joshua Elmshaeuser
Section 6.4.1, Page 154
Mission Configuration – 10g Payload – Lunar Descent
Fig. 6.4.1-4 Visualization of the 10kg Lander on the Moon.
(Joshua Elmshaeuser)
Author: Joshua Elmshaeuser
Section 6.4.1, Page 155
Mission Configuration – 10g Payload – Lunar Descent
Section 6.4.1, Page 156
Structure
The most important consideration in designing our structure of the Lunar Lander for the
10kg payload is the mass of the frame. We identify the size and mass of the Lunar
Lander at the first lunar touchdown as driving factors for the total mass of the Lunar
Lander frame structure.
The two additional engines needed to do the hop for the 10kg payload case increase the
size of the Lunar Lander frame. The height of the chambers for the hop engines is larger
than the height of the chamber for the main engine. The height of the Lunar Lander is
increased from the 100g payload case partially because of the large chamber of the hop
engines. However, the main reason for designing a Lunar Lander frame height of 1.1m
(as depicted in Table 6.4.1-1) is because the line of fire for the hopper thrusters needs to
be through the center of mass at initial lunar touchdown. The height increase of the
Lunar Lander to 1.1 meters allows for the placement of several Lunar Lander
subcomponents higher up in the frame. This increases the height of the center of mass for
the Lunar Lander. Since the center of mass is higher, the angle at which these thrusters
fire is smaller which results in the need for propellant. This results in the reduction of the
entire Lunar Lander mass including the frame. The increase of the Lunar Lander height
brings the total volume of the Lunar Lander to 1.15 m3.
Table 6.4.1-1 Basic Frame Dimensions for 10kg payload Lunar Lander
Variable
Height
Bottom Diameter
Top Diameter
Length of Legs
Value
1.1
1.3
1.0
0.606
Units
meters
meters
meters
meters
The basic shape of the Lunar Lander for the 10kg payload case is a conic frustum.
Figure 6.4.1-1 illustrates the main frame components for this Lunar Lander. The frame
floor consists of a circular outer ring and circular inner ring connected by four rectangular
Author: Ryan Nelson
Mission Configuration – 10g Payload – Lunar Descent
Section 6.4.1, Page 157
floor supports. Another circular ring makes up the top of the Lunar Lander frame. This
top circular ring connects to the frame floor by four hollow circular side support beams.
These side support beams are able to support the various loads subjected to the Lunar
Lander throughout the mission while maintaining a low thickness. A low thickness
enables the storage of the four Lunar Lander legs within these side supports during Earth
launch and Lunar Transfer. A 0.5mm magnesium skin around the entire Lunar Lander
frame protects against micrometeorites and provides thermal protection.
The lower
density of magnesium than aluminum makes magnesium more attractive for the Lunar
Lander skin.
Fig 6.4.1-1 Schematic of Basic Lunar Lander Frame Components.
(Ryan Nelson)
As depicted in Fig. 6.4.1-1, the addition of the two engines needed to make the hop
results in some extra structural mass to connect these engines. The mass needed to
integrate the two engines to the main frame of the Lunar Lander is about 5 kg. The
addition of this extra structural mass for engine integration combined with the total mass
of the Lunar Lander at initial touchdown brings the frame mass to 19.97 kg.
Author: Ryan Nelson
Mission Configuration – 10g Payload – Lunar Descent
Section 6.4.1, Page 158
Material Selection
The 10kg Lunar Lander is made of the same materials that the 100 gram Lunar Lander is
fabricated from. The floor and shear panels are fabricated from magnesium AZ31, while
all other structural members are made of aluminum 6065. All bolts, rivets, fasteners, and
connectors are made of AIS 1015 low carbon steel.
Author: Korey LeMond
Mission Configuration – 10kg Payload – Lunar Descent
Section 6.4.2, Page 159
6.4.2 – Mission Operations
Overview
In Section 5.4.2, we described the descent profile of the Lunar Lander for the 100g
payload. The descent profile of the 10 kg payload is very similar. Differences arise in
the thrust of the descent main engine, burn times, amount of propellant used, and perilune
altitude where final descent begins. In this case, descent begins at an altitude of 13.8 km.
Tables 6.4.2-1 and 6.4.2-2 highlight the additional differences for the 10 kg payload.
Table 6.4.2-1 Final Descent Action Breakdown
Action
Radial Burn
Rotation
Vertical
Coast to Land
Time (s)
000 - 177
177 - 183
183 - 199
199 - 210
Thrust (N)
1650
165
1650
180
Table 6.4.2-2 Descent Propellant Masses
Action
Time (s)
De-Circularize
Radial Burn
Rotation
Vertical
Coast to Land
Unusable
Contingency
Total
000 - 177
177 - 183
183 - 199
199 - 210
Propellant
Mass (kg)
0.5
97.8
0.3
8.5
0.9
3.5
1.0
112.5
Footnote:
An additional 8.7 kg of propellant is needed for
the hop, bringing the total propellant mass in
LPO to 121.2 kg.
Author: Kara Akgulian
Mission Configuration – 10kg Payload – Lunar Descent
Section 6.4.2, Page 160
Landing Location
We chose to land at the same landing site as the 100g payload case. Please refer to
Section 5.4.2.
Author: Kara Akgulian
Mission Configuration – 10kg Payload – Lunar Descent
Section 6.4.3, Page 161
6.4.3 – Propulsion
Overview
The Lunar Lander main engine we designed for the 10kg payload is of the same type and
configuration as the engine described in Section 5.4.3 for the 100 g payload case. Fig.
6.4.3-1 shows the engine dimensions resulting from higher thrust requirements and
includes the 100g payload Lunar Lander engine for a size reference.
Fig. 6.4.3-1 Main Engine Dimensions.
(Thaddaeus Halsmer)
Performance/Operating Parameters
When we scale the main engine for the 10kg payload case, only the thrust and burn time
change to the specifications given in Table 6.4.3-1. Isp, chamber pressure and nozzle area
ratio are the same as given in Section 5.4.3.
Author: Thaddaeus Halsmer
Mission Configuration – 10kg Payload – Lunar Descent
Section 6.4.3, Page 162
Table 6.4.3-1 Hybrid Engine Performance Specifications
Parameter
Thrust Max / Min
Burn Time
Specification
1650 / 165
190.4
Units
[N]
[s]
Propellant Feed System
The propellant feed system operates in the same manner we discussed in Section 5.4.3,
although three isolation solenoid valves were made necessary by adding the two
locomotion engines that are discussed in Section 6.5.1. The isolation solenoid valves
allow any engine to be purged and fired individually using the common feed system and
oxidizer flow control valve. The fluid system schematic given in Fig. 6.4.3-2 shows the
integration of the additional valves and engines.
Fig. 6.4.3-2 Lander Propulsion Fluid Schematic
(Thaddaeus Halsmer)
Author: Thaddaeus Halsmer
Mission Configuration – 10kg Payload – Lunar Descent
Section 6.4.3, Page 163
Table 6.4.3-2 Additional Fluid control components not in Table 5.4.3-2
Item
SV03
SV04
SV05
Description
Locomotion engine #1 ISO Solenoid Valve
Main Engine ISO Valve
Locomotion engine #2 ISO Solenoid Valve
Author: Thaddaeus Halsmer
Mission Configuration – 10kg Payload – Lunar Descent
Section 6.4.3, Page 164
Power Requirements (during operation)
Adding the three solenoid valves to the propellant feed system increases the maximum
power requirement to 150 watts.
H2O2 Storage
Employing the same technique as in Section 5.4.3, we determine the amount of H2O2
decomposed over the period of one year. The hydrogen peroxide is stored in a sealed
spherical Aluminum 1060 tank. The amount of H2O2 decomposed during the mission is
0.121 kg. As compared to the 121 kg of total propellant carried, this value is negligible.
The hydrogen peroxide tank is sized based on the tank operating pressure and the amount
of H2O2 used during the mission. Table 6.4.3-3 describes the pressure, size and mass of
the H2O2 tank.
Table 6.4.3-3 Pressure, Size and Mass of the H2O2 tanks
Parameter
Operating Pressure
Volume
Thickness
Diameter
Mass
Value
3.07
0.0867
1.26
0.550
3.35
Units
MPa
m3
mm
m
kg
The total volume of the oxidizer tank is the sum of the volume of the H2O2 required for
the mission (0.0834 m3: including the usable propellant volume for: main engine +
hopper engine, boil-off volume and the volume of unusable propellant) and a 4% ullage
volume (0.00333 m3).
Pressure Feed System
Helium is used as the pressurant gas for the pressure-fed system. We chose Titanium as
the material for the He tanks (Humble). Table 6.4.3-4 shows the detailed specifications
of the pressurant tank as well as the mass of the Helium gas required of the whole landing
mission.
Author: Saad Tanvir
Mission Configuration – 10kg Payload – Lunar Descent
Section 6.4.3, Page 165
Table 6.4.3-4 Pressurant System parameters
Parameter
Operating Pressure
Volume
Thickness
Diameter
Mass of the Tank
Mass of the Helium
Value
21
0.0397
6.70
0.41
1.34
1.37
Units
MPa
m3
mm
m
kg
kg
Propulsion System Inert Mass
The inert mass of the propulsion system is obtained as described in Section 5.4.3. The
inert masses for the additional locomotion engines are largely dependent on the amount
of propellant we use during the locomotion (hop) and the average thrust the engine
produces during locomotion. Using the propellant masses and the average thrust obtained
from mission operations we size the additional nozzles, chambers and feed system for the
two locomotion engines. The oxidizer (H2O2) is obtained from the same tank that
provides H2O2 to the main Lunar Lander engine. The total propulsion system inert mass
for the hybrid landing engine plus the two additional locomotion engines equals 45.67 kg.
Table 6.4.3-5 gives a detailed breakdown of the propulsion system inert masses.
Author: Saad Tanvir
Mission Configuration – 10kg Payload – Lunar Descent
Section 6.4.3, Page 166
Table 6.4.3-5 Propulsion system inert mass breakdown
Parameter
Main Lander Engine
Oxidizer (H2O2) Tank
Pressurant Gas (He)
Pressurant Tank (He)
Nozzle
Combustion Chamber
Injector
Skirts and Bosses
Feed System
Structural Supports
Valves
Hopper Engines
Nozzles
Combustion Chambers
Injectors
Skirts and Bosses
Feed Systems
Structural Supports
Valves
Propulsion System Inert Mass
Value
Units
3.35
1.37
1.34
15.74
3.30
0.99
0.33
0.99
2.74
6.94
kg
kg
kg
kg
kg
kg
kg
kg
kg
kg
6.30
0.29
0.09
0.03
0.09
0.68
1.10
45.67
kg
kg
kg
kg
kg
kg
kg
kg
Thermal Analysis
As in Section 5.4.3, similar thermal analysis is performed for this case. We find that to
keep the H2O2 at 283 K temperature during the lunar transfer phase we require a constant
power supply of 37 W.
During lunar descent, the temperature drop in the hydrogen peroxide tank is 5.8 K (from
283 K to 277.2 K). Similar temperature drop from the valves and the feed lines are 2.2K
and 2.0K respectively. Because the drop in temperature of these components during lunar
descent is small, its affect on engine performance is negligible. Hence, there is no power
required for thermal control on the propulsion system during lunar descent.
Author: Saad Tanvir
Mission Configuration – 10kg Payload – Lunar Descent
Section 6.4.4, Page 167
6.4.4 – Attitude Control
The attitude control system for our10kg payload case is identical to the system we use on
the lunar descent vehicle in the 100 gram payload case. For both payload cases the
attitude control system is used to counteract the moment that the lunar descent engine
creates. The main engine creates a torque on the spacecraft when the line of action of the
thrust does not pass through the center of mass of the descent module. Therefore we use
the same attitude system for both cases due to the similarities in the amount of thrust
generated by the lunar descent main engine. We also use the same attitude detection
devices from the previous case for the 10kg payload case. The sun sensors and star
sensors are located on the lunar descent vehicle for the duration of the mission and are
used for both the orbital transfer phase and the lunar descent phase. The configuration of
the thrusters on the lunar descent vehicle is found in the attitude control section for the
100 gram payload. The amount of attitude propellant is the only portion of the attitude
control system that changes between the 100 gram and 10kg payload cases. The total
mass for the attitude control system masses can be seen in Table 6.4.4-1.
Table 6.4.4-1 Attitude System Masses
Attitude Component
De-Orbit Burn Attitude Propellant
Descent Attitude Propellant
Inert Attitude Control
Total
Author: Josh Lukasak
Mass (kg)
0.5
5.13
10.18
15.71
Mission Configuration – 10kg Payload – Lunar Descent
Section 6.4.5, Page 168
6.4.5 – Communication
Communication Hardware/Configuration
The communication equipment we are using on the Lander for the 10kg payload mirrors
the 100g payload very closely, which can be found in Section 5.4.5. However, the 10 kg
payload differs from the 100g payload in one place. While we need two communication
systems for the 100g payload we now only need one system because the Space Balls are
not used. Table 6.4.5-1 lists the total mass, power usage and purchase price for the
system.
Table 6.4.5-1 Communication Equipment Totals
Mass (kg)
2.17
Power Usage (W)
54.53
Purchase Price (2009 $)
295,167
The communication systems consist of:
1) Two antennae to communicate with Earth.
2) Two antennae mount pivots.
3) A transmitter to transmit to Earth.
4) A receiver to receive signals from Earth
5) A video camcorder to record video and take pictures
6) A Computer board to control these systems.
Communication Power Requirements
Please refer to Section 5.4.6 for details regarding the communications power
requirements for the Lunar Lander/Hopper.
Communication Timeline (and completion of GLXP requirements)
Please see Section 5.4.5 for the communication timeline and GLXP requirements for the
10kg payload case, as they have gone unchanged and the way that we will achieve them
is unchanged as well.
Author: Trenten Muller
Mission Configuration – 10kg Payload – Lunar Descent
Section 6.4.6, Page 169
6.4.6 – Thermal Control
Lunar Day
The thermal control for the 10kg payload case is similar to the thermal control for the
100g payload case. We describe the main system in the 100g payload case, Section 5.4.6.
This section describes the differences between the two payload cases. Fig. 6.4.6-1 is the
diagram of the day thermal control system.
Fig. 6.4.6-1 Schematic of Passive Thermal Cooling System.
(Josh Elmshaeuser)
We use the multi-layer insulation (MLI) blanket on the entire outside of the Lunar
Lander, like the 100g payload case. The difference between the two cases is the
locomotion mode. The MLI blanket needs to cover the hopper engines. Also, the hopper
Author: Kelly Leffel
Mission Configuration – 10kg Payload – Lunar Descent
Section 6.4.6, Page 170
engines will need a resistance heater. The total system needs 9.5 Watts during Lunar
Transfer.
The other sections of the Lunar Lander thermal control will be the same. The length of
the heat pipe remains at five feet, and the communication plate remains the same size. All
the other thermal control devices scale based on the size of the Lunar Lander. A
breakdown of the subsystem masses are in Table 6.4.6-1. The total mass of the system is
9.56 kilograms.
The total mass was calculated in the MATLAB program
LanderThermalControl.m.
Table 6.4.6-1 Day Thermal Control of Lunar Lander
Components
MLI Blanket
Aluminum Plate
Heat Pipe
Radiators
Ammonia
Heaters
Total
Mass (kg)
2.38
1.40
2.51
2.80
0.0215
0.45
9.56
Surviving the Lunar Night
We will encounter the same environmental conditions for the 10 kilogram payload
configuration as with the 100 gram payload. The most pronounced difference is the
slightly greater surface area of the protected enclosure so that the heat loss is
approximately 11.5 Watts instead of 10 Watts
We are using the same heating and shutdown system as described in Section 5.4.6. The
only difference is the amount of hydrazine is increased by about 5% as is the tank volume
to compensate for the increased heat loss.
Author: Kelly Leffel
Mission configuration – 10 kg Payload – Lunar Descent
Section 6.4.7, Page 171
6.4.7 – Power
Battery
We size the battery for the Lunar Lander based on power requirements from the Lander
engine, communication and attitude systems. The LL engine, the communication gear
and the attitude systems all depend on the battery during the landing phase.
The LL’s engine requires power to operate the various valves and pumps to make the
engine function, which total 148 Watts and will be operating for 450 seconds, the first
250 seconds is for the landing phase, then the engine is ignited again for 200 seconds
after landing and this is the hopping phase, where the LL travels 500 meters and then
lands again.
The communication gear uses a peak total power of 54.53 Watts, to ensure that we can
maintain constant communication between the Earth and the LL throughout the entire
landing phase. We provide enough power for the communication gear to transmit for 30
minutes at peak power consumption. The attitude system uses a peak total power of 25.4
Watts to power our system and requires 15 minutes of peak power to make sure the LL
arrives at its destination. Table 6.4.7-1 highlights the total power requirement for all
systems and their individual components on the 100 g Lander.
Table 6.4.7-1 Power breakdown of system on Lander
Item
Propulsion
Communication
Attitude
Total
Power
148
54.53
25.4
227.93
Units
Watts
Watts
Watts
Watts
Footnote:
Complete description of each device can be found in
Section 5.4.2
For a complete background on each device within each system, please refer to Tables A6.4.7-1 through Tables A-6.4.7-3 in Section A-6.4.7.
Author: Adham Fakhry
Mission configuration – 10 kg Payload – Lunar Descent
Section 6.4.7, Page 172
We purchase the batteries from the company Yardney Technical Products, Inc., who
provide readymade lithium ion batteries that are used in space vehicles.
We select a 21 Ampere-hour battery, but the battery only requires 16.95 Ah. We decided
that as a safety precaution the battery should have a little bit extra power, in case of any
problems that might occur after separating from the OTV. Refer to Section A-5.4.7 to see
how we calculated the capacity of the battery in Ah.
In order to avoid loss of charge when not in use, which happens with all lithium ion
batteries, our battery on the LL will be completely drained during integration with the
entire spacecraft. The team will send a command that will allow the battery to charge
when it approaches the moon, using the OTV solar cells before separation. We compute
the battery size using lander10kg.m.
Solar Array Sizing
We use solar cells to provide enough power to the communication system during the twoweek long lunar day. The peak power required is 54.53 W to run all of the
communication equipment, refer to Section A-6.4.7 to see breakdown of power
requirements for each device within the communication equipment.
The design for the solar cells for the 10 kg case is the same as the 100 g case, since the
power requirements decreases by 6 watts and the solar arrays are designed to provide
extra power, we decided to keep solar arrays same size as the 100g LL. Refer to Section
5.4.7 for a detailed analysis of the solar cells. Figure 6.4.7-1 displays the power
consumed and the power generated by the cells through a lunar day.
Author: Adham Fakhry
Mission configuration – 10 kg Payload – Lunar Descent
Section 6.4.7, Page 173
Fig. 6.4.7-1 Maximum Potential Power of solar cells during a lunar
day. The communication gear has enough power to meet its peak
power requirements on the 1.5st day of the lunar day. At maximum
power, the solar cells provide 74.5% more power than required at
day 7 of the lunar day.
(Adham Fakhry)
The red line represents the power used by the communication gear, the blue line
represents the power produced by the cells. The solar cells provide more than enough
power required by the communication gear. Also, the excess power ensures the dust
accumulated on the lunar surface or impacts from meteorites, will not severely affect the
power collected by the cells as they provide extra power. We sized the solar cells to meet
the communication system’s power requirements during data transmission from the Earth
to the Moon.
The solar cells are provided by Able Engineering, who will be supplying solar cells for
the Orbital Transfer Vehicle and the Lander. The Lander solar cells will cost $250,000
and the solar cells will also weigh a maximum of 2 kilograms. Solar array sizing is
Author: Adham Fakhry
Mission configuration – 10 kg Payload – Lunar Descent
Section 6.4.7, Page 174
automated in lander.m. Refer to Section A-5.4.7 for breakdown of costs and the
constants used to size the solar cells.
DC-DC Converters
The DC-DC converters for the 10kg case depend on the number of component within the
system and the voltage they each require to operate, the DC-DC converters change the
voltage from the battery and the solar cells to a higher voltage for each individual device
will require. Tables A-6.4.7-1, A-6.4.7-2 and A-6.4.7-3 in Section A-6.4.7 refer to the
various components that require a DC-DC converter. We choose DC-DC converters for
each device by using the data in Tables A-5.4.7-6 through A-5.4.7-10, which list all the
DC-DC converters we are considering to use. Table 6.4.7-2 below displays the DC-DC
converter system for the 10 kg payload.
Table 6.4.7-2 DC-DC converter system
Component
Mass
Dimensions
Temperature Range
Cost
Quantity
0.815
0.033 x 0.033 x 0.033
-50 to 150
$ 57,500
Units
kg
m
C
-
Footnote:
Cost is in 2009 US dollars and is the cost for all the units.
All the converters will be placed in a single aluminum box that will be connected to the
PCDU (power conditioning and distribution unit) that will then branch out to each system
and their various components.
Power Conditioning and Distribution Unit (PCDU)
A power conditioning and distribution unit (PCDU) is needed to safely manage the power
distribution throughout the LL during mission. Please refer to Section 5.4.7 for the
complete analysis of why we need a PCDU, what the unit is used for and its
specifications.
Author: Adham Fakhry
Mission Configuration – 10kg Payload – Locomotion
Section 6.5.1, Page 175
6.5 - Locomotion
6.5.1 - Trajectory
The requirements for the 10kg payload locomotion phase are to travel 500m on or over
the surface of the Moon using a propulsion system on the Lunar Lander to perform a
“hop”. The Lunar Lander begins and ends the trajectory in an upright position and uses
the attitude control system to change orientation during flight.
Throttling of the
propulsion system will be used throughout the flight to make the trajectory more efficient
and ensure that the Lunar Lander touches down safely.
In order to keep the overall 90% mission probability of success, the main lunar descent
engine cannot perform the hop. Instead, one hybrid thruster orientated at 32.3° from the
Lunar Lander’s axis of symmetry provides the necessary thrust to perform the hop. This
thruster is one of a pair needed to have the necessary redundancy in the propulsion
system. This thruster configuration has a significant effect on the trajectory as the Lunar
Lander will need near zero horizontal velocity in order to land. In order to achieve this,
the Lunar Lander will have a horizontal velocity in a direction opposite that of the thrust
during the last part of the hop. This increases the required propellant and the time of
flight compared to the case where a vertical thruster provides thrust for the hop.
Fig. 6.5.1-1 shows the hop trajectory as well as the orientation of the Lunar Lander at
various points in the trajectory. Table 6.5.1-1 provides the results for this trajectory.
Author: Alex Whiteman
Mission Configuration – 10kg Payload – Locomotion
Section 6.5.1, Page 176
200
Moon
Hop Trajectory
150
altitude (m)
100
50
0
Touchdown
Liftoff
-50
-100
-1000
-900
-800
-700
-600
-500
-400
range (m)
-300
-200
-100
0
Fig. 6.5.1-1 Hop trajectory with orientation visuals at various points during flight.
(Alex Whiteman)
Table 6.5.1-1 Hop Trajectory Results
Variable
Time of Flight
Propellant Used
Average Thrust
Value
134.5
8.6
197
Units
sec
kg
N
Author: Alex Whiteman
100
Mission Configuration – 10kg Payload – Locomotion
Section 6.5.2, Page 177
6.5.2 – Propulsion
Overview
We use hybrid engines to complete the locomotion requirement of 500 meters. The
locomotion engines use the same chemical propellants and share a common H2O2 feed
system with the Lunar Lander main engine. However, the locomotion engine
configuration is changed to a simple axial flow engine with a single circular port. While
the main engine and locomotion engines share a common oxidizer feed system, they can
each be fired independently of each other. Fig. 6.5.2-1 shows the configuration of the
locomotion engine along with its dimensions.
Fig. 6.5.2-1 Locomotion engine configuration and dimensions
(Thaddaeus Halsmer)
Performance/Operating Parameters
We designed the locomotion engine to provide a median thrust to weight ratio of 1 during
the locomotion maneuver. Like the main engine the locomotion engine can be throttled to
10% of maximum thrust by controlling the H2O2 mass flow rate. Because of the fuel
grain port configuration change and reduced burn time, relative to the main engine, the Isp
of the locomotion engine is higher than the that of the main engine. Table 6.5.2-1 shows
the primary performance specifications of the locomotion engine.
Author: Thaddaeus Halsmer
Mission Configuration – 10kg Payload – Locomotion
Section 6.5.2, Page 178
Table 6.5.2-1 Hybrid Locomotion Engine Performance Specifications
Parameter
Thrust (average)
Isp, vacuum average
Chamber Pressure Max / Min
Nozzle Area ratio Ae/At
Burn Time
Specification
192
330
2.1 / 0.21
100
134.5
Units
[N]
[s]
[MPa]
-[s]
Propellant Feed System
The locomotion engines use the same oxidizer feed system used for the main engine
which was described in Section 6.4.3.
Power Requirements (during operation)
The power requirements during the locomotion maneuver remain unchanged from those
described in Section 6.4.3.
Author: Thaddaeus Halsmer
Mission Configuration – 10kg Payload – Locomotion
Section 6.5.3, Page 179
6.5.3 – Attitude
The attitude control thrusters rotate the lunar descent craft and orient it in a fashion that
will allow for optimal hop thruster firing. The attitude sensing and control devices are the
same as for the lunar descent phase. These specifics can be seen in Section 6.4.5. What is
important to note is that the hydrogen peroxide thrusters will only require 0.4 kg of
propellant for the locomotion phase of this payload case. This attitude control propellant
is used not only for counteracting hopper engine torque but also to orient and rotate the
descent vehicle to apply the hopper engine thrust in the proper direction. Exact
specifications on how the attitude control thrusters are used in the locomotion trajectory
can be seen in Section 6.5.2.
Author: Josh Lukasak
Mission Configuration – 10kg Payload – Integration
Section 6.6, Page 180
6.6 – Mission Integration
Much of the integration process for the 10kg payload case was exactly the same as for the
100 gram payload case. In this case, the only differences lie in the fact that there is no
longer a Space Ball, and two extra engines had to be integrated into the OTV electronics
module. The Lunar Lander has three engines for reliability purposes that have to be
integrated onto the vehicle. We do this by moving the electronics components further
away from the axis of symmetry of the OTV to clear room for the extra nozzles.
Once this arrangement is completed, the entire set of vehicles are integrated together in
the same manner as the 100 gram case and the mission is ready to launch.
Author: Korey LeMond
Mission Configuration – 10kg Payload – Integration
Fig. 6.6-1 Final Mission Configuration.
(Korey LeMond)
Author: Korey LeMond
Section 6.6, Page 181
Mission Configuration – 10kg Payload – Risk Analysis
Section 6.7, Page 182
6.7 - Risk Analysis
The probability of success of the mission is computed by combining the probability of
success for each vehicle in the mission. We add a redundant hybrid engine to the hop
phase in order to boost the probability of success for the mission. A table describing the
probability of success for each vehicle is shown below:
Table 6.7-1 Vehicle Success Rate
Vehicle
Dnepr Launch Vehicle
Orbital Transfer Vehicle
Lunar Lander
Hop (redundant)
Mission Success Rate
Success Rate
94%
88%
88%
99%
72%
In order to satisfy the 90% mission success rate requirement, a contingency launch of a
duplicate mission is planned. This duplicate mission will be launched immediately
following any failure of the first mission.
Table 6.7-2 Mission Success Rate
Number of Missions
One
Two
Success Rate
72%
92%
Although the individual mission success rate is not acceptable, planning a contingency
launch artificially increases our overall system reliability to satisfy the mission
requirement. The necessity of a contingency launch is closely connected to total mission
cost – a lower mission success rate allows for a less costly mission. We are taking a
gamble that our first mission will work correctly and win the GLXP purse. If our first
mission fails, it is prudent to fix the problem, build, and launch a duplicate mission to win
the GLXP purse in order to reduce capital loss.
Author: Solomon Westerman
Mission Configuration – 10k Payload – Cost Analysis
Section 6.8, Page 183
6.8 - Cost Analysis
Table 6.8-1 Total Mission Cost
Expense
Dnepr Launch Vehicle
Orbital Transfer Vehicle
Lunar Lander / Hopper
Overhead
Total Mission Cost
$/kg Payload
Cost ($M)
5.77
9.52
5.71
8.58
29.6
2.96
Our costing method tends to underestimate the total mission costs. We anticipate actual
mission costs will be higher than our estimate. Table 6.8-2 tabulates net profit if we
consider the GLXP purse money.
Table 6.8-2 GLXP Purse and Relative Cost
Expense
Total Mission Cost
GLXP Purse
Net Profit
Cost ($M)
(29.6)
22.3
(7.3)
We anticipate a net loss in capital in this mission.
Author: Solomon Westerman
Mission Configuration – Large Payload
Section 7, Page 184
7 – Mission Configuration Large Payload
Author: Solomon Westerman
Mission Configuration – Large Payload – System Overview
Section 7.1, Page 185
7.1 – System Overview
We now modify our Google Lunar XPRIZE (GLXP) mission architecture to deliver an
arbitrarily large payload to the lunar surface. Our goal is to resupply an existing lunar
base. The requirements of this mission are to travel to the moon, descend to within
100 m of the surface, and then travel 500 m over the lunar surface in at least one minute.
For the large payload mission configuration, we launch aboard the Falcon 9 launch
vehicle. Our Orbit Transfer Vehicle (OTV) and Lunar Lander are based on the designs
for the GLXP size payloads, modified to accommodate the larger mass we now carry to
the Moon. Figure 7.1-1 shows an overview of our vehicle configuration for the large
payload case.
Fig. 7.1-1 Overview of mission vehicle stack for large payload
(Korey LeMond)
Author: Christine Troy
Mission Configuration – Large Payload – System Overview
Section 7.1, Page 186
Since our requirements for this mission configuration do not stipulate that we must land
on the lunar surface before beginning our locomotion phase we incorporate a hover phase
to achieve our required locomotion. The Lunar Lander descends to 100 m above the
lunar surface, comes to rest, and then travels laterally 500 m. Figure 7.1-2 shows a view
of our Lunar Lander completing its locomotion phase.
Fig. 7.1-2 Lunar Lander completing locomotion for large payload mission configuration.
(Korey LeMond/Christine Troy)
The general timeline of our large payload mission is given in Table 7.1-1. The elapsed
times in this table are referenced from the time of arrival into low lunar orbit. It takes the
combined OTV and Lunar Lander stack one year to travel from low earth orbit to low
lunar orbit. We transfer from the Earth to the Moon using an electric propulsion system
and spiral transfer orbit. Descent to near the lunar surface and hover locomotion are
accomplished with our hybrid descent engine. We land in Mare Cognitum, near the
landing site of Apollo 12.
Author: Christine Troy
Mission Configuration – Large Payload – System Overview
Section 7.1, Page 187
Table 7.1-1 Mission Timeline for large payload configuration
Elapsed Time
(ddd:hh:mm)
-365:00:00
0:00:00
0:00:01
0:00:56
0:00:59
0:01:00
Event
Launch
Arrive in Low Lunar Orbit/Separate from OTV
Transfer to Lunar Descent Transfer Orbit
Begin Final Lunar Descent burn
Come to rest 100 m above surface/begin hover locomotion
Touch down on lunar surface
Footnote: Elapsed Time given in days, hours, and minutes
Table 7.1-2 summarizes our key mass deliverables at important waypoints along our
mission timeline. The total mission cost and cost per kilogram of payload are also
included.
Table 7.1-2 Large payload system masses and cost
Parameter
Injected Mass to Low Earth Orbit
Injected Mass to Low Lunar Orbit
Mass on Lunar Surface
Payload Delivered to Lunar Surface
Total Mission Cost
Cost per kg of Payload to Moon
Mass
9953
4545
2325
1743
222.6
128,000
Author: Christine Troy
Units
kg
kg
kg
kg
Million $
$/kg
Mission Configuration – Large Payload – Launch Vehicle
Section 7.2.1, Page 188
7.2 – Launch Vehicle
7.2.1 – Launch Vehicle/Site
We select the Falcon 9 rocket to place the large payload into a low Earth orbit (LEO).
The Falcon 9 is a new reusable launch vehicle designed by Space Exploration
Technologies (SpaceX). Although the Falcon 9 has not completed any missions to date,
successful tests on the payload separation mechanism and engine configuration have been
accomplished. Additionally SpaceX has taken extra steps in the design of the Falcon 9 to
increase mission reliability. Specifically the Falcon 9 has been designed to a 1.4 factor of
safety making it safe for human flight. The first and second stages also have significantly
higher margins of safety because they are designed to be reusable. The Falcon 9 is
propelled by ten Merlin 1C liquid propellant rocket engines (nine for the first stage, one
for the second stage). SpaceX has limited the number of stages to two to minimize
separation events and possible failures.
Fig 7.2.1-1. Falcon 9 Launch Profile.
(Falcon 9 Launch Vehicle Payload User’s Guide)
Author: Zarinah Blockton
Mission Configuration – Large Payload – Launch Vehicle
Section 7.2.1, Page 189
Table 7.2.1-1 Characteristics of the Falcon 9 Launch Vehicle
Variable
Height
Diameter
Lift-off Mass
Mass to 400km orbit
Number of Stages
Orbit Inclination
Cost per kilogram to LEO
Value
55
3.66
333,400
9,953
2
28.5
3,600
Units
m
m
kg
kg
-degrees
$/kg
The Falcon 9 will launch from Space Launch Complex 40 (SLC-40) on Cape Canaveral
Air Force Station (CCAFS) in Florida, USA. Cape Canaveral is located adjacent to the
Atlantic Ocean; which ensures safe stage drops. With 55 launches to date, SLC-40 is also
the former home of the Titan IV heavy lift rockets. Launches from Cape Canaveral are
capable of having an initial orbital inclination between 28° and 57°. The launch site is
located 28° north of the equator making it a good spot to maximize launch vehicle
performance
Fig. 7.2.1-2 Map of Launch Site.
(Google Maps)
Author: Zarinah Blockton
Mission Configuration – Large Payload – Launch Vehicle
Section 7.2.2, Page 190
7.2.2 – Earth Parking Orbit
Atmospheric Drag
We analyze the effects of atmospheric drag for the large payload in the same manner as
for the 100g payload in Section 5.2.2. We reach the same conclusions and select a
parking orbit of 400 km.
Parking Orbit Selection
Similar to Section 5.2.2, we choose the lowest altitude parking orbit possible. We select
a 400 km circular parking orbit as it overcomes drag effects and yields the minimum cost.
Author: Andrew Damon
Mission Configuration – Large Payload – Launch Vehicle
Section 7.2.3, Page 191
7.2.3 – Attitude determination in LEO
This mission configuration uses the same attitude determination subsystems and
methodology as outlined in Section 5.2.3 with no changes.
Author: Kris Ezra
Mission Configuration – Large Payload – Lunar Transfer
Section 7.3.1, Page 192
7.3 – Lunar Transfer
7.3.1 – Structure
Much of the information presented in this section contains differences between Orbital
Transfer Vehicle (OTV) configurations in the 100g and Large Payload missions. For a
better explanation on various topics discussed in this section please refer back to Section
5.3.1. Refer to Section A-5.3.1 for the details and methodology used to determine our
design for the large payload mission.
Limit Loads
We obtain launch loads or limit loads from the Dnepr User’s Manual, and use them to
evaluate the capability of our spacecraft design just as in the 100g payload case.
Similarly, we only concern ourselves with the launch loads for sizing and designing our
spacecraft, because we assume they will provide an upper bound limit for our vehicle.
See Section 5.3.1 for details on how we select these loads.
Margin of Safety
We apply the same factor of safety (FS) of 1.5 when we conduct our analysis. See
Section 5.3.1 for details on how we select this factor of safety.
Configuration
The configuration and layout for the OTV is based on the same design for the 100g
payload mission. Changes in the diameter and height of the OTV are made to increase
volume.
We increase the height and diameter of the OTV to 2.58 m and 3.6 m,
respectively. We increase the volume of the OTV, so we can fit more equipment and
larger components inside. The propulsion system is much larger in this mission along
with many other mission components. Changes to individual structural members include
different cross section thicknesses and dimensions to handle larger loads produced from a
heavier payload. These changes are illustrated in further detail in Section A-7.3.1. The
Author: Tim Rebold
Mission Configuration – Large Payload – Lunar Transfer
Section 7.3.1, Page 193
main difference between the two missions comes from this change in payload. During
launch the payload produces forces on the OTV. The payload or Lunar Lander (LL) is
significantly heavier in the large payload mission. As a result the structural members
experience stronger loads from what they experienced in the 100g payload mission. We
can scale many of the components in this design because of the similarities in OTV
configuration. Below we discuss the changes to individual components and systems.
Stiffeners
The stiffeners in this mission have a larger cross section to deal with larger bending
moments and compressive loads. Unlike the previous two missions, the size of the
stiffener cross section is determined from the bending loads transferred through them.
Buckling was the limiting failure mode in the previous designs, and determined the
needed cross section properties. See Section 5.3.1 for more details on the configuration
and layout of the support assembly.
Propulsion Module
We make the members of the propulsion frame stronger and stiffer to support a heavier
propulsion system. The cross section dimensions will again increase to handle stronger
loads. The propulsion system will also have 4 additional thrusters, which will add
complexity to the support structure. We size the frame the same way we size the frames
in the other designs. We then choose to scale the frame by a factor of 4 to account for the
added complexity. The complexity arises from extra structural support needed to mount
and harness the additional thrusters into the frame securely. We finally add another 15
percent of structural mass to account for our findings in Section 5.3.1. We found that the
frame was not strong enough based on our analyses and needed to add approximately 15
percent of additional structural mass to satisfy requirements. See Section 5.3.1 for details
on the configuration and layout of the support assembly and Section A-7.3.1 for complete
details on cross section dimensions.
Author: Tim Rebold
Mission Configuration – Large Payload – Lunar Transfer
Section 7.3.1, Page 194
Electronics Module
The electronics module remains relatively similar to the other two missions with the
addition of scaling up structural members to handle stronger loads. See Section 5.3.1 for
details on the configuration and layout of the support assembly.
Skin
The skin will remain at 0.5 mm and cover the same areas of the OTV as in the previous
two missions. The increased size and volume of the OTV will drive up the surface area
of the skin. See Section 5.3.1 for details on the configuration and layout of the skin
Integration
Unlike the previous two designs, the large payload OTV has the same diameter as the
Lunar Lander. As a result, the LL skirt is not as complex. We use a scaling factor to
determine the size and mass of both Payload Attach Fitting (PAF) and LL skirt
integration components. See Section 5.3.1 for details on the configuration and layout of
the support assembly and Section A-7.3.1 for the new LL skirt and PAF masses.
Sizing
We use the same basic principles that we used to size the OTV in the 100g payload
mission for the large payload mission. See Section A-5.3.1 for complete details on this
process.
Author: Tim Rebold
Mission Configuration – Large Payload – Lunar Transfer
Section 7.3.1, Page 195
Material Selection
The large payload case orbital transfer vehicle is made of many of the same materials that
the 100 gram and 10kg OTV’s are fabricated from. The floor and shear panels are the
only difference, being fabricated from aluminum 6065 in this case for extra strength and
support. All other structural members are made of aluminum 6065 as well. All bolts,
rivets, fasteners, and connectors are made of AIS 1015 low carbon steel.
FEA Analysis
Unlike the previous missions, we cannot perform a finite element analysis (FEA) for the
large payload mission, because we do not have enough time to model a new Orbital
Transfer Vehicle (OTV). The OTV for this mission is significantly different from the
previous mission configurations. Therefore, there will be no FEA analysis for our large
payload design.
Author: Korey LeMond
Mission Configuration – Large Payload – Lunar Transfer
Section 7.3.2, Page 196
7.3.2 – Mission Operations
Trajectory
Using the analysis explained in Section A-5.3.2 we designed the path for the large
payload. Table 7.3.2-1 summarizes the parameters of the trajectory.
Table 7.3.2-1 Large Payload Trajectory Configuration
Parameter
Payload Mass (kg)
Thrust (mN)
Mass Flow Rate (mg/s)
Earth Phase Angle (deg)
Moon Phase Angle (deg)
Parking Orbit Altitude (km)
Capture Orbit Altitude (km)
Initial Mass (kg)
Flight Time (days)
Fig. 7.3.2-1 depicts the trajectory for the large payload.
Fig. 7.3.2-1 Large Payload Trajectory.
(Kara Akgulian)
Author: Kara Akgulian
Value
3542
2118
93.5
216
270
400
25
9953
365
Mission Configuration – Large Payload – Lunar Transfer
Section 7.3.2, Page 197
Lunar Capture Orbit
The lunar capture orbit altitude is 25 km. This is the same altitude used as in the 10g
payload case. Please refer to Section 5.3.2 for more information.
Trajectory Correction Maneuver
One requirement for this project includes the capability to perform a 50 m/s burn for
course correction.
Due to factors beyond the scope of this project, such as Sun
perturbations, the spacecraft will deviate from our trajectory design during actual flight.
To account for this bias, we carry additional propellant for the main engine.
The
propellant is available for making small corrections in flight as necessary. We calculate
the propellant necessary for a 50 m/s burn and record the results in Table 7.3.2-2.
Table 7.3.2-2 Correction Maneuver Configuration
Parameter
Isp (s)
mo (kg)
Propellant for Correction (kg)
Author: Levi Brown
Value
2250
9953
22.5
Mission Configuration – Large Payload – Lunar Transfer
Section 7.3.3, Page 198
7.3.3 – Propulsion on the Orbital Transfer Vehicle
Propulsion System Overview
We select a cluster of five High-Power Hall Effect Thrusters (HPHETs) as the primary
propulsion for the OTV. Each thruster is substantially larger than the one used in the
100g and 10kg missions, however its supporting propulsion system is fundamentally the
same. Figure 7.3.3-1 shows the general layout of the EP system on the Orbital Transfer
Vehicle.
PPU
FCS
Thruster
Fig. 7.3.3-1 Propulsion System general layout inside the Orbital Transfer Vehicle.
(Brad Appel)
We design the EP system to deliver cargo from Low Earth Orbit to Lower Lunar Orbit at
the cheapest price per kilogram. The main system parameters are listed in Table 7.3.3-1.
Table 7.3.3-1 Propulsion System Totals
Variable
Wet Mass
Dry Mass
Required Power
Burn time
Payload Capability
Value
3,810
520
38,773
365
4,545
Author: Brad Appel
Units
kg
kg
Watts
days
kg
Mission Configuration – Large Payload – Lunar Transfer
Section 7.3.3, Page 199
One composite overwrapped tank provides the Xenon storage for all thrusters. The Flow
Control System is slightly more sophisticated than in the other missions, since we branch
out the feed lines into a five different paths. Each branch contains essentially the same
components found in the 100g and 10 kg missions.
Each thruster is a Busek BHT-8000. We optimize the combination of specific impulse,
mass flow rate, and number of thrusters based on cost. Table 7.3.3-2 provides the
operating conditions of each of the five thrusters.
Table 7.3.3-2 Specifications for the BHT-8000 Hall Thruster
Variable
Thrust
Specific Impulse
Mass Flow Rate
Power Input
Efficiency
Mass
Value
424
2250
19.2
7,600
0.64
25
Units
mN
s
mg/s
W
-kg
By using five thrusters we also increase the system reliability. We include enough margin
in the required power so that even if one of the thrusters fails, the remaining four could
complete the mission. Also, we have the ability to apply a controlled torque to spacecraft
by throttling particular engines. Here, throttling means adjusting the power input, not the
mass flow rate.
We achieve thermal control in a similar fashion to the 100g and 10kg missions. Even
though the waste heat out of the thrusters totals to almost 14 kW, there is enough surface
area in
the bottom
skin
of the spacecraft for passive radiation
Author: Brad Appel
control.
Mission Configuration – Large Payload – Lunar Transfer
Section 7.3.4, Page 200
7.3.4 – Attitude
Attitude Control
As mentioned in Section 5.3.4, three systems are integrated to perform attitude control of
the OTV.
1. Sensors – The same set of sensors are used for all payloads: a sun sensor and star
sensor, both made by VFCT
2. Reaction Wheels – Due to the increased mass of the OTV, a larger reaction wheel
must be used to maintain attitude stability throughout the translunar phase. VFCT
manufactures the VF MR 19.6, a wheel capable of creating adequate torque for the
system.
3. Thrusters – The 𝐻2 𝑂2 thruster hardware needed is the same as the 100g and 10kg
case. An increase in inert mass does occur as shown in Table 3 below
Our three attitude subsystems: sensors, reaction wheels, and thrusters combine to create a
complete system accomplishing the mission of controlling the OTV during the trans lunar
phase. Below, Table 7.3.4-1 provides an overview of our ACS specifications for the large
payload.
Table 7.3.4-1 OTV ACS Budget for Large Payload
Device
VF STC 1 (star sensor)
VF SNS (sun sensor)
VF MR 19.6
H202 thruster
H202 Propellant
Inert Mass
Totals
Mass (kg)
6.4
0.7
42
0.36
10.6
13.84
73.9
Cost ($)
133,333
133,333
133,333
2,500
100
1,000
404,000
Power Required(W) Volume (m^3)
20.4
0.0163
5
0.0163
280
0.0258
-0.00003
---0.00957
305.4
0.068
Author: Brian Erson
Mission Configuration – Large Payload – Lunar Transfer
Section 7.3.4, Page 201
Attitude Propellant Use
The large payload case presents a significant increase in propellant mass. Although the
thruster hardware is the same, much more propellant must be used to create enough
torque for attitude stabilization. The increase in propellant yields the need for larger and
more massive tanks. This increase is not large enough to merit a change in propellant.
𝐻2 𝑂2 still fits the bill as a powerful and adequate propellant for the large payload.
Space Environment Perturbations
Reaction wheels are used to counteract the attitude perturbations due to the space
environment. Refer to Section 5.3.4 for further details.
Author: Brian Erson
Mission Configuration – Large Payload – Lunar Transfer
Section 7.3.5, Page 202
7.3.5 – Communication
Communication Hardware/Configuration
For the large payload we will be using the same systems as for the 100g payload. Refer to
Section 5.3.5 and Section 5.4.5 for more information.
Communication Link Budget
Please see Section 5.3.5 for the communication link budget during the lunar transfer. The
communication requirements and equipment are unchanged with the change in vehicle
and payload mass.
Communication Ground Stations
For the arbitrary payload the ground station usage is the same as the 100g payload and
can be found in Section 5.3.5.
Communication Antenna Pivot
The patch antenna pivot discussed in Section 5.3.5 will be used again for the large
payload case, and will be essentially unchanged.
Author: Michael Christopher
Mission Configuration – Large Payload – Lunar Transfer
Section 7.3.6, Page 203
7.3.6 – Thermal Control
Electronics Thermal Control
Similar to the 100g and 10kg payload cases, for the large case we thermally control the
electronics, the xenon tank, and the electric thruster. The electronic components put out a
total of 4189W of heat. To cool these components, we make use of the same thermal
control system as in the 100g and 10kg cases. We see the mass and volume breakdown
of the thermal control components for the electronics in Table 7.3.6-1.
Xenon Tank Thermal Control
For the large payload case, we also employ a simple wire heater to keep the xenon within
its storage temperatures.
Electric Thruster Thermal Control
The electric thruster generates approximately 16523W of heat that we need to dissipate to
keep the thruster under its maximum operating temperature of 473K (200ºC). We see the
dimensions of the electric thruster thermal control system as well as the rest of the OTV’s
thermal control system in Table 7.3.6-1.
Table 7.3.6-1 Large Payload OTV Thermal Control Dimensions
System
Electronics
Xenon Tank
Electric Thruster
Component
Ammonia
Mass (kg)
~0
Heat Pipes
15.43
Radiators
Wire Heater
Aluminum Shroud
23.4
1
N/A
Author: Ian Meginnis
Dimensions
ID = 3.04cm; OD = 3.34cm;
Length = 5m
Total Cross-Sectional Area =
17.31m2
(8 fins @ 2.163m2 each)
N/A
Mission Configuration – Large Payload – Lunar Transfer
Section 7.3.7, Page 204
7.3.7 – Power
Introduction
For the large payload case, we cannot use the Ultraflex-175 solar arrays that we
employed with the 100g and 10kg cases. We instead use rectangular solar arrays that can
deploy to a larger area.
Power Budget
We see the adjusted power budget for the OTV during the translunar phase in Table
7.3.7-1. The largest contributing factor to the OTV budget still arises from the electric
thruster. This device consumes over 90% of the entire OTV power budget. To ensure
that we meet all of the power needs, we increase the total power production by 5%.
Table 7.3.7-1 Large Payload OTV Power Budget by Group
Group
Power
Units
Propulsion
Communication
Attitude
Power
Lunar Lander
(during translunar)
TOTAL
38773
0
305.4
1731.5
Watts
Watts
Watts
Watts
105
Watts
42960
Watts
Solar Array Sizing
With the power budget, we determine the sizes of the solar arrays. For the large payload,
we use rectangular arrays. The radii of the circular arrays are limited by the length of the
vehicle, and as such, we cannot physically attach large enough circular arrays to the side
of the OTV to support the power needs. Rectangular arrays offer more area for the same
length since they can deploy out several times their length.
This OTV needs
approximately 43kW kilowatts of power, which can be provided by two arrays each with
dimensions of 5m x 14.3m. The solar arrays are composed of triple-junction gallium-
Author: Ian Meginnis
Mission Configuration – Large Payload – Lunar Transfer
Section 7.3.7, Page 205
arsenide photocells. We see the dimension of the mass and deployed area for two arrays
in Table 7.3.7-2.
Table 7.3.7-2 Large Payload OTV Solar Array Dimensions (total)
Parameter
Value
Units
Mass
286.4
kg
Deployed Area
143.2
m2
Similar to the 100g and 10kg payload cases, one of the largest driving factors in
determining the OTV’s mission configuration for the large case is the high cost of the
solar arrays. The solar arrays cost a total of $42.96 million.
Battery Sizing
For the large payload case, we make use of the same lithium-ion battery cells that we use
for the 100g and 10kg cases. We see the new dimensions of the battery in Table 7.3.7-3.
Table 7.3.7-3 Large Payload OTV Battery Dimensions
Parameter
Mass (includes housing)
Volume
Total Energy
Value
271.7
0.0957
34261
Units
kg
m3
W-hr
Although the OTV’s battery is still made of the same types of lithium-ion cells, we have
222 cells instead of ten or 13 cells. This results from the increased battery energy
capacity for the large case.
The components for the OTV’s power system for the 100g and 10kg payload cases are
also used for the large case. We reference Fig. 5.3.7-2 in section 5.3.7 for a diagram of
the OTV power system components.
Author: Ian Meginnis
Mission Configuration – Large Payload – Lunar Transfer
Section 7.3.7, Page 206
Table 7.3.7-4 Large Payload OTV Power Subsystem Dimensions
Component
Solar Arrays
Battery
PCDU
DC/DC Converters
Mass (kg)
286.4
271.7
194.8
1
Volume (m3)
5.0 (stowed)
0.0957
0.3437
0.000218
Author: Ian Meginnis
Mission Configuration – Large Payload – Lunar Descent
Section 7.4.1, Page 207
7.4 – Lunar Descent
7.4.1 – Structures
We determine the mass of the frame structure for the large payload case by scaling up the
100g payload case structure. The Lunar Lander for the large payload case scales from the
100g case because the design of the Lunar Lander does not involve two extra hopper
engines as with the 10kg payload case. These extra hopper engines force the height of
the Lunar Lander frame to be larger. The scaled size of the propellant system (main
engine nozzle, chamber and propellant tank) compares to the 100g payload case. Table
7.4.1-1 compares of the frame dimensions and scale factors between the 100g payload
frame and large payload frame. The frame mass for the large payload case is 104.86 kg.
Table 7.4.1-1 Scale up of Large Frame Dimensions from 100g case
Structural Component
Diameter of Floor Outer Ring
Diameter of Engine support Ring
Length of Rectangular Floor supports
Length of Side supports
Diameter of Top Ring
Length of Legs
100g
Dimension
1.3 m
0.315 m
0.493 m
1.011m
1.0 m
0.607 m
Large
Dimension
3.6 m
0.872 m
1.364 m
2.022 m
2.4 m
1.214 m
Scale
Factor
2.77
2.77
2.77
2
2.4
2.4
The side supports of the Lunar Lander for the large payload case scale up by a factor of 2
while the floor components of the Lunar Lander frame scale up by a factor of 2.77 as
depicted in Table 7.4.1-1. The top ring and legs increase by a factor of 2.4. The 3.6 meter
diameter Falcon 9 payload fairing constrains the Lunar Lander base. We use a scaling
factor of 2.77 to fit the base within the faring. Scaling the other aspects of the Lunar
Lander on such a large scale is unnecessary because this leads to empty space within the
Lunar Lander. As the amount of unused volume in the Lunar Lander increases, the mass
needed for thermal control also increases. With these scaling factors, the total volume of
the Lunar Lander is 14.33 m3.
Author: Ryan Nelson
Mission Configuration – Large Payload – Lunar Descent
Section 7.4.1, Page 208
It is impractical to design the basic Lunar Lander frame for the large payload case the
same as the Lunar Lander frames for the 100g or 10kg payload cases. With the previous
two cases, the small Lunar Lander frame sizes result in light frame masses without
getting too complex. Designing the frame for the large case in the same way the frames
are designed for the 10kg and 100g payload cases results in extremely thick frame
components and unnecessarily large frame masses. This is primarily because the four
rectangular beams running from the floor’s outer floor ring to the engine floor support
ring experience high bending loads. Figure 5.4.1-1 represents a schematic of this design.
In order to get a more accurate depiction of the Lunar Lander frame mass for the large
case Lander_Frame_hover.m is run using the dimensions of the 100g Lunar Lander.
This yields a mass for each of the Lunar Lander components that are then scaled by the
respective factors seen in Table 7.4.1-1.
Material Selection
The large payload case Lander is made of many of the same materials that the 100g and
10kg Landers are fabricated from. The floor and shear panels are fabricated from a
magnesium alloy, while all other structural members are made of aluminum 6065. All
bolts, rivets, fasteners, and connectors are made of AIS 1015 low carbon steel.
CAD
The CAD model for the large payload Lander is designed the same manner as the Lander
for the 10kg payload. The only differences are that the large payload Lander is scaled up
to accommodate the larger payload and the large payload Lander does not need the two
small hopper engines. If you are interested in the design and layout of the internal
systems please reference Section 6.4.1.
Author: Ryan Nelson
Mission Configuration – Large Payload – Lunar Descent
Section 7.4.2, Page 209
7.4.2 – Mission Operations
Overview
In Section 5.4.2, we described the descent profile of the Lander for the 100g payload.
The descent profile of the large payload is very similar. Differences arise in the thrust of
the descent main engine, burn times, amount of propellant used, and perilune altitude
where final descent begins. In this case, descent begins at an altitude of 12.5 km. Also
note that the Lunar Lander is brought to a stop at 100 m where the hover trajectory takes
over. Tables 7.4.2-1 and 7.4.2-2 highlight the additional differences for the 10 kg
payload.
Table 7.4.2-1 Final Descent Burn Times
Action
Radial Burn
Rotation
Vertical
Time (s)
000 - 169
169 - 174
174 - 188
Thrust (N)
34500
3450
34500
Table 7.4.2-2 Descent Propellant Masses
Action
Time (s)
De-Circularize
Radial Burn
Rotation
Vertical
Unusable
Contingency
Total
000 - 169
169 - 174
174 - 188
Propellant
Mass (kg)
9.5
1952.0
5.7
161.0
66.4
12.0
2206.6
Footnotes: An additional 79.6 kg of propellant
is needed for the hover, bringing the total
propellant mass in LPO to 2286.2 kg.
Landing Location
The landing location is in Mare Cognitum, which is the same location as the 100g
payload. To read more about the landing site please refer to Section 5.4.2.
Author: John Aitchison
Mission Configuration – Large Payload – Lunar Descent
Section 7.4.3, Page 210
7.4.3 – Propulsion
Overview
The Lunar Lander main engine we designed for the large payload case is nearly identical
to the main engines used on the two smaller payload cases described in Section 5.4.3.
The only configuration change we make, is increasing the number of fuel grain plates,
shown in Fig. 7.4.3–1.
The additional fuel grain plates allow us to optimize the
combustion chamber dimensions and corresponding mass.
Fig. 7.4.3–1 Radial flow hybrid engine dimensions and configuration.
(Thaddaeus Halsmer)
Performance/Operating Parameters
This engine is a scaled up version of the engine discussed in Section 5.4.3 and the only
performance parameters that change are thrust and burn time.
Author: Thaddaeus Halsmer
Mission Configuration – Large Payload – Lunar Descent
Section 7.4.3, Page 211
Table 7.4.3 -1 Hybrid Engine Performance Specifications
Parameter
Thrust Max / Min
Burn Time
Specification
34500 / 3450
188.9
Units
[N]
[s]
Propellant Feed System
The propellant feed system architecture is unchanged from the system described in
Section 5.4.3.
Power Requirements (during operation)
Because the propellant feed system is scaled up for this payload case, the power
requirement for the fluid control components is 275 watts during operation.
H2O2 Storage
Employing the same technique as in Section 5.4.3, we determine the amount of H2O2
decomposed over the period of one year. The hydrogen peroxide is stored in a sealed
spherical Aluminum 1060 tank. The amount of H2O2 decomposed during the mission is
2.29 kg. As compared to the 2,290 kg of total propellant carried, this value is negligible.
The hydrogen peroxide tank is sized based on the tank operating pressure and the amount
of H2O2 used during the mission. Table 7.4.3-3 describes the pressure, size and mass of
the H2O2 tank.
Table 7.4.3-3 Pressure, Size and Mass of the H2O2 tanks
Parameter
Operating Pressure
Value
3.07
Units
MPa
Volume
Thickness
Diameter
Mass
1.652
3.376
1.467
63.87
m3
mm
m
kg
Author: Thaddaeus Halsmer/Saad Tanvir
Mission Configuration – Large Payload – Lunar Descent
Section 7.4.3, Page 212
The total volume of the oxidizer tank is the sum of the volume of the H2O2 required for
the mission (1.588 m3: including the usable propellant, boil-off volume and the volume of
unusable propellant) and a 4% ullage volume (0.0635 m3).
Pressure Feed System
Helium is used as the pressurant gas for the pressure-fed system. We chose Titanium as
the material for the He tanks (Humble). Table 7.4.3-4 shows the detailed specifications
of the pressurant tank as well as the mass of the Helium gas required of the whole landing
mission.
Table 7.4.3-4 Pressurant System parameters
Parameter
Operating Pressure
Value
21
Units
MPa
Volume
Thickness
Diameter
Mass of the Tank
Mass of the Helium
0.756
1.78
1.13
25.49
26.08
m3
cm
m
kg
kg
Propulsion System Inert Mass
The inert mass of the propulsion system obtained as described in Section 5.4.3. The total
propulsion system inert mass for the hybrid landing engine equals 277.18 kg. Table 7.4.35 gives a detailed breakdown of the propulsion system inert masses.
Author: Saad Tanvir
Mission Configuration – Large Payload – Lunar Descent
Section 7.4.3, Page 213
Table 7.4.3-5 Propulsion system inert mass breakdown
Parameter
Oxidizer (H2O2) Tank
Pressurant Gas (He)
Pressurant Tank (He)
Nozzle
Combustion Chamber
Injector
Skirts and Bosses
Feed System
Structural Supports
Valves
Value
63.87
26.08
25.49
115.50
8.67
2.60
0.87
2.60
24.57
6.94
Units
kg
kg
kg
kg
kg
kg
kg
kg
kg
kg
Propulsion System Inert Mass
277.18
kg
Thermal Analysis
As in Section 5.4.3, similar thermal analysis is performed for this case. We found that to
keep the H2O2 at 283 K temperature during the lunar transfer phase we require a constant
power supply of 37 W.
During lunar descent, the temperature drop in the hydrogen peroxide tank is 2.3 K (from
283 K to 280.7 K). Similar temperature drop from the valves and the feed lines are 2.2 K
and 3.7 K respectively. Because the drop in temperature of these components during
lunar descent is small, its affect on engine performance is negligible. Hence, there is no
power required for thermal control on the propulsion system during lunar descent.
Author: Saad Tanvir
Mission Configuration – Large Payload – Lunar Descent
Section 7.4.4, Page 214
7.4.4 – Attitude Control
Much larger attitude control devices are required for this payload case due to the increase
in descent vehicle size. Due to the fact that the attitude sensing devices are not mass
dependent the same Valley Forge attitude sensing devices will be used for the large
payload case. In order to combat the larger thrust that the main lunar descent engine
creates the attitude control thrusters must be larger as well.
The attitude thrusters are aligned in the same fashion as the previous lunar descent
modules and are composed of twelve 335 N General Kinetics hydrogen peroxide attitude
control thrusters. As in the previous two cases we use the hydrogen peroxide from the
main lunar descent engine tank for attitude control (General Kinetics). The use of the
attitude control thrusters are for a brief de-orbit burn and then for the descent phase to a
height of 100 meters above the lunar surface.
The de-orbit burn using the attitude control thrusters will require 9.5 kg of hydrogen
peroxide. For the descent burn, which occurs until the vehicle is 100 meters above the
lunar surface, we use the attitude control thrusters to control the rotation of the space craft
due to torque created by main engine offset. The thrusters are oriented in the same
fashion as in the previous payload cases and are fired in opposing pairs to offset
detrimental torques and to insure there will be no translational motion from the attitude
thrusters. During the descent phase it will cost 92.9 kg of hydrogen peroxide attitude
propellant to counteract the torque we create with the lunar descent main engine.
Other mass concerns for the attitude system include the piping that we use to tap off the
main engine propellant tank and a laser altimeter used for determining the height of the
spacecraft above the lunar surface. With all components factored in the total mass of the
attitude control system is 125.38 kg for the descent portion of the large payload mission.
Author: Josh Lukasak
Mission Configuration – Large Payload – Lunar Descent
Section 7.4.5, Page 215
7.4.5 – Communication
Communication Hardware/Configuration
For the arbitrary payload the communication equipment onboard the Lander is the same
as the 10 kg payload and can be found in Section 6.4.5.
Communication Power Requirements
Please refer to Section 5.4.6 for details regarding the Communications Power
Requirements for the Lunar Lander/Hopper.
Communication Timeline (and completion of GLXP requirements)
Please see Section 5.4.5 for the communication timeline and GLXP requirements for the
large payload case, as they have gone unchanged and the way that we will achieve them
is unchanged as well.
Author: Michael Christopher
Mission Configuration – Large Payload – Lunar Descent
Section 7.4.6, Page 216
7.4.6 – Thermal Control
Lunar Day
We use the other two payload cases’ thermal control system to design the large payload
case. The difference between this case and the other two is the locomotion. The case does
not need the extra thermal control due to the Space Balls or the hopper engines. This
design includes only one main engine and oxidizer tank needing thermal control. Figure
7.4.6-1 is the diagram of the system.
Fig. 7.4.6-1 Schematic of Passive Thermal
Cooling System.
(Josh Elmshaeuser)
The differences in this payload’s thermal control system are the heat pipe’s length and the
components needing a multilayer insulation (MLI) blanket or heater. The heat pipe’s
length increase is due to the increase in the Lunar Lander’s area. The heat pipe extends
from five feet in the other cases to twelve feet in this payload case. The MLI blanket
covers the Lunar Lander, main engine, and oxidizer tank. The main engine, oxidizer tank,
and the radiators require a heater. The total power during Lunar Transfer needed for
Author: Kelly Leffel
Mission Configuration – Large Payload – Lunar Descent
Section 7.4.6, Page 217
thermal control is 21 Watts. The other systems scale with the increase in the surface area.
The breakdown of the masses is found in Table 7.4.6-1. The total mass of the thermal
control system is 56.54 kilograms. The total mass was calculated in the MATLAB
program LanderThermalControl.m.
Table 7.4.6-1 Day Thermal Control of Lunar Lander
Components
MLI Blanket
Aluminum Plate
Heat Pipe
Radiators
Ammonia
Heaters
Total
Mass (kg)
21.4
1.40
10.53
22
0.1727
1.03
56.54
Lunar Night
We will encounter the same environmental conditions for the large payload configuration
as with the 100 g payload. The most pronounced difference is the greater surface area of
the protected enclosure so that the heat loss is approximately 119 W instead of 11 W.
We are using the same heating and shutdown system as described in Section 5.4.6. The
main difference is the amount of hydrazine is increased by a factor of 11 as is the tank
volume to compensate for the increased heat loss.
Author: Kelly Leffel
Mission Configuration – Large Payload – Lunar Descent
Section 7.4.7, Page 218
7.4.7 – Power
Battery
We size the battery for the Lander based on power requirements from the Lander engine,
communication and attitude systems. The Lander engine, the communication gear and the
attitude systems all require power from the battery during the landing phase.
The Lander’s engine requires power to operate the various valves and pumps to make the
engine function. A total of 275 W are required and the engine operates for 450 seconds.
The communication gear uses a peak total power of 54.53 Watts, to ensure that we
maintain constant communication between the Earth and the Lander throughout the entire
Landing phase. We provide enough power for the communication gear to transmit for 30
minutes at peak power consumption. The attitude system uses a peak total power of 25.4
Watts to power our system and requires this peak power for 15 minutes to make sure the
Lander arrives at its destination. Table 7.4.7-1 highlights the total power requirement for
all systems and their individual components on the 100 g Lander.
Table 7.4.7-1 Power breakdown of system on Lander
Item
Propulsion
Communication
Attitude
Total
Power
275
54.53
25.4
354.93
Units
Watts
Watts
Watts
Watts
Note: Complete description of each device can be found in
Section 5.4.2
For a complete background on each device within each system, please refer to tables A6.4.7-1 through tables A-6.4.7-3 in Section A-6.4.7.
We purchase the batteries from the company Yardney Technical Products, Inc., which
provides commercial off-the-shelf lithium ion batteries that are used in space vehicles.
We select a 21 Ampere-hour battery to support the operations of the Lander.
Author: Adham Fakhry
Mission Configuration – Large Payload – Lunar Descent
Section 7.4.7, Page 219
Solar Array Sizing
We use solar cells to provide enough power to the communication system during the twoweek long lunar day. Refer to Section A-6.4.7 to see breakdown of power requirements
for each device within the communication equipment.
Because the power requirements for the Lander do not change for the large payload case,
the design and size for the solar cells is the same as for the 100 g and 10 kg cases. Figure
7.4.7-1 displays the power consumed and the power generated by the cells through a
lunar day.
Fig. 7.4.7-1 Maximum Potential Power of solar cells during a lunar Day.
The communication gear has enough power to meet its peak power
requirements on the 1.5 day of the lunar day. At maximum power, the solar
cells provide 74.5% more power than required at day 7 of the lunar day.
(Adham Fakhry)
From Fig. 7.4.7-1, the cells still provide enough power for the Lander during its entire
time on the lunar surface. Since the solar cells remain the same, the cost of the solar
arrays is the same and is listed in Section 5.4.4.
Author: Adham Fakhry
Mission Configuration – Large Payload – Lunar Descent
Section 7.4.7, Page 220
Solar array sizing is calculated in lander.m. Refer to Section A-5.4.7 for breakdown of
costs and the constants used to size the solar cells.
DC-DC Converters
Table 7.4.7-2 displays the DC-DC converter system for the large payload.
Table 7.4.7-2 DC-DC converter system
Component
Quantity
Units
0.985
kg
0.033 x 0.033 x 0.033
m
-50 to 150
C
$ 68,500
-
Mass
Dimensions
Temperature Range
Cost
Note: Cost is in 2009 US dollars and is the cost for all the units.
All the converters will be placed in a single aluminum box that will be connected to the
PCDU (power conditioning and distribution unit) that will then branch out to each system
and their various components.
Power Conditioning and Distribution Unit (PCDU)
Please refer to Section 5.4.7 for the complete analysis of why we need a PCDU, what is
used for and its specifications.
Author: Adham Fakhry
Mission Configuration – Large Payload – Locomotion
Section 7.5.1, Page 221
7.5 – Locomotion
7.5.1 – Hover Trajectory
The requirements for the arbitrary payload locomotion phase are to travel 500m on or
over the surface of the Moon using a propulsion system on the Lander to perform a
“hop”. However, for this payload the Lander does not have to touch down on the Moon’s
surface before beginning its locomotion. We choose to travel the 500m by using a
combination of the main lunar descent engine and the attitude control system to perform a
translational hover and landing. The main lunar descent engine will be throttled to control
the Lander’s vertical movement while the attitude control system will control the
Lander’s horizontal movement.
The Lander comes to a complete stop at an altitude of 100m above the lunar surface
before the hover trajectory starts. Then the attitude control system fires to move the
Lander horizontally while the main engine throttles down to begin the descent. During
the last part of the trajectory, the attitude control system fires in the opposite direction to
stop the horizontal movement while the main engine throttles up to stop the Lander’s
vertical movement. This trajectory saves propellant and time compared one where the
Lander touches down first.
Figure 7.5.1-1 shows the hover trajectory and Table 7.5.1-1 provides the results of this
trajectory.
Author: Alex Whiteman
Mission Configuration – Large Payload – Locomotion
Section 7.5.1, Page 222
150
Moon
Hover Trajectory
altitude (m)
100
50
0
-50
-700
-600
-500
-400
-300
range (m)
-200
Fig. 7.5.1-1 Lander hover trajectory.
(Alex Whiteman)
Table 7.5.1-1 Hover Trajectory Results
Variable
Value
Units
Time of Flight
Propellant Used
61.4
19.9
sec
kg
Average Thrust
1011
N
Author: Alex Whiteman
-100
0
100
Mission Configuration – Large Payload – Locomotion
Section 7.5.2, Page 223
7.5.2 – Propulsion
Overview
The Lunar Lander main engine continues to operate during the Locomotion Phase using
its variable thrust capability to perform the hover, locomotion and landing maneuvers.
The Lunar Lander main engine used for locomotion is fully described in Section 7.4.3.
Performance/Operating Parameters
For the large payload case the Lunar Lander main engine is operating continuously
throughout the Lunar Descent and Locomotion phases. Average thrust and burn time
during the Locomotion phase is given in Table 7.5.2–1.
Table 7.5.2-1 Engine Performance Specifications
Parameter
Thrust (average)
Burn Time
Specification
3954
61.4
Units
[N]
[s]
Propellant Feed System
The propellant feed system architecture is unchanged from the system described in
Section 7.4.3.
Power Requirements (during operation)
During this phase of flight the power requirement is unchanged from what was given in
Section 7.4.3.
Author: Thaddaeus Halsmer
Mission Configuration – Large Payload – Locomotion
Section 7.5.3, Page 224
7.5.3 – Attitude Control
For translating the lunar descent module the required 500 meters we will use the attitude
control thrusters. The thrusters are oriented in a fashion that allows us to maintain a hover
condition with the lunar descent main engine. It can be seen in Fig. 7.5.3-1 that the
thrusters oriented in the cluster of four and control the rotation about the b3 vector are
able to induce translational motion when fired in pairs with the thrusters on the opposite
side of the vehicle.
At the end of the 500 meter translation the control thrusters are used to rotate the craft
and orient the main engine thruster in a direction that will allow for a soft landing. The
benefit of using the attitude control thrusters for the translational motion is that it will
allow us to adhere to the requirement that the hover phase must take over one minute.
The amount of hydrogen peroxide propellant the attitude thrusters use from the time the
craft reaches its hover altitude at 100 meters until touchdown on the lunar surface is 3.6
kg.
b3
b2
Fig. 7.5.3-1 The above figure shows the orientation
of the thrusters about the lunar descent vehicle. It
should be noted that the thrusters that control the
rotation about the b3 axis are able to cause
translational motion when fired in the proper
pairs. Not to scale.
(Josh Lukasak)
Author: Josh Lukasak
Mission Configuration – Large Payload – Vehicle Integration
Section 7.6, Page 225
7.6 – Mission Integration
Much of the integration process for the large payload case would be similar to the 100
gram and 10kg payload cases. In this case, though, the integration was only broadly
touched on. We simply did not have the time to go into great depth on where things were
going to be placed in this vehicle due to the fact we only have one semester to complete
the project. With two other missions being concurrently integrated, that left little time to
accomplish great detail in integrating this mission. For the most part, components present
on the 100g and 10kg OTV’s simply scale up to accomplish the large payload mission.
With a similarly scaled up frame, plenty of room exists within the body to place the
components and still have an OTV that fits in the Falcon 9 fairing.
One notable difference in the mission integration of the large payload case is the lander
and OTV now have the same diameter, and a mating skirt is no longer required.
Once this was completed, we integrate the entire set of vehicles in the same manner as
the 100 gram case and the mission is ready to launch.
Author: Korey LeMond
Mission Configuration – Large Payload – Risk Analysis
Section 7.7, Page 226
7.7 – Risk Analysis
The probability of success of the mission is computed by combining the probability of
success for each vehicle in the mission. We compute the probability of success for the
large payload case using a slightly modified version of the method used for the GLXPsized payloads. We assume that higher-cost missions tend to have a higher probability of
success than low-budget missions. We increase the probability of success for each
system under the assumption that a larger integration expense tends to increase the
reliability of the system.
Unlike the GLXP-sized payload cases, the large payload does not include a separate
locomotion phase. Reusing the lunar lander for locomotion increases the probability of
success for the mission. A table describing the probability of success for each vehicle is
shown below.
Table 7.7-1 Vehicle Success Rate
Vehicle
Falcon 9 Launch Vehicle
Orbital Transfer Vehicle
Lunar Lander/Hover
Mission Success Rate
Success Rate
99%
96%
97%
92%
This mission success rate satisfies the 90% mission requirement in one launch. Our large
payload mission can reliably deliver payload to a lunar base.
Author: Solomon Westerman
Mission Configuration – Large Payload –Cost Analysis
Section 7.8, Page 227
7.8 – Cost Analysis
Table 7.8-1 Total Mission Cost
Expense
Falcon 9 Launch Vehicle
Orbital Transfer Vehicle
Lunar Lander / Hover Vehicle
Overhead
Total Mission Cost
$/kg Payload
Cost ($M)
38.2
106.6
28.6
49.1
222.6
$128k
Our costing method tends to underestimate the total mission costs. This total cost is
difficult to compare with historical values. Our team could not locate total mission cost
estimates for the Surveyor or Ranger programs. We believe the Surveyor program is an
ideal comparison to this mission, as their mission requirements are similar and have
comparable masses.
Due to lack of information, we are forced to make a cost comparison to the Apollo
program. The Apollo program, a manned lunar mission, is at least an order of magnitude
more complex than our unmanned mission. In addition, the mass of the Apollo descent
stage is roughly three times the mass of our descent stage, reducing the overall accuracy
of the comparison. Using the marginal cost of an Apollo mission, the cost per kilogram
of payload to the lunar surface is approximately $8M per kilogram, adjusting for
inflation. This is many times the cost of our mission. However, we do not believe this is
a fair comparison to our mission.
Author: Solomon Westerman
Alternative Designs
Section 8.1.1, Page 228
8 – Alternative Designs
Author: Solomon Westerman
Alternative Designs – Launch Vehicle
Section 8.1.1, Page 229
Placeholder for overview section?
Author: Solomon Westerman
Alternative Designs – Launch Vehicle
Section 8.1.1, Page 230
8.1 – Launch Vehicle Alternative Designs
8.1.1 – Launch Vehicle Alternatives
Prior to our final selection of launch vehicles we gave consideration to many other
vehicles. We evaluated each vehicle based on the launch cost to LEO for the smaller
payload cases and cost per kilogram to LLO for the large payload case.
Table 8.1.1-1 Small Launch Vehicle Characteristics (Futron Corporation)
Vehicle
Taurus (USA)
Start (Russia)
Falcon 1e (USA)
Kosmos 3M(Russia)
Rockot (Russia)
Long March 2C (China)
Strela (Russia)
Long March 2D (China)
Dnepr (Russia)
Payload to LEO-200 km (kg)
1380
632
1010
1500
1850
2400
1700
3500
4400
Cost per kg to LEO ($/kg)
13,768
11,687
9,100
8,667
7,297
7,031
6,300
5,000
4,800
Table 8.1.1-2 Medium Launch Vehicle Characteristics (Futron Corporation)
Vehicle
Atlas 2 (USA)
Ariane 44L (ESA)
Delta 3 (USA)
Long March 2F (China)
Long March 2E (China)
Soyuz (Russia)
Falcon 9 (USA)
Payload to LEO-200 km
(kg)
8640
10200
8291
8400
9200
7000
10454
Author: Zarinah Blockton
Cost per kg to LEO ($/kg)
11,029
11,314
10,000
6,700
5,500
5,400
3,600
Alternative Designs – Launch Vehicle
Section 8.1.1, Page 231
Table 8.1.1-3 Large Launch Vehicle Characteristics
Vehicle
Delta IV M (USA)
Delta IV M (5,2) (USA)
Long March 2C (China)
PSLV (India)
Delta IV M (5,4) (USA)
Delta IV M(4,2) (USA)
Long March 2E (China)
Delta IV H (USA)
Atlas V (USA)
Dnepr (Russia)
Ares V (USA)
Falcon 9 (USA)
Falcon 9 H (USA)
Payload to LLO (kg)
6360
7705
1517
2206
9874
8900
4466
18545
7836
2108
80564
6780
22234
Cost per kg to LLO ($/kg)
26,536
24,680
23,568
22,470
20,595
19,798
17,288
17,251
16,733
12,521
6,788
6,782
4,973
Selection
We select the Dnepr-1 for the small payload missions because it yields the lowest launch
cost. We select the Falcon 9 for the large payload mission because it yields the least
expensive cost per kilogram to LLO. Although the Falcon 9 Heavy configuration
outperforms the medium lift version, the amount of thrusters needed is not feasible for
our Orbital Transfer Vehicle.
Author: Zarinah Blockton
Alternative Designs – Lunar Transfer
Section 8.2.1, Page 232
8.2 – Lunar Transfer Alternative Designs
8.2.1 – Trajectory Alternatives
Bielliptic Transfer
While performing preliminary analysis on various methods for lunar transfer, we consider
a bielliptic transfer. A bielliptic transfer involves three impulse burns for a total transfer
of 360o. Figure 8.2.1-1 created in Satellite Toolkit (STK) shows an example of the
transfer shape.
Fig. 8.2.1-1 Example of Bielliptic Transfer.
(Levi Brown)
Burn 1 accelerates the spacecraft from the initial parking orbit (periapsis of 1st transfer
arc) to some intermediate radius r (apoapsis of 1st transfer arc). Burn 2 accelerates the
spacecraft again, so the resultant periapsis of the 2nd transfer arc equals the final orbit
radius. For a lunar transfer, the final orbit radius is the Moon’s semi-major axis of
384400 km. Burn 3 decelerates the spacecraft to circularize the orbit at the final orbit
radius.
Author: Levi Brown
Alternative Designs – Lunar Transfer
Section 8.2.1, Page 233
We perform an analysis of a bielliptic transfer and compare the results to a Hohmann
transfer. As we see in Table 8.2.1-1, the bielliptic transfer requires more ΔV, which
consequently results in a larger initial OTV mass. We conclude that for a transfer
employing chemical propulsion, a Hohmann transfer is the most cost effective method.
Table 8.2.1-1 Bielliptic vs. Hohmann Transfer Result Comparison
Earth Parking
Lunar Parking
Intermediate
Orbit
Orbit
radius
Parameter
Hohmann
km
km
km
ΔV (km/s)
4.0
200
110
1 x 106
TOF (days)
5
Author: Levi Brown
Bielliptic
4.2
81
Alternative Designs – Lunar Transfer
Section 8.2.1, Page 234
Weak-Stability Boundary
A weak-stability boundary transfer saves 25% ∆v when compared to a Hohmann
Transfer. The savings can nearly double the amount of allowable payload placed into
low lunar orbit (Belbruno). This transfer begins with the OTV exiting the Earth capture
orbit and performing a fly by maneuver with the Moon. Enough energy has been gained
from the Moon’s gravity so that the vehicle can now reach the weak-stability boundary of
the Earth. At this location the OTV performs another maneuver to travel back to the
Moon and be ballistically captured. We must align and time all bodies correctly to
successfully execute this trajectory. This analysis would be a massive project that would
span beyond the scope of this feasibility study. Because of the complexity of the weakstability boundary transfer and its highly theoretical background it does not make this a
viable option to perform.
Author: Kara Akgulian
Alternative Designs – Lunar Transfer
Section 8.2.1, Page 235
Hohmann Transfer with Lunar Capture
The first trajectory that we consider for the tranlunar phase is the well-known Hohmann
transfer. A Hohmann transfer consists of two impulsive engine burns, typically carried
out by chemical propulsion systems.
The Hohmann transfer and lunar capture is
illustrated in Fig. 8.2.1-1. We note the second ΔV is a braking maneuver to capture into a
circular lunar orbit.
ΔV2
ΔV1
Fig. 8.2.1-1 Earth-Moon Hohmann transfer generated using Satellite Tool Kit (STK).
(Andrew Damon)
ΔVpc for a simple plane change maneuver is also calculated. All ΔV values are tabulated
in Table 8.2.1-1.
Author: Andrew Damon
Alternative Designs – Lunar Transfer
Section 8.2.1, Page 236
Table 8.2.1-1 Gravitational and Orbital Parameters for the
Earth-Moon Hohmann Transfer
Variable
Value
Units
ΔV1
ΔV2
ΔVtot
ΔVpc
3.131
0.820
3.952
2.166
km/s
km/s
km/s
km/s
Author: Andrew Damon
Alternative Designs – Lunar Transfer
Section 8.2.1, Page 237
Trajectory Alternative Selection
We compare the methods described in this section to determine the optimum mission
architecture.
We find that weak stability boundary poses several problems in the design process. The
method is primarily theoretical, so easily making changes for different payloads becomes
troublesome. Based on the design cost to employ this method, we discard it.
Comparing a Hohmann and bielliptic transfer, we see that the Hohmann consistently
requires less ΔV and time. No matter the configuration, initial OTV mass is less for a
Hohmann transfer.
We compare the required initial OTV mass for a Hohmann and spiral transfer. We find
that the ΔV for a Hohmann transfer is approximately 4.0 km/s. Previously, we calculated
the initial OTV mass that results for this ΔV for different propellants. Additionally we
find the initial OTV mass taking advantage of a spiral trajectory (See table 8.2.3-1).
Table 8.2.3-1 OTV Mass Comparison
Transfer Type
Hohmann
Spiral
Initial OTV Mass (kg)
992
550
We see that the initial OTV mass is significantly less for a spiral transfer, which
consequently leads to a lower cost. We select a low thrust spiral transfer as the mission
architecture.
Author: Levi Brown
Alternative Designs – Lunar Transfer
Section 8.2.2, Page 238
8.2.2 – Propulsion Alternatives
Chemical Propulsion Alternative
In the case of chemical propellants, we size the Orbital Transfer Vehicle (OTV) using the
ideal rocket equation and historical values of finert, Isp and oxidizer to fuel ratio. A ΔV of
4 km/s is assumed for the Lunar Transfer phase. Table 8.2.2-1 shows the OTV propellant
masses and the total OTV wet masses using different chemical propellants.
Table 8.2.2-1 OTV wet masses for different chemical propellants
Propellant
OTVwet mass (kg)
OTVwet Propellant mass (kg)
LH2/LOx
992
594
LCH4/LOx
RP-1/LOx
2084
2350
1470
1681
MMH/N2O4
Solid
2514
4690
1812
3550
Based on the results shown in Table 8.2.2-1, we see that LH2/LOx is the most feasible
chemical propellant option. However, due to high complexity, cost as well as mass
compared to electrical propulsion system, the chemical option was retired.
Author: Saad Tanvir
Alternative Designs – Lunar Transfer
Section 8.2.2, Page 239
EP Gimbal
In an effort to decrease the mass of the OTV, we considered many propulsion
alternatives. A notable alternative among those is adding a main thruster gimbal. The
basic concept is to give the OTV the capability to move and point its main thruster so that
some attitude control and maneuvering can be done without engaging attitude control
thrusters. The primary benefit of this alternative is the potential to decrease the attitude
control propellant which, in turn, decreases the total mass and the total cost of the
mission. The main disadvantage of adding a gimbal mount for the main engine is that the
mount introduces extra complexity, another failure point, and adds mass. According to
Vaughan (2005), a low complexity 3-axis gimbal mount could be developed, built, and
tested in house for less than about $1000 and with about 6 kg of material. This system is
shown in Fig. 8.2.2-1 with the OTV geometry and has a maximum range of about 20
degrees.
Fig. 8.2.2-1 This is a representation of the OTV
equipped with a 3-axis gimbal mount on the main
thruster. It also includes the location of the
center of mass and critical OTV dimensions at the time
the alternative was considered.
(Kristopher Ezra)
Author: Kristopher Ezra
Alternative Designs – Lunar Transfer
Section 8.2.2, Page 240
At the stage of development when we considered this alternative, it was noted that
onboard reaction wheels (spinning wheels which orient the OTV by pointing their
angular momentum vectors in a specific way) could completely manage attitude control.
Because of this, the only remaining task is de-saturating the reaction wheels periodically
(about 6 times) throughout each day so the wheels can continue to function properly.
Since the attitude control thrusters are located on the external body of the OTV, they
necessarily have a longer moment arm and produce more torque with less thrust. It was
this distance and increased torque that showed, after some analysis, that it was more
costly in terms of mass to gimbal the main engine than to use attitude control thrusters for
de-saturation instead. In all, the de-saturation propellant mass required by the attitude
control thrusters was found to be approximately 6.3 kg whereas the de-saturation
propellant mass required by the main engine was found to be 17.9 kg (We credit Brian
Erson with the development of the desaturation propellant cost model). This large mass
difference drives our decision to fix the main engine and use attitude control thrusters to
de-saturate reaction wheels.
Author: Kristopher Ezra
Alternative Designs – Lunar Transfer
Section 8.2.2, Page 241
Lunar Transfer Propulsion Selection
Early in the design process, we see that the electric propulsion system is capable of
delivering the same payload to the moon while using less than half the mass of the
chemical propulsion system. The cost of this mass advantage is a long time of flight and a
large power requirement, but we can still accomplish the mission in less than a year. The
EP system offers further advantages because it requires fewer moving parts, takes up less
space, and is easier to obtain than the LOx/LH2 chemical propulsion system. Therefore,
we find that electric propulsion is the best choice for a cheap and simple cargo mission to
the Moon.
Author: Brad Appel
Alternative Designs – Lunar Transfer
Section 8.2.3, Page 242
8.2.3 – Attitude Alternatives
Hydrogen Peroxide vs. Hydrazine Attitude Control Thrusters
We need the OTV to have the ability to de-saturate (or de-spin) the reaction wheels
during the Lunar Transfer phase. The most common method of performing these
maneuvers is with small attitude control thrusters. These thrusters could be useful for
attitude control during lunar descent as well (about 20Nm is needed for the 100g and
10kg payloads). One of the proposed main propulsion methods is to use hydrogen
peroxide (H2O2) as an oxidizer, so it may be possible to utilize this for attitude control as
well. Valves make both hydrazine and H2O2 throttleable (Marotta, Circle Seal Controls).
As we show in tables A-8.2.3-1 and A-8.2.3-2, hydrogen peroxide thrusters are both less
massive and significantly more affordable than hydrazine thrusters (Astrium, Rocket
Research Corporation, Wernimont, Astronautix).
Electrothermal Hydrazine
A summary of the characteristics of two types of electrothermal hydrazine thrusters,
resistojets and arcjets, is shown in Table 8.2.3-1. We note that the alternative hydrazine
monopropellant system uses less than one watt per thruster. Based on cost, power
consumption, and complexity of system integration, neither of these options was selected
for our mission (Delft, Sellers, Pinero).
Table 8.2.3-1 Resistojet and Arcjet Thruster Data
Criterion
VacuumThrust
Input Power per thruster
Maximum Isp
Life Span
System Mass (4 thrusters
+ PPU)
Resistojet
0.5-1
500
300
380
Arcjet
0.1-0.23
300
500
1250
Units
N
W
sec
hours
6.5
1.22
kg
Note: PPU is Power Processing Unit; Cost is about $150,000 for either option
Author: Brittany Waletzko
Alternative Designs – Lunar Transfer
Section 8.2.3, Page 243
8.2.3 – Power Alternatives
Introduction
We consider several different options for providing power to the OTV during the Lunar
Transfer phase. Based on the power system designs of past successful spacecraft, we
analyze four alternatives: nuclear radioisotope thermoelectric generator (RTG), fuel cells,
batteries, and solar arrays.
Nuclear Power
Nuclear power on spacecraft has been used primarily on missions past the orbit of Mars
where limited solar power is available, or where there are extended periods of darkness.
Power outputs for these systems have always been less than 1 kW electrical.
Radioisotope thermal generators or RTG’s transform thermal energy from natural
radioactive decay into electricity by means of thermocouples that are only about 10%
efficient (Angelo, 1985). We would then have nine times as much waste heat to remove.
With our 2500 watt system this would mean 22,500 watts of heat energy that would have
to be removed which would require increased mass for large radiators negating any
benefit.
Fuel Cell
For long duration missions fuel cells are often used as the primary power source. They
have advantages of long working life and a high energy density with a steady output.
Disadvantages are their expense, complexity, and need for volatile cryogenic fluids.
Another possible disadvantage of using fuel cells on this mission could be maintenance.
Fuel cells have to be purged of contaminants twice daily and the products of the chemical
reactions need to be expelled as they would be dead weight (Dumoulin, 2009).
Space-rated fuel cells have chemical efficiencies of only about 10% so it is impractical to
use them for our electric propulsion system when efficiencies of greater than 25% are
obtained by using the fuel in a chemical rocket engine (Patel, 2005).
Author: Tony Cofer
Alternative Designs – Lunar Transfer
Section 8.2.3, Page 244
Primary Batteries
Primary batteries are batteries that cannot be recharged because of non-reversibility in the
chemical reaction undergone to produce electrical power. The only portion of the mission
that these types of batteries are useful for are portions of the mission that do not include
solar panels. The only phase that did not use solar panels is the locomotion phase which
contains the Space Ball and the most mature model of the Rover. All other portions of the
mission contain the need to recharge batteries and therefore primary batteries are not
considered in these phases. The criteria for selecting the Space Ball primary battery is
described in section A-5.5.5.
Secondary Batteries
Secondary batteries are batteries that have a chemical composition that can be reversed
and can be recharged after it is depleted. This makes them necessary for any portion of
the mission that uses solar panels as the primary source of power. There are several
chemical make-ups that can be recharged and are shown in table A-8.2.4-1.
Author: Jeff Knowlton
Alternative Designs – Lunar Transfer
Section 8.2.3, Page 245
Solar Array Power
One of the major advantages that solar arrays have over other forms of energy production
is their minimal storage volume. We can fold and stow solar arrays during times of
inactivity (launch) and can deploy the arrays when they are needed (translunar phase).
Current solar array technology is also very resistant to temperature changes and radiation
damage. These two characteristics allow us to stow the solar arrays outside of the
spacecraft, providing volume inside of the spacecraft for other critical systems. This also
reduces the mass of insulation we need to protect the power system from the harsh
environment of space.
Solar arrays are ideal for spacecraft whose power demands are less than 10,000W and
whose mission durations are greater than a few days (Griffin, 2004). For the 100g and
10kg payload cases, our power requirements for the OTV are less than 3kW; this makes
our design a good fit for solar arrays. In addition, because of the long duration of our
mission, we also select solar arrays for the large payload case. Solar arrays are also very
reliable due to their successful integrations with numerous past and present spacecraft.
Table 8.2.3-1 Solar Cell Technology Comparisons
Cell Type
Efficiency
W/m2
$/Watt (2008)
Maturity
Silicon
GaAs/Ge
12-14%
18-19%
150
225
112
337
In Use
In Use
GaInP2/GaAs/Ge
30%
300
337
New & In Use
Amorphous Silicon
CuInSe
8%
10%
80
135
100
100
In Development
In Development
Note: This table was developed using data from Patel
There are several different types of solar cells from which we can fabricate our solar
arrays. These options are detailed in Table 8.2.3-1. Because of their wide availability
and relatively low costs, silicon-based solar cells have been most commonly used in past
spacecraft, including the International Space Station. These types of cells are limited in
capability, however, due to the fact that they are low in efficiency, relatively
susceptibility to radiation-induced degradation, and have a higher mass per power output
Author: Ian Meginnis
Alternative Designs – Lunar Transfer
Section 8.2.3, Page 246
than other types of cells. The efficiency of silicon solar cells is around 13%, which we
see from Table 8.2.3-1 is less than half as efficient as other types of cells (Patel, 2005).
From this fact, we see that in order to equal to the power output of other types of solar
cells, the area of the silicon arrays would need to be twice as large as the area as other
more efficient cells. It is evident to us that these characteristics would increase the mass
and the subsequent launch costs associated with silicon cells. Based on the high launch
cost per kilogram for this mission, we determine that the savings in purchase costs would
be well offset by the additional launch costs from increased mass. From the drawbacks
of silicon-based solar cells, we easily determine that silicon cells, while common on past
spacecraft, are not suitable for this mission.
Based on the properties in Table 8.2.3-1, we select triple-junction gallium-arsenide solar
cells for the OTV’s solar arrays.
These cells are very efficient because they are
constructed with three layers of sunlight-absorbing materials. These materials consist of
GaInP2 (gallium-indium-phosphorus), GaAs (gallium-arsenide), and Ge (germanium).
Not only are these solar cells significantly more efficient than silicon cells, but they are
much thinner than silicon cells (~0.25mm, including protective coverglass). Galliumarsenide cells are also much more resistant to adverse changes in efficiency at higher and
lower temperatures.
The cells are also much less susceptible to radiation-induced
degradation (Patel). All of these characteristics greatly contribute to reducing the overall
required mass for the solar arrays.
From Table 8.2.3-1, we see that the largest disadvantage to the triple-junction solar cells
is the high cost per watt of power.
These cells are approximately three times as
expensive as the silicon-based cells. These costs, however, are offset by the two-fold
increase in power generating capabilities. This characteristic allows the solar arrays to be
smaller in size and, subsequently, smaller in mass than the silicon arrays. The savings in
launch costs due to this reduction in mass significantly offset the increase in price per
watt of the solar cell.
Author: Ian Meginnis
Alternative Designs – Lunar Descent
Section 8.3.1, Page 247
8.3 – Lunar Descent Alternative Designs
8.3.1 – Landing Alternatives
High Energy Tangent Landing
The High Energy Tangent Landing (HETL) alternative involved the Lander impacting
the Lunar surface with a significant portion of its orbital velocity. We wanted to explore
this design because it would lower propellant requirements for a soft landing. The Matlab
file tangent_landing.m computes the horizontal distance our Lander will slide with a
given horizontal velocity and mass. In order to compute this distance some basic
assumptions were used:
1) The Lander would not skip.
2) The Lander would not dig into the Lunar soil.
3) The Lander would not be completely destroyed upon impact.
Author: Trenten Muller
Alternative Designs – Lunar Descent
Section 8.3.1, Page 248
Figure 8.3.1-1 shows the horizontal slide distances of our Lander that were computed.
30
25
distance (km)
20
15
10
5
0
10g
15g
20g
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
horizontal velocity (km/sec)
0.8
0.9
1
Fig. 8.3.1-1 Horizontal slide distance of Lander for three landing g-force cases.
(Trenten Muller)
Besides the large slide distance one other concern came about during this analysis. With a
significant portion of its orbital velocity our Lander would more likely create a crater and
be destroyed upon impact. Because our mission involved certain goals after landing this
risk would not be acceptable.
This design was not chosen because of this risk and because of the large slide distances.
Author: Trenten Muller
Alternative Designs – Lunar Descent
Section 8.3.1, Page 249
Gas Spring Landing Gear
Our vehicle orbiting the moon has high kinetic energy.
The Lunar Lander engine
dissipates this energy during descent. Some of this energy could be stored during landing
and used for lunar surface locomotion as illustrated in Fig. 8.3.1-2. We examine gas
springs consisting of a piston and cylinder. The springs compress at touch-down and
later release the stored pressure to bounce the Lunar Lander along the lunar surface.
Preliminary analysis shows our 500 m locomotion could be achieved with bouncing.
However, gas springs have temperature and acceleration limits that are exceeded during
lunar landing (Enidine, 2009). Gas springs are unsuitable for landing gear because of
these violations.
Fig. 8.3.1-2 Schematic of Lunar Lander hopping with gas spring landing gear.
(Christine Troy)
Author: Christine Troy
Alternative Designs – Lunar Descent
Section 8.3.1, Page 250
Accordion
The communication equipment was found to take the least amount of force on the Lunar Lander,
10 g’s. We find the forces acting on a three-dimensional object may cause deformation, causing
additional stress forces within the object. Newton’s second law of momentum leads us to find
that the longer the contact time, the less the force on the object will be. Therefore, the greater the
deformation of a material, x, the less amount of force exerted on the object. Using a honeycomb
material at the bottom of the Lunar Lander would increase the amount of material will deform
while not adding as much mass if it was just solid. We can see an example of the honeycomb
core in Fig. 8.3.1-3.
Fig. 8.3.1-3 Honeycomb core schematic.
(Caitlyn McKay)
Author: Caitlyn McKay
Alternative Designs – Lunar Descent
Section 8.3.1, Page 251
The assumption for the force exerted on the Lunar Lander is there would be complete
deformation of all hollow areas in the global buckling of the core. For the following
information the Lunar Lander used in the Matlab code theAccordion.m, has a dry mass
of 230.8 kg excluding the material being deformed, radius of 1 meter, and a plate face
thickness of 0.001 meters. The results of the optimal honeycomb core type we see in
Table 8.3.1-1.
Table 8.3.1-1 Honeycomb core length and mass of material crushed at different velocities for a
force of 10 Earth G’s
Honeycomb core
3/8-5056-.0007p
1 m/s
0.0252 m
1.25 kg
5 m/s
0.1283 m
6.45 kg
10 m/s
0.512 m
25.74 kg
15 m/s
1.151 m
57.87 kg
20 m/s
2.050 m
102.84 kg
From the data we determined that the mass added to the Lunar Lander for the crushable
core to slow impact time was greater than that of the fuel that would be saved. Other
materials could have been looked into, but a concern would be that the material would
have to be strong enough to withstand the 8.5 Earth G’s of launch. The Lunar Lander
could be supported during launch, but that would add more complexity and weight to the
orbital transfer vehicle. Even if the mass of the core was not a problem, the fact that the
crushable material has to be below the engine is a nuisance, not to mention that the
material would only buckle globally once which would not work for the 10 kg payload
vehicle. It was decided that incorporating shocks into the legs of the Lunar Lander would
be simpler than dealing with these problems.
Author: Caitlyn McKay
Alternative Designs – Lunar Descent
Section 8.3.1, Page 252
Airbag Landing
We consider using inflatables or shock absorbers as an alternate landing method because
of the potential for mass savings over a soft thruster landing. We investigate an airbag
system comprised of six cylindrical airbags arranged circumferentially on the base of the
OTV, as shown in Fig. 8.3.1-4.
2.0 m
Fig. 8.3.1-4. Arrangement of six airbags on the base of the OTV.
(Andrew Damon)
We assume that the base of the OTV is 2.0 meters, and that each airbag is a cylinder of
length 0.5 meters. The airbags are evaluated under an impact of 10 times Earth gravity
acceleration (10 g’s), which is equivalent to a deceleration of 98.1 m/s2. If we assume
the mass of the combined lander and locomotion phases is 250 kg, then the force on the
airbags is equivalent to 24525 N. If this force is divided evenly among the six airbags,
each airbag must withstand a force of 4087 N. The proposed airbag would be made of
Kevlar, with a valve or flap designed to yield at a specified pressure in order to prevent
the lander from bouncing. This “burst valve” is shown on the drawing of a single airbag
in Fig. 8.3.1-5.
Author: Andrew Damon
Alternative Designs – Lunar Descent
Section 8.3.1, Page 253
Burst valve
0.3 m
0.5 m
Fig. 8.3.1-5. Side view of a single Kevlar airbag with burst valve.
(Andrew Damon)
Using a density value for Kevlar of 1440 kg/m3, we determine that an airbag thickness of
2 mm will withstand the shock of landing. The length of the airbag is 0.5 m, and the
diameter is 0.3 m.
We next determine the mass of the system in order to evaluate it against the additional
propellant mass needed for a soft landing. Six of the airbags described above will have a
total mass of 5.03 kg. In order to inflate the airbags, gas generators are necessary. The
estimated mass of gas generators needed to inflate airbags of this volume is 10 kg.
Miscellaneous structural mass necessary to contain the uninflated airbags and support the
gas generators is estimated at 5 kg.
Total Airbag System Mass: 20 kg.
Author: Andrew Damon
Alternative Designs – Lunar Descent
Section 8.3.1, Page 254
Momentum Transfer
A preliminary alternative proposed for landing a payload on the lunar surface was the
idea of momentum transfer. In this scenario, the dry OTV traveling at orbital velocity
(approximately 1.7 km/s) ejects a portion of its mass toward the moon. By conservation
of momentum or a work-energy construction, we expect that the portion of the dry OTV
containing the Lander slows by some amount to allow it to make a semi-soft landing on
the lunar surface. In this analysis it is assumed that mass is ejected in one piece by a
spring-like force and any energy lost is attributed to this ejection. Based on preliminary
analysis, it happens that this approach is entirely infeasible.
To apply a mathematical model, we assume that a reasonable collision (or explosion)
distance is on the order of 1-2 meters. This short distance implies that, to dissipate the
full orbital velocity of 1.7 km/s, the Lander must be subjected to 1x105 Earth gravity
accelerations. This high acceleration is achieved when the dry OTV ejects the most
possible mass away from the Lander and so it is the minimum value we expect to be
reached over such a short collision distance. The limiting factor is the communication
equipment which can withstand an acceleration of at most 10 Earth g’s. At this low
value, the collision/explosion required would have to act over a length of 150 km to slow
the Lander to a stop at the lunar surface. Figure 8.3.1-6 demonstrates the acceleration felt
by the Lander is subjected to versus the collision distance required. These accelerations
are tremendously high for short range collisions. Since the accelerations and forces are
probably greater than what is sustainable by the molecular bonds in the Lander, we
observe that a new approach to landing is probably necessary and discard the momentum
transfer alternative.
Author: Kristopher Ezra
Alternative Designs – Lunar Descent
5
3
Section 8.3.1, Page 255
Work/Energy Based Accelerations
x 10
2.5
Accelleration (g)
2
1.5
1
0.5
0
0.5
1
1.5
2
2.5
3
3.5
Collision Distance (m)
4
4.5
5
Fig. 8.3.1-1 Plot of the acceleration that the Lander is subjected to versus the collision distance
required to slow the Lander from 1.7 km/s to 0 km/s. These accelerations are in multiples of Earth's
gravity (g's).
(Kristopher Ezra)
Author: Kristopher Ezra
Alternative Designs – Lunar Descent
Section 8.3.1, Page 256
Semi-Soft Landing
We considered a semi-soft landing as an alternative to a soft landing to see if it would
save mass or cost on the lander.
We determined a 15g maximum loading would
constitute a semi-soft landing. Initially, we thought the benefits of this landing would be
a savings in complexity and propellant mass. However, after looking at this alternative in
more detail, we determined that the propellant savings would be at most a few kilograms
unless the lander could impact the lunar surface with a significant percentage of its orbital
velocity. This scheme, however, would require a large addition to the structural mass of
the lander in order for it to withstand such high loading. Consequently, this alternative
was discarded in favor of a soft landing.
Author: Alex Whiteman
Alternative Designs – Lunar Descent
Section 8.3.1, Page 257
Spinning Tether
An alternative explored for placing our Lander on the lunar surface was the concept of a
spinning tether. In this alternative, the Lander and dry OTV are connected by a tether
and spin about their center of mass during descent. If the system spins so that the Lander
is traveling at the orbital velocity of 1.7 km/s around the center of mass, then when the
Lander moves toward the Moon it travels with a relative speed of 3.4 km/s and when it
moves away it has zero relative speed (this only occurs when the tether is exactly
perpendicular to the velocity vector of the entire system). To capitalize on this effect, we
ignite a charge exactly when the Lander is at or near the lunar surface and has zero
relative speed. The dry OTV is lost, but we place the Lander on the surface of the Moon
with little or no propellant. Figure 8.3.1-7 illustrates the setup. For simplicity, the figure
shows the center of mass of the system to be in the center of the tether, this is an extreme
simplification and is obviously not the case in the actual analysis.
w
v
2v
v=0
Fig. 8.3.1-2 The system of the dry OTV with the Lander moves toward
the Moon at a rate v while rotating at a rate w. The goal is that the
Lander moves with a net speed of zero relative to the Moon at some
point along its trajectory.
(Kristopher Ezra)
We observe that this alternative is subject to the same constraints as the other landing
alternatives in that the maximum acceleration sustained by the Lander cannot exceed 10
Author: Kristopher Ezra
Alternative Designs – Lunar Descent
Section 8.3.1, Page 258
Earth g’s, the system must travel at 1.7 km/s initially, and for the system to be viable it
must have a mass less than the proposed propellant mass. With these constraints in mind,
we assume that the dry OTV and Lander are connected with a length of Kevlar (whose
safety factor is 1.25 in all cases observed) and use a work/force/energy model to recreate
the scenario. At the point in time this alternative was considered, the dry OTV had a
mass of 251 kg, the Lander had a mass of 138.3 kg, and the descent propellant mass was
75 kg. Fig. 8.3.1-8 shows the required length of tether for a 1.7 km/s orbital velocity is
about 50 km.
Linear Velocity vs Tether Length
2
Magnitude of Linear Velocity (km/s)
1.8
1.6
1.4
1.2
1
0.8
0.6
0.4
0.2
0
0
10
20
30
40
Tether Length (km)
50
60
Fig. 8.3.1-3: Plot of the magnitude of the orbital velocity versus the tether length required to spin the
Lander at this rate without exceeding the 10 Earth g's constraint.
(Kristopher Ezra)
Author: Kristopher Ezra
Alternative Designs – Lunar Descent
Section 8.3.1, Page 259
Given this tether length, we compute the net mass savings if the tether replaces the
descent propellant. The density of the tether remains constant and so does its cross
section (1.3 mm radius including a 1.25 safety factor), but as the length of the tether
increases so does its total mass. Fig. 8.3.1-9 shows the mass savings as a function of
tether length and it is clear to us that, even if we could support a tether with a length that
is half of the orbital height above the moon, the mass of the tether is too large. A mass
savings of -325 kg indicated on the plot corresponds to a tether with a mass 325 kg
greater than the descent propellant. This equates to a tether with a weight of 400 kg.
This is not feasible. For this reason, the spinning tether alternative was discarded.
Because of the very low accelerations sustainable by the communications equipment (10
g limit), the Lander and payload cannot be placed on the Lunar surface by means of a
spinning tether system.
Author: Kristopher Ezra
Alternative Designs – Lunar Descent
Section 8.3.1, Page 260
Mass Savings
100
50
0
Mass Savings (kg)
-50
-100
-150
-200
-250
-300
-350
-400
0
10
20
30
40
Tether Length (km)
50
60
Fig. 8.3.1-4: Total mass savings obtained when the tether replaces the descent propellant. The
function decreases with tether length and rapidly becomes negative.
(Kristopher Ezra)
Author: Kristopher Ezra
Alternative Designs – Lunar Descent
Section 8.3.1, Page 261
Alternative Selection
There are a variety of creative landing trajectories with which we could have chosen for
our mission but in the end the most elegant and safest landing trajectory is a soft landing.
Every landing trajectory explored aside from soft landing is an unproven method for
lunar landing. The lack of a lunar atmosphere makes any alternative except a soft landing
very risky and reduces our probability of mission success. The mass additions that are
required for the some of the alternative landing trajectories (spinning tether and
accordion) are simply just too costly to implement and explore. The structural mass
increases to protect the electronics on board the lunar lander during a semi-soft or hard
landing are far too large and complicated to use a landing trajectory. In summary the
reduction in mission success probably, implementation cost, and mass increases drove us
to choose the soft landing alternative for the lunar descent phase in all payload case
instances.
Author: Josh Lukasak
Alternative Designs – Lunar Descent
Section 8.3.2, Page 262
8.3.2 – Structural Alternatives
Structural Design
We use a conic frustum shape for the Lunar Lander in each of the payload cases for two
reasons.
This shape allows for the comfortable housing of all the Lunar Lander
subsystems without leaving too much unused volume. This simple design shape also
allows for easy recalculation of the frame mass as the size and mass of the Lunar Lander
subsystems change.
We freeze the conic frustum shape itself in all payload cases, but alternative design
concepts for the frame around this shape are examined.
The rectangular floor support
beams contribute the largest portion of mass to the Lunar Lander frame. The floor
consists of an outer circular ring connected to an inner engine support ring by these four
rectangular beams as depicted in Fig. 5.4.1-1. We examined the frame mass in all three
payload cases for three and five of these rectangular floor beams. We find that with three
of these rectangular floor beams each beam becomes thicker in order to support the
bending loads. The increased thickness causes an unnecessary increase in Lunar Lander
frame mass. With the examination of five rectangular floor beams, we find that although
each beam becomes thinner, the addition of the extra beam results in a higher Lunar
Lander frame mass.
A similar analysis was done on the side support beams which did not play a large roll in
the overall frame mass, but did contribute. Four side support beams are used for all three
payload cases but three and five supports are analyzed. The results compare with the
floor supports. For smaller payload masses such as the 100g and 10 kg cases, a Lunar
Lander frame with four side supports and four legs yields the lowest overall frame mass.
These alternate frame designs are not analyzed for the frame with the arbitrary payload.
Since the basic Lunar Lander frame design depicted in Figure 5.4.1-1 is impractical for
the frame with the arbitrary payload case, we used a scale up from the four side support
and four leg 100g payload Lunar Lander.
Author: Ryan Nelson
Alternative Designs – Lunar Descent
Section 8.3.3, Page 263
8.3.3 – Trajectory Alternatives
Linear Tangent Steering Law
Although our final mission descent trajectory employs a relatively simple radial/vertical
burn scheme, other descent trajectories were considered. The most promising of these
alternatives analyzed was the linear tangent steering law (LTSL), as presented to us by
James M. Longuski, Professor of Astrodynamics, Purdue University. The goal of this
section is to describe the analysis that was performed using the LTSL method, as well as
illustrate the cost-benefit conclusions of the analysis. Finally, we will explain why the
radial/vertical burn scheme was chosen for all three mission configurations.
The linear tangent steering law reduces analysis of a launch or descent to a two
dimensional problem. The variable of interest determined using the LTSL is the angle at
which a spacecraft is thrusting, here we use 𝜃, as a function of time. Figure 8.3.3-1
provides an overview of the problem; y is in the vertical direction pointing up from the
surface of the moon.
y
f
𝜃
g
x
Fig. 8.3.3-1 Overview of descent from parking
orbit to landing, which is at the origin in this
figure.
(John Aitchison)
From Professor Longuski’s presentation on the linear tangent steering low, for the Lunar
Lander we calculate the x and y components of position and velocity as a function of 𝜃,
𝜃0 , a, and f. Please see the Matlab codes LinearTangentSteeringLaw.m as well as
LinearTangentSteeringLaw_func.m for the process used to accomplish this.
Author: John Aitchison
Alternative Designs – Lunar Descent
Section 8.3.3, Page 264
After using the LTSL to determine an optimized function for theta, the amount of
propellant used was compared to that required of our radial/vertical burn scheme. In each
case, the flight times using the LTSL were much greater than those obtained using the
radial/vertical burn scheme. These longer burn times using the LTSL resulted in higher
propellant masses based on a constant mass flow assumption. The LTSL did, in cases
where we started descent from very high altitudes, result in propellant savings. These
savings were seen when descent started altitudes above approximately 40 km.
As a result of the lower propellant mass usage of the radial/vertical burn scheme, we
chose not to use the LTSL. Although the LTSL did provide propellant savings when
starting descent at high altitudes, the total propellant usage from descent at low altitudes
trumped these savings. Also, a significant note is that the LTSL analysis assumed
constant acceleration of the spacecraft. This simplification would have significantly
reduced the validity of the descent trajectory analysis. Using the radial/vertical burn
scheme, we were able to model acceleration as a function of time using precise equations
of motion, which provided a more accurate model of the Lunar Lander descent
Author: John Aitchison
Alternative Designs – Lunar Descent
Section 8.3.4, Page 265
8.3.4 – Propulsion Alternatives
Bi-Prop vs. Hybrid Engine of Lander Engine
The use of a bi-propellant (LH2/LOx) system leads to a total system mass savings of ~6
kg for the 100 g and the 10 kg payload cases. However, no existing bi-propellant engine
exists that can satisfy the thrust levels that we require. Developing a new system from
scratch is infeasible for the mission time frame and extremely costly as compared to the
radial flow hybrid engine.
Furthermore, the bi-propellant system is more complex than the currently employed
hybrid system. If we use the bi-propellant engine, we will be adding another propellant
tank, increase driving gas volume for the pressurant system, increase the number of feed
lines and double the control valves. Moreover, power required for thermal control
approximately doubles and the overall volume occupied by the propulsion system
increases, which leads to an increase in Lander mass and volume.
As a result we eliminate the bi-propellant alternative for the Lander engine for all three
payload cases.
Author: Saad Tanvir
Alternative Designs – Lunar Descent
Section 8.3.4, Page 266
Solid Rocket Motors
Solid rocket motors (SRM) were one of the propulsion design alternatives we considered
for Lunar Descent. The advantage of this approach is the elimination of the relatively
complex propellant feed system associated with liquid propellants. By using multiple
solid rocket motors fired in series, analysis of the Lunar Descent shows that a semi–soft
landing is theoretically feasible. Depending on the impulse of the engines we find that
the complete Lunar Descent mission requires a total of 5–8 engines.
ATK was the only solid rocket motor supplier found that sold the desired size of motors.
When priced we found that the purchase cost of each engine was approximately
$700,000.
Table 8.3.4–1 contains the motor types, cost and the number required to
perform the lunar descent mission using SRM’s. The initial conditions for the analysis
were based on an initial descent altitude at the Perilune of a 110x15 km altitude elliptical
orbit. Because of the prohibitively high cost of this approach, further analysis of the
SRM lunar descent design was not pursued.
Table 8.3.4 – 1 SRM purchase costs for 85kg payload lunar descent
QTY
4
3
ATK Solid Rocket Motor (STAR Series)
STAR 13b (provide impulse to zero circular velocity)
STAR 6b (provide impulse to zero vertical velocity)
Total cost for lunar descent motors 
Author: Thaddaeus Halsmer
Cost Each
$700000
$700000
$4.9 million
Alternative Designs – Lunar Descent
Section 8.3.4, Page 267
Lunar Descent Propulsion Selection
The main incentive for using solid rocket motors for Lunar Descent is their simplicity –
but at a cost of several million dollars, we find it is worth it to go for a less-simple but
cheaper system. At the other extreme, a bi-prop descent engine requires an enormous
amount of complexity – and runs into the same problem as solids with commercial
availability. Monopropellant systems do not offer the competitive specific impulse
needed for a low-mass system. We find a happy medium with a radial Hybrid engine.
The Hybrid engine is relatively simple, offers good performance, and we can develop it
for a price cheaper than the other options.
Author: Brad Appel
Alternative Designs – Lunar Descent
Section 8.3.5, Page 268
8.3.5 –Attitude Alternatives
Spin Stabilized Lunar Lander
Much like a spinning top stands stably upright, a spacecraft spun around its central axis
remains fixed in space. Since space is a very low friction environment, once a vehicle is
spun up, it remains spinning will very little or no additional force input. For these
reasons, we consider spinning the Lunar Lander for attitude control. Spin stabilization
only requires propellant for initial spin-up and for spin-axis reorientation. Figure 8.3.5-1
is a schematic of the spinning Lunar Lander.
Fig. 8.3.5-5 Diagram of spinning Lunar Lander configuration.
(Christine Troy)
We calculate the propellant needed for both the spin up and axis reorientation and
compare this value to the propellant mass needed for 3-axis control. The mass of
propellant saved is only around 2.2 kg. There is also 0.54 kg mass savings because six
thrusters are eliminated. However, to implement the spinning Lunar Lander, the descent
engine and landing gear each need extensive redesign. These redesigns add significant
mass to the systems, outweighing the mass saved in propellant. Additional attitude
sensing equipment is also needed for the spinning Lunar Lander concept. The spinning
Lander concept is more complex and riskier than 3-axis stabilization and offers no mass
savings. For the above reasons, the spinning Lunar Lander alternative design is not used.
Author: Christine Troy
Alternative Designs – Lunar Descent
Section 8.3.5, Page 269
Attitude Control-Thrusters
The two main options for lunar descent attitude propellant are hydrazine and hydrogen
peroxide. An important factor when we choose attitude control propellant is the specific
impulse of the propellant. The specific impulse of hydrazine is approximately 230
seconds (Astronautix) while hydrogen peroxide has a specific impulse of approximately
140 seconds (General Kinetics, 2009). The reason that specific impulse is such an
important factor when determining what attitude propellant can be seen here:
(Rauschenbakh, 2003)
𝑚̇ =
|𝑀|
𝑔 ∗ 𝐼𝑠𝑝 ∗ 𝐿
(8.3.5 − 1)
In Eq. 8.3.5-1 𝑚̇ is the mass flow rate in kg/s of propellant. Our goal when using this
equation is to keep 𝑚̇ to a minimum, thus reducing mass and cost. When applying this
equation to the descent phase of the10kg payload case and comparing the results when
using hydrazine or hydrogen peroxide the mass savings for using hydrazine is 2.61 kg,
which is a fairly large amount. The issue with hydrazine though is the high cost of it as a
propellant and for thrusters that employ hydrazine. Hydrazine thruster systems cost
nearly $100,000 each while hydrogen peroxide thrusters cost on the order of thousands of
dollars. This cost savings alone is worth the extra 2.61 kg we have on the system due to
our choice of hydrogen peroxide. Not only the thrusters but hydrazine itself is in the
hundreds of dollars per kilogram for only the propellant while propellant grade hydrogen
peroxide is under ten dollars per kilogram (Peroxide Propulsion, 2009). The large cost
savings brought about by using hydrogen peroxide more than offsets the slight mass
advantage we gain by using hydrazine attitude control thrusters.
Alternative Selected
For the lunar descent module and all payload cases the attitude control system we chose
was hydrogen peroxide thrusters and 3-axis stabilized landing as opposed to spin
stabilized. In summary hydrogen peroxide thrusters cause a slight increase in mass on our
lunar descent vehicle but save us hundreds of thousands of dollars in system costs. 3-axis
Author: Josh Lukasak
Alternative Designs – Lunar Descent
Section 8.3.5, Page 270
stabilized landing is a less complex attitude control solution than spin stabilization. Spin
stabilization creates a reduction in attitude propellant requirements but the mass savings
would need to be translated into extra structures to deal with the spinning descent. For the
large masses we are dealing with spin stabilization is not an advantageous solution for
attitude control.
Author: Josh Lukasak
Alternative Designs – Lunar Descent
Section 8.3.5, Page 271
Alternative Selection
In previous sections various modes of attitude control were discussed and decisions were
made based on the feasibility of implication, cost, and any possible risks involved. The
first decision made for the orbital transfer vehicle is what type of attitude control thrusters
are to be used. When looking at various types of thrusters including cold gas, electric
propulsion, and hydrazine it was apparent that hydrogen peroxide thrusters are the
optimal solution for attitude control. In general hydrogen peroxide is a relatively cheap
propellant when compared to other attitude propellants. Cold gas thruster systems
themselves are quite expensive and electric propulsion attitude thrusters require a large
amount of power which we cannot divert from the main electric propulsion engine. We
can note that having hydrogen peroxide thrusters require their own separate tanks but the
price benefits of hydrogen peroxide for attitude control far outweigh the slight mass
increase in the overall system and the performance advantages of certain attitude
thrusters.
The second alternative for attitude control is spin stabilization as opposed to a 3-axis
stabilized spacecraft. The benefit of a spin stabilized space craft is a reduction in attitude
control propellant required. We chose 3-axis stabilization for our space craft during the
lunar injection phase because it is difficult, without adding considerable structural mass,
to spin the space craft but also keep the solar panels and communication antennas
properly aligned in their respective directions.
Author: Josh Lukasak
Alternative Designs – Lunar Descent
Section 8.3.6, Page 272
8.3.6 – Thermal Control Alternatives
Thermal- Active Control
We considered a mostly active thermal control system because past space missions
include active systems. Instead of a heat pipe, we could use a cryogenic cooler could
perform the necessary task. Upon doing research several different coolers would cool the
Lunar Lander. The problems associated with the coolers included high mass and power.
For example, a NGST HEC weighs 7 kilograms by itself, additional 5 kilograms for other
thermal control. The cooler requires 120 Watts to run (Donabedian, 2004).
Another form of active control, we considered using heaters to warm the Lunar Lander
during the lunar night. The idea is not ideal for our case due to the large amount of power
that heaters require. Our system saves over 2 kg, 120 Watts, and $30,000.
Author: Kelly Leffel
Alternative Designs – Lunar Descent
Section 8.3.7, Page 273
8.3.7 – Power Alternatives
Solar Array Deployment
Two different alternatives were considered for deploying the arrays after storage in the
Lunar Lander during descent. The active system includes a solar array deployment
system. The passive system includes a solar hinge. The passive system would add over 1
kilogram in deployment and the solar array deployment system would add 2.27 kilograms
(Davis, 2009). For both of these alternatives the solar arrays would need to be protected
during descent.
The current design for the solar arrays on the Lunar Lander attaches the arrays to the top
of the Lunar Lander. We determine this simple design to be the best alternative for our
particular mission. Both these options became invalid after the decision was made to
place the solar arrays on the top of the Lunar Lander. The advantages for our decision
include less weight and power. Also the system is more reliable since the solar arrays do
not need protection during descent, unlike storing in the Lunar Lander, and the ability to
collect power without the delay between landing and deploying.
Author: Kelly Leffel
Alternative Designs – Rover
Section 8.4, Page 274
8.4 – Rover
Introduction
The Rover design, as the locomotion phase, comes from the Apollo’s Lunar Rover, Mars’
Rovers Opportunity and Spirit, and the new Rover prototype currently being designed by
NASA. Since NASA has consistently used rovers in the past, our first designs for the
locomotion stage are rovers devised for our purposes.
There were two different designs for the Rover. The first was to have a slower moving
Rover that would travel the 500 meters in 2 Earth days designed for the 100g and 10kg
payload cases. These Rovers had to be carefully thermally controlled as to not overheat.
The second Rover is designed to travel the 500 meters in 13 minutes and then overheat
for the 10kg payload case. The mass break down for the second designed rover, given by
the rover design team, we see in Table 8.4-1.
Table 8.4-1 Mass breakdown for the
second Rover design, 10kg payload case
Item
Battery
Antenna
Transmitter
CPU (computer)
Camera
Motor (4)
Mounts (8)
Gearhead (4)
Thermal Control
Magnesium Body
Space Blanket
Wheel (4)
Power (Extra)
Payload
Total
Mass in kg
0.422
0.1
0.21
0.2
1
0.092
0.176
0.03
0
0.4
0.58
0.58
1.92
10
15.71
The Rover was not chosen for either for the 100 gram or 10 kg case because at that point
in time other systems of travel were found to be lighter, which made the entire vehicle
Author: Caitlyn McKay
Alternative Designs – Rover
Section 8.4, Page 275
system less expensive. When the Hopper, the current 10kg payload design, reached the
same maturity level of the Rover the cost and mass were very similar. We can conclude
that this means that there is more than one feasible solution to the problem.
Author: Caitlyn McKay
Alternative Designs – Rover
Section 8.4.1, Page 276
8.4.1 – CAD/Integration
We modeled the small rover for the 100 g and 10 kg payloads in CATIA. Figure 8.4.1-1
shows how the rover components are assembled. The four wheel drive assemblies are
mounted on the underside of the rover with aluminum motor mounts. We place the
electrical components including the camera, battery, CPU, and transceiver, inside the
rectangular rover body. The figure however does not reflect the true size of the battery for
the faster Kamikaze rover as design worked stopped before the CAD model was updated.
CPU
Transceiver
Camera
Battery
Wheel Drive Assemblies
Fig. 8.4.1- 1 The system configuration for the small rover design.
(Ryan Lehto)
Author: Ryan Lehto
Alternative Designs – Rover
Section 8.4.2, Page 277
8.4.2 – Communications
While the communication systems for the rover were not fully developed due to the
switching to the Space Ball design, the mass and size of many of the systems could have
been reduced.
We selected larger heavier radiation hardened equipment for the Rover. We also planned
on communicating with the Earth using S-Band frequencies instead of using the lighter
UFH systems and communicating to the Lander.
We originally selected a dish antenna for communication back to Earth. As well as a
transponder and a receiver which were powerful enough the send the signal all the way
home. In the end the Rover just like the communication gear it was equipped with were
upgraded and replaced with lighter, more efficient products.
Author: Joshua Elmshaeuser
Alternative Designs – Rover
Section 8.4.3, Page 278
8.4.3 – Propulsion
System Overview
We analyzed a rover propulsion system that consisted of four individual drive assemblies
attached to the underside of the rover body seen in Fig. 8.4.1-1. Figure 8.4.3-1 shows
each drive assembly includes a D/C motor, a planetary gear head, and a wheel.
Motor
Gear
Wheel
0.076 m
Fig. 8.4.3-1 Assembly of the drive system components
the D/C motor, planetary gear head and the wheel.
(Ryan Lehto)
Placement Consequences
The placement of rover drive assembly exposes the motors and gear heads to the
harshness of the vacuum of space. The motors we use are recommended for space
applications; however the drive assemblies must be lubricated with a special space rated
lubricant that will not leak out nor freeze in the extreme hot and cold of space.
Motor and Gearing
We chose a 19 mm diameter D/C motor with an accompanying 22 mm planetary gear
head. The motors share the load of the rover mass equally. The initial rover design
moved at a velocity of 0.01 m/s. We believed a larger velocity would be too risky as the
rover could become hung-up on an obstacle or overturned. Each D/C motor connects
Author: Ryan Lehto
Alternative Designs – Rover
Section 8.4.3, Page 279
with a planetary gear head to amplify the torque provided to the wheels. All the motors
combine to produce a torque of 0.85 N-m to power the rover. The rover turns similar to a
tracked vehicle by reversing two drive assemblies on the same side. The system allows
the rover to turn 360° in place for a turn radius of 0 m.
The wheels are designed to traverse the lunar terrain and absorb vibrations during roving
phase of mission.
The space inside the wheel is filled with foam to prevent the
accumulation of lunar dust and debris. The average sized obstacle expected will be less
than 0.20 m, see Section 5.5.6. The optimum wheel size and spacing is 0.076 m 0.214 m
respectively. The curved spokes act as a dampener, absorbing the vibrations caused by
rover’s movement. No further analysis of spoke effectiveness at absorbing vibrations
was performed due to the decision to use the Space Ball concept for the small payload.
Author: Ryan Lehto
Alternative Designs – Rover
Section 8.4.4, Page 280
8.4.4 –Thermal Control
The proposed thermal control for the Rover Alternative was to use a solid copper block
heat sink. The copper heat sink would absorb heat from the electronics and transfer it to
heat pipes to conduct it to the surroundings. The primary focus would be to dissipate
heat from the processors of the electronic systems and the incoming solar energy. The
drawbacks of this system are the large weight of the copper heat sinks as well as the
inefficient dissipation.
Author: John Dixon
Alternative Designs – Rover
Section 8.4.5, Page 281
8.4.5 – Power
The power system for the rover changed several times, most recent official number use a
single battery that outputs 3.6 volts and 12 ampere-hours. The total energy contained in
the battery cell is 43.2 Watt-hours. In addition to this 0.422 kg battery cell there is
addition hardware that pushed the total mass of the power system to 2.3 kg. The
additional hardware contained a kg of battery housing and DC-DC converters to increase
the voltage to the proper levels for each electrical component.
Author: Jeff Knowlton
Alternative Designs – Rover
Section 8.4.6, Page 282
8.4.6 – Deployment
We use a linear shaped charge to deploy the Rover. When we ignite the charge there is
an explosive but straight cut through the material below. Linear shaped charges will be
made in-house using copper lining and C-4 as the explosive material (Kane, 2009).
Copper Lining
Explosive Charge
Fig. 8.4.6-1 Linear shaped charge cross section.
(Caitlyn McKay)
We cut three sides of the Lunar Lander using the charge; a fourth side has a hinge that
opens. The Rover deployment system includes a platform in which the Rover is situated
that is lowered by a steel cable. An example of the deployment system is seen in Fig.
8.4.6-2. The lowering platform increases the complexity as well as the mass during
deployment as seen in Table 8.4.6-1.
Author: Caitlyn McKay
Alternative Designs – Rover
Section 8.4.6, Page 283
Motor
Support Beam
Platfor
m
Steel
Cable
Fig. 8.4.6-2 An example of the deployment system used to lower the Rover from the Lunar
Lander to the surface of the Moon.
(Caitlyn McKay)
Table 8.4.6-1 Mass breakdown for Rover deployment
Item
Linear shaped charge
SOLIMIDE foam
Steel cable
Platform
Motor
Support Beams
Total
Mass in kg
0.580
0.030
0.82
0.025
0.025
0.13
1.610
Footnote: The table is an estimate of masses.
Author: Caitlyn McKay
Alternative Designs – Rover
Section 8.4.7, Page 284
8.4.7 – Structural Analysis
We initially designed the Rover using magnesium alloy as the material. Magnesium
while not as strong, has a lower density than aluminum. The magnesium alloy still
satisfied all of the Rover’s stresses. For the second design of the kamikaze rover, we
were considering Lexan material. Since the Rover design was decided against, stress
analysis was not done on the Lexan Rover.
The stresses the Rover endures are caused by the forces exerted during different phases
and the mass of the systems inside the Rover. These forces cause bending moments,
deformations and buckling. Sheet buckling will occur if the support is not thick enough
and the Rover cannot withstand the compression, shear or bending stresses.
The design of the Rover wheel, seen in Fig. 8.4.7-1, is an airless wheel commonly
referred to as a twheel developed by Michelin. The wheel is made from aluminum with
chevrons added for traction and is filled with Solimide foam to keep dust and rocks out.
Fig. 8.4.7-1 Schematic of Rover wheel
(Ryan Lehto)
We designed the wheel size to increase the probability of a successful mission while
maintaining a low mass. Rocks the Rover would have to drive over were considered to
be steps that needed to be climbed. The optimum size for the base of the Rover is also
found by the radius of the wheel. The optimum base size allows the Rover to drive up a
45 degree angle.
Author: Caitlyn McKay
Alternative Designs – Other Locomotion
Section 8.5.1, Page 285
8.5 – Other Locomotion
8.5.1 – Sled Alternative
We investigated a tracked vehicle steered by skis as a means to traverse the lunar surface
for the small 100g and 10 kg payloads see Fig. 8.5.1-1. In our analysis, we assumed the
track vehicle to have same mass as the rover (see Section 8.4). We determined that 0.13
m by 0.017 m track is needed to power the vehicle.
Fig. 8.5.1- 1 Track vehicle schematic.
(Ryan Lehto)
The track system was not an attractive alternative as the advantages for the light payload
were minimal versus the disadvantages.
Advantages

Increases traction

Does not sink into soil
Disadvantages

Heavier than most other systems

Adds complexity

Kicks up lunar dust
Author: Ryan Lehto
Alternative Designs – Other Locomotion
Section 8.5.2, Page 286
8.5.2 – Spring Launch Alternative
For the purpose of locomoting our payload 500 meters across the lunar terrain we
investigated the feasibility of using a spring launch system. We calculated the potential
energy of different springs to determine the maximum distance the payload could
traverse. First, a few assumptions were made; launching from ground level, initially the
system is at rest, and a 45° initial trajectory angle to maximize range. The maximum
distance is calculated from basic dynamics.
We looked into launching our payload the full 500 meter distance to determine if this
design should be given further consideration. For the 100 gram payload case the spring
system could launch the payload over 1000 meters. However for the 10 kilogram payload
case a multiple spring configuration is needed to get the payload 500 meters away from
the base. For this reason the possibility of using a spring launch for the large payload case
is impractical.
We continued to look at the viability of a spring launch system for the 100 gram and 10
kilogram payload cases. The first calculation did not include the weight of a container for
the payload and other gear necessary to accomplish the mission; including the camera
and communication equipment. While developing the spring launch alternative the Space
Ball concept was introduced. The Space Ball has a shell that protects the payload and it
allows the equipment to travel across the lunar surface. However, the Space Ball could
not roll on a 45° inclined slope and therefore needed an external system to accomplish
this task. We decided that the spring launch concept would be developed concurrently
with the Space Ball design. At the time of calculation, the potential Space Ball design
weighed 7.6 and 18.5 kilograms for the 100 gram and 10 kilogram payload cases,
respectively.
Author: Zarinah Blockton
Alternative Designs – Other Locomotion
Section 8.5.2, Page 287
Table 8.5.2-1 Spring Launch System Results
Variable
Mass
Free Length
Spring Constant
Velocity
Number of Springs
Maximum Height
Distance
100 gram
7.6
0.1524
18879
41.6
30
270
540
10 kilogram
18.5
0.1524
18879
40.7
70
260
520
Units
kg
m
N/m
m/s
-m
m
Table 8.5.2-1 shows that for both the smaller payload cases the spring launch design
completes the X PRIZE locomotion requirement. We determined the Space Ball shell can
withstand 30 g’s upon impact; therefore necessary analysis was done on the g loads at
launch and landing. The results prove the spring launch system is incapable of delivering
the Space Ball a substantial distance away from the landing site because of the structural
limitations. The limiting factor of this concept is the g loads upon launch. We looked into
to the possibility of increasing the length of the springs to reduce that, however the
necessary spring length exceeds the scope of feasibility.
Table 8.5.2-2 Spring Launch System Results
Variable
Distance
Launch g’s
Landing g’s
Spring length for 30 g
launch
100 gram
540
3549
26
18
10 kilogram
518
3402
25.5
17
Units
m
--m
We considered launching a partial distance away from the lunar landing site rather than
the full 500 meter distance to reduce the length of the springs and the g loads upon
launch. The maximum distance the payload can launch, and still be within feasible limits,
is 200 meters. However, the requirement for the Space Ball to move up a 45° inclined
slope was relaxed therefore eliminating its need for an external system to achieve that
task. The added complexity of creating a system to store and un-store compressed
springs, which still do not accomplish the assigned task, is the final deciding factor for
the elimination of the spring launch system from project Xpedition.
Author: Zarinah Blockton
Alternative Designs – Other Locomotion
Section 8.5.3, Page 288
Ski Alternative
The ski has an advantage over a wheel based system in that, where wheels require
traction in order to provide forward motion, a ski actually benefits from low friction
environments. Instead of gripping the surface, it slides along the top providing a platform
of support. The large contact area of a ski distributes weight over a large area which
allows it to sit on top of fine materials (i.e. lunar regolith) without sinking in.
The ski is a natural spring which lets it act as a suspension system. The spring like quality
of the ski is helpful in dampening any vibrations created by moving across the lunar
surface.
The ski is not able to turn without carving an edge into the surface it glides on. A separate
motor system would be required to provide turning. The lunar regolith is twenty times the
density of snow thus requiring significant power required to turn. Though the ski would
slide over the regolith with low friction, a ski is still being pushed or dragged along. The
power required to overcome friction would need to be accounted for.
The ski system requires a propulsion source. Unless the Lander could land on top of a tall
hill, the ski system would still need an additional propulsion system. In the event that the
ski rides up onto a large obstacle, the chance of tipping the craft over or removing our
propulsion system from the ground is quite high. Due to these risks, we chose another
alternative for our locomotion design.
Author: Cory Alban
About The Authors
Section 9, Page 289
9 – About the Authors
John Aitchison – As part of Mission Operations, John’s primary responsibility was
design and analysis of the lunar descent trajectory. After graduation in May 2009, he will
be heading to Northrop Grumman Corporation in El Segundo, CA to begin work in the
space systems business division. Eventually he plans on starting his own company, either
in the aerospace or adventure travel industries, possibly both. His ultimate goal is to
work one day a week, and spend the other six kayaking the meanest whitewater in the
world and piloting commercial sub-orbital flights.
Kara Akgulian – Kara is a member of the Mission Operations technical group and the
Locomotion phase group. In May 2009 she will be graduating from Purdue University
with a major in design and a minor in dynamics and control. This summer Kara will be
getting married and plans to start working full time in the Aerospace Industry. In the
future she will continue her education with a Masters degree in Aeronautical and
Astronautical Engineering.
Cory Alban – Cory is a member of the Mission Operations group with most of his work
being done on the lunar locomotion phase. He will be working with Rolls-Royce in
Indianapolis this summer. Cory spends his time ballroom dancing, playing guitar, and
living the dream.
Brad Appel – Brad is the group lead for the Propulsion Group and focused on the
Electric Propulsion for the OTV. After graduation Brad will continue at Purdue with
graduate research in propulsion. He was born in the backseat of a greyhound bus. It was
rolling down highway 41. After graduate school he will be looking for a good time down
on the Bayou.
Author: Solomon Westerman
About The Authors
Section 9, Page 290
Zarinah Blockton – Zarinah is a member of the mission operations group. She was
responsible for the launch vehicle selection for all three payload cases. She was also
involved in the locomotion phase of the mission. She researched the possibility of using a
spring launch system for the smaller payload cases. After graduation Zarinah plans to
teach overseas, over a year and a half period, in Vietnam and Chile. Upon her return to
the United States Zarinah will pursue a Master’s Degree in Aerospace Design. Zarinah’s
ultimate goal is to become an astronaut for the National Aeronautics and Space
Administration (NASA).
Levi Brown – After obtaining a Bachelor of Science in Aeronautical and Astronautical
Engineering, Levi will continue his education in the pursuit of a Master’s of Science in
Aeronautical and Astronautical Engineering. He will primarily study propulsion and
astrodynamics. Upon receiving a Master’s he will seek employment in the aerospace
industry. His long term goal is to become involved in trajectory analysis and mission
design for America’s manned or unmanned space programs.
Michael Christopher – Michael Christopher is an undergraduate senior in the
department of Aeronautics and Astronautics at Purdue University. Originally Michael is
from Buffalo, NY. Michael’s major contributions to Project Xpedition were in the
communications group. Upon graduation, in December of 2009, he plans on pursuing
employment most likely in the field of power generation. In the more immediate future,
during the summer of 2009 Michael will be interning for Solar Turbine in San Diego,
CA.
Tony Cofer – Tony hails most recently from Orange County Indiana and will graduate in
May. He is planning on going to graduate school at Purdue. His major is in propulsion
and minor in aerodynamics.
Author: Solomon Westerman
About The Authors
Section 9, Page 291
Andrew Damon – Andrew, hailing from Indianapolis, IN, will graduate in May 2009
with a BSAAE. He hopes to start a career working in manned spaceflight and also work
toward a graduate degree. Perhaps most importantly, he hopes to move out of the
midwest.
John Dixon – After finishing with my undergraduate education at Purdue University,
John will be attending the University of Nevada – Las Vegas to achieve a Masters in
Aerospace Engineering. Shortly after achieving the aforementioned degree he will then
begin his lifelong dream of World Domination. Falling short of that he will single
handedly invade Canada with nothing more than a bent paperclip, a piece of chewing
gum, a ball point pen, and a can of Pepsi, MacGruber style.
Joshua L. Elmshaeuser – Joshua’s main contribution to project Xpedition was serving
as a member of the communications group. Within the communications group he
researched hardware options for the Lunar Transfer phase as well as the locomotion
phase. He served as the communications liaison for the translunar phase, and assisted the
locomotion group with some of the space ball design analysis. The best experience for
him was modeling the lander system in CATIA providing an encompassing perspective
of the spacecraft we are to put on the Moon. His future plans are to leave the field of
Aeronautical and Astronautical Engineering behind for a more hands on application in air
and space. Upon graduation he will be a Second Lieutenant in the United States Air
Force, stationed at Columbus AFB Mississippi training as a Pilot.
Brian Erson – Brian will be continuing at Purdue through the fall of 2009. After
graduation, he plans on pursuing a Masters of Business Administration, then help develop
marketing systems for engineering products.
Kris Ezra – Kris is a member of the Attitude Group working primarily in the Lunar
Transfer phase and focusing on landing alternatives.
Author: Solomon Westerman
He has one more year of
About The Authors
Section 9, Page 292
undergraduate studies after which he will graduate with a bachelor of science in
astronautics with minors in applied physics, mathematics, and Spanish language and
literature. After graduation Kris will attend graduate school and will eventually pursue a
job with private defense contractors. He enjoys ballroom dancing, playing the piano, and
riding his motorcycle.
Adham Fakhry – Adham is a member of the Power group, focused on designing the
power systems for the Lander for all mission configurations and also worked on the
surviving the lunar night for the Lander. Adham is graduating this spring and hoping to
find an internship for the summer. After that he will be attending graduate school at
Purdue University to get his Masters degrees in Aeronautical and Astronautical
Engineering.
Thaddaeus Halsmer – Senior Design responsibilities were in the Propulsion group on
the Lunar Descent phase with primary focus on the Lunar Descent Vehicle main and
locomotion engine design. He will graduate in May 2009 with a Bachelor of Science in
Aeronautical and Astronautical Engineering and has accepted a position as a Rocket
Propulsion Development Engineer with Space Exploration Technologies (SpaceX) in
Hawthorne, CA.
Jeff Knowlton – Jeff plans to continue school until August when he graduates. His goal
is to become an astronaut, what path he will take to reach that goal is not set in stone, but
may include the military.
Kelly Leffel – Kelly worked on the Structures and Thermal Group for Senior Design.
Kelly spent most of the time working on the thermal control for the Lunar Lander. After
graduation in May, she will return to her hometown of Greensburg, Indiana to work a
summer job. Kelly plans on returning to Purdue University to receive a master’s of
Author: Solomon Westerman
About The Authors
Section 9, Page 293
science in Aeronautics and Astronautics Engineering. After graduating with a master’s,
she plans on seeking full time employment.
Ryan Lehto – After obtaining a Bachelor of Science in Aeronautical and Astronautical
Engineering with a minor in Computer Graphics Technology, Ryan will be working in
aerospace industry as a supplier technical support engineer contractor for Sikorsky
Aircraft. Long term goals include returning to school to obtain a Masters in engineering
and one day working as a test engineer for rocket propulsion firm.
Korey LeMond – Korey’s main contribution to Project Xpedition was to serve as the
Structures Group Lead and as the systems engineer for integration of the vehicles and
missions. Korey will be graduating in May. Upon graduation he will be moving to
Houston where he will be working with Cameron International designing subsea oil
pipelines and wellheads.
Josh Lukasak – Josh was given way too much responsibility in this project and was
named Attitude Group Lead and Lunar Descent Phase Manager. Josh is graduating in
May ’09 and is getting married in July because he’s not as smart as you think.
Unfortunately for Josh the change promised by President Obama hasn’t had a chance to
reach him before graduation. As a result he will follow his wife and live off her meager
teacher’s salary while attempting to optimize burger patty trajectories or something just
as demeaning. All the while Josh will continue to brag that he went to the same
University as Neil Armstrong.
Caitlyn McKay – Caitlyn traveled from South Dakota to go to school at Purdue. She
was part of the Structures and Thermal group.
This semester she focused on the
alternative rover design and was a 10 kg integration manager. She will commission as a
2nd Lieutenant in the United States Air Force in May 2009 when she graduates. Caitlyn
Author: Solomon Westerman
About The Authors
Section 9, Page 294
will be stationed at Randolph AFB near San Antonio, TX to begin navigation training
starting one week after graduation.
Ian Meginnis – Ian served as the group leader for the Power Systems group and the
phase leader for the Lunar Transfer phase. He also worked on developing the thermal
control system for the orbital transfer vehicle. His major is design and his minor is
aerodynamics.
After graduation, Ian plans on attending graduate school and then
working at NASA-JSC.
Trenten Muller – Trent was a part of the Communications group where he worked on
communication equipment as well as ground stations. He also explored an alternative
landing design that had the Lander impacting with a significant portion of its orbital
velocity. After graduating in May, 2009 he plans on joining the Air Force and working in
the aviation or space based field.
Ryan Nelson – This spring Ryan mostly did work on the frame design on the Lander for
each of the three payload cases. Early in the project he did some research and analysis on
the launch loads and vibrations for the Delta II launch vehicle as well as the Dnepr. After
senior design, he plan on graduating in May ’09. Following graduation he plans on
attending graduate school in the Midwest and further expand on his understanding of
structures and propulsion as they apply to the aerospace industry.
Tim Rebold – Tim was part of the Structures team for project Xpedition, and worked
primarily on the OTV. He was responsible for designing the configuration of the OTV
and sizing its structural members. Tim is majoring in Dynamics and Control with a
minor in Structures while he attends Purdue University.
He plans to continue his
education at Purdue University by seeking a Master’s Degree in Aeronautical and
Astronautical Engineering. When not busy with school, Tim enjoys time with his wife
and daughter, and playing ultimate Frisbee and racquetball.
Author: Solomon Westerman
About The Authors
Section 9, Page 295
Saad Tanvir – Saad has primarily been working on developing the radial flow hybrid
engine for Lunar descent phase of the mission. After obtaining a Bachelor of Science in
Aeronautical and Astronautical Engineering, Saad will continue his education by
pursuing of a Master’s of Science in Aeronautical and Astronautical Engineering. His
major area of research focuses on alternative energy with applications to the aerospace
industry. His long term plan is to return to Pakistan and help improve its aerospace
industry.
Christine Troy – Christine served as Project Xpedition’s Assistant Project Manager,
Webmaster, and was in the attitude control technical area. Specifically, she worked on
attitude control for lunar descent. After graduation, she plans to work in either the
aerospace or wind energy industries.
Brittany Waletzko – Brittany is a member of the attitude control group and is also the
integration manager for the 100g payload for project Xpedition. After graduating this
spring, she is going straight into graduate school. Beginning with a math course over the
summer, she'll be working on getting her Master's degree in the area of Astrodynamics &
Space Applications at Purdue University's graduate school. Brittany expects the nonthesis option will take less than two years to complete, after which she will be seeking
full-time employment.
Solomon Westerman – Solomon is the Project Manager for Project Xpedition. While
not devoting his time to Project Xpedition, Solomon enjoys cycling, guitar, and highpower rocketry. He has accepted a Guidance, Navigation, and Control Engineer position
at Space Exploration Technologies (SpaceX) in Hawthorne, CA, where he plans to live it
up. You can check his website at http://web.ics.purdue.edu/~skwester and contact him at
solomon.westerman@gmail.com.
Author: Solomon Westerman
About The Authors
Section 9, Page 296
Alex Whiteman – After finishing senior design, Alex will be graduating from Purdue
and then looking for a job. He also hopes to get a masters degree in Astronautical
engineering in the near future.
Author: Solomon Westerman
References
Section 10, Page 297
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