This document proposes an Avalanche Risk Assessment mission consisting of 3 satellites in low Earth orbit. The satellites will use laser altimetry to measure snow depth and slope of mountainous regions to improve avalanche prediction models. Data will be downlinked in near real-time to warn authorities of potential avalanche risks so preventative actions can be taken. The mission aims to prevent avalanches in over 80% of risk situations. It is estimated to save over 100 lives per year and $5 billion in economic impacts over the 10 year lifetime of the mission. The technical details of the spacecraft, sensors, communications, and operations are provided.

1. Avalanche Risk Assessment
AE 4356 Final Proposal
Brian Hardie
Sang Jeong
Sung Kim
Jonathan Saenger
Eric Stoker-Spirt
April 25, 2016
HONOR AGREEMENT
Having read the Georgia Institute of Technology Academic Honor code, I understand and accept
my responsibility as a member of the Georgia Tech community to uphold the Honor Code at all
times. In addition, I understand my options for reporting honor violations as detailed in the code.
____________Brian Hardie___________
___________Sang Jeong_____________
___________Sung Kim______________
___________Jonathan Saenger________
___________Eric Stoker-Spirt_________
Executive Summary
The Avalanche Risk Assessment (ARA) mission is a space mission that will introduce an
innovative method for determining avalanches. With current avalanche prediction methods being
unreliable, the need for better accuracy was determined. This mission will measure snow depth

2. 1
and land slope to more accurately predict avalanches. Snow depth and land slope will be
measured using laser altimetry, Success will be achieved if all launch operations are completed
and if sufficient data is collected to be able to prevent at least 80% of all avalanche-risk related
situations. The mission assessment concludes that the long term benefits of the mission will
outweigh the costs of the mission. It is recommended that this mission concept or similar
alternatives be explored.

3. 2
Table of Contents
1. Mission Objectives and Requirements Definitions
1.1. Mission Objectives
1.2. Impact of the Mission
1.3. Driving Requirements
2. Mission Implementation
2.1. Mission Architecture
2.2. Mission Description
2.3. Trajectory and Maneuver Design
2.4. Models Used
3. Flight System
3.1. Spacecraft Architecture
3.2. Launch Vehicle
4. Sensors
4.1. Payload Description
5. System Engineering, Risks, and Mission Operations
5.1. Driving Requirements and Trade Studies
5.2. Technical Resources
5.3. Risk Identification and Mitigations
5.4. Mission Operations
6. Management, Schedule, and Cost
6.1. Management Plan
6.2. Program Schedule
6.3. Cost Estimate
6.4. Descope Options
APPENDICES
Bibliography
Nomenclature

4. 3

5. 4
List of Tables
1.0 The lighting report summary from STK………………………………………………...11
2.0 The Mass Allocation Table……………………………………………………………...20
3.0 The Propellant Budget…………………………………………………………………..20
4.0 Power Allocation………………………………………………………………………..21
5.0 Link Budget……………………………………………………………………………..21
6.0 Risk Matrix……………………………………………………………………………...22
List of Figures
1.0 The Operational View………………………………………………………………………..8
2.0 3-D Orbit Visualization………………………………………………………………………9
3.0 2-D Ground Tracks…………………………………………………………………………...9
4.0 The map of the flow of information………………………………………………………….10
5.0 CAD Model…………………………………………………………………………………..12

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1.0 Mission Objectives and Requirements Definition
1.1 Mission Objectives
Primary:
● To take measurements of snow depth and land slope and use this information to predict
and report areas at risk of avalanches.
● To refine and improve a mathematical model to predict avalanches.
Secondary:
● To support meteorologists in snow measurements.
● To provide information on land elevation to update topographical maps.
● To give individuals and businesses confidence that mountains are safe.
Mission Requirements
● The mission must cover all locations on Earth that are heavily populated, mountainous,
and subject to large amounts of snow
● The mission must be able to collect data on these areas at a spatial separation of no more
than 20 meters between adjacent points
● The mission must have be capable of pointing accurately enough to take measurements of
all important measurements for every pass.
● The data must be downlinked sufficiently quickly to take action before an avoidable
avalanche occurs
The Avalanche Risk Assessment mission shall be considered fully successful if:
● All three satellites are inserted into their desired orbits.
● Sufficient data is collected over all inhabited mountain ranges.
● Data is downlinked, reduced, and sent to authorities of corresponding areas in adequate
time to employ preventative measures in over 80% of avalanche-risk situations.
● It provides data to update mathematical prediction models for avalanches.

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The Avalanche Risk Assessment mission shall be considered minimally successful if:
● Data is downlinked, reduced, and sent to authorities of corresponding areas in adequate
time to employ preventative measures in over 20% of avalanche-risk situations. This
corresponds to the prevention of slightly less economic damage than the cost of this
mission, plus minimal death and injury prevention.
● It provides data to update mathematical prediction models for avalanches.
1.2 Impact of the Mission
This mission will improve upon current avalanche prediction models by measuring the
snow depth at discrete points throughout a mountain range. Mathematical models calculate the
likelihood of avalanches based off of snow depth, land slope, temperature, wind speed, and other
factors. However, models currently estimate the snow depth from the snowfall of an entire
region. This is inaccurate, since snow does not fall evenly on mountains due to high wind speeds
and sloped surfaces. Avalanches tend to start at one point, so to accurately predict avalanches,
data needs to be gathered at discrete points, not an entire region. Our mission’s data will improve
avalanche prediction accuracy.
90% of avalanches in which humans are involved are caused by human activity, so if an
avalanche is known to be likely, an area can be closed off until containment methods are
employed, thus preventing the avalanche entirely. Therefore, this mission will be able to prevent
avalanches by predicting them.
Avalanches kill an average of 122 people per year, as well as injuring or otherwise
affecting 2,401 people.
Economically, $27,844,000 per year is lost to property damage, another $27,844,000 is
lost to legal fees, court settlements cost about $250,000,000, and rescues cost $1,000,000 per
year. Ski resorts are usually very cautious, closing more often than they need to in order to
minimize the risk of avalanche. However, with a more accurate model, these resorts would not
need to close down as often. If every resort stays open 1% more, an extra $200 million in
revenue can be gained per year. In total, the economic impact is about $506 million per year.
Over the 10-year lifetime of this mission, it has the potential to make a 1,220 life, $5 billion
impact.

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1.3 Driving Requirements
● Data downlink rate (Data needs to be sent down with very little delay so as to act as
quickly as possible)
● Coverage (All populated mountain ranges should be monitored)
● Vertical spatial resolution (snow measurements must be precise enough to be useful)
● Horizontal spatial resolution (measurements must be close enough together that no
important points are missed)
● Survivability (the longer this mission lasts, the more damage it can prevent.

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2.0 Mission Implementation
2.1 Mission Architecture
For the mission architecture, the subject of this mission is to observe the Earth to measure
the slope levels of mountain ranges in order to provide accurate determination of the potential
avalanche happening around the area. There are a total of 3 satellites utilized to be part of this
mission: each spaced out 120 degrees apart to cover the full 360 degrees with a 55 degrees
inclination with an altitude of 500 km (Low Earth Orbit). A variety of case studies and
evaluations were completed to determine which mission design would be the most suitable to the
condition of the mission while meeting the minimum requirements. TOPSIS analysis was not
considered after completing the Pugh Evaluation, as it provided enough evidence for the team to
conclude that operating this mission with 3 satellites would be most ideal. The cost was
determined to be less than three times as much, while gathering three times amount of the data
but also providing three times the amount of coverage surface area.
Figure 1. Operational View.

10. 9
Figure 2. 3-D Orbit Visualization.
Figure 3. 2-D Ground Tracks.
2.2 Mission Description
The data from the three ARA satellites will measure snow depth and land slope, which
will be downlinked to the ground station. The spacecraft position and velocity will be sent to the
ground station from the Space Surveillance Network (SSN), and weather information and wind
speed will be determined from meteorology data. This information will be autonomously input
into a computer, which will calculate the risk of avalanche at every measured location. If it is
determined that an area is at risk, a warning will automatically be sent to the authorities of that
area via a predetermined warning system.

11. 10
Figure 4. A map of the flow of information.
2.3 Trajectory and Maneuver Design
In the case that our launch vehicle misses the desired orbit, we can perform a plane
change maneuver or Hohmann transfer as needed. If the equatorial spacing between the three
separate satellites needs to be changed, two hohmann transfers can be used to change the altitude
to a lower/higher orbit then back to the original orbit.
Because we are in a low-earth orbit, periodic burns will be needed to offset the influence
of atmospheric drag on our orbit. To know when these burns are needed, we will acquire all three
satellite’s orbital states from the Space Surveillance Network (SSN). This position and velocity
data will be sent to our ground station where an autonomous program will decide if the orbit has
lowered enough to require an altitude-raising burn. If so, the ground station will send a command
to the appropriate satellite.

12. 11
2.4 Models Used
STK was used to model many different architectures to determine ground station access
time and mountain range coverage time. This information was critical in the trade study of
mission architecture, discussed in the “Systems Engineering” section.
In addition, STK was used to generate a lighting report. An orbit was simulated over the
course of a week. This simulation was used to determine the amount of time in eclipse vs.
sunlight, so that solar panels could be adequately sized. The lighting report was several pages
long, but is summarized in Table 1 below:
Table 1. The lighting report summary from STK.
Sunlight
Penumbra (treated as eclipse)
Umbra (eclipse)

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Figure 5. CAD
Model
3.0 Flight System
3.1 Spacecraft Architecture
3.1.1 Structures
The satellite body is made up of
space-grade aluminum. This material was
chosen for its strength-to-weight ratio as
well as its availability. The instrument
module, a cube, houses the dual laser
altimetry system. An attached octagonal
prism houses the rest of the satellite’s
subsystems.
The solar panels are separated from
the main body and angled at fifty-five
degrees. This, combined with servo motors
that rotate the panels, allows the panels to
stay perpendicular to the sun’s rays, allowing for maximum power generation.. A sun sensor on
the solar panels grants knowledge of the location of the sun and enables the satellite to correctly
angle the solar panels. The solar panels fold flat against the instrument module for storage.
3.1.2 Electrical Power System
The most important function of EPS is to provide power to other subsystems. This is
accomplished with solar panels, since a primary battery cannot survive for the design lifetime of
this mission, 10 years. The solar panels were sized to generate the power necessary to keep the
satellite in data collection mode 35% of the orbit, non-collecting mode 60% of the orbit, and
transmitting 5% of the orbit. Based on an STK lighting report, the satellite is in eclipse 32.71%
of the time. The solar panels must generate all necessary energy in the time spent in sunlight,
even at the end of life.

14. 13
Triple Junction Gallium Arsenide solar cells were selected due to their high efficiency
and slow degradation rate. With a 20% power margin, the solar cells were sized to 3.5 m2. See
power budget for more details.
The solar panels with charge a secondary battery. Over 10 years, the battery will undergo
58,000 charge-discharge cycles. For this cycle life, Nickel Hydrogen batteries can have a depth
of discharge of 60%, while Li-Ion can have 15%, and Ni-Cd can have about 7%. For this reason,
Ni-H₂ battery cells will be used. The battery has 81 cells of diameter 3 cm and height 6 cm and
will weigh 13.5 kg.
An extra 50 kg is included in the EPS mass, which accounts for extra electronic
components such as the EPS board and power distribution module. This is sized by analogy.
3.1.3 Thermal Management
The radiant heat sources that will be confronted by the spacecraft surfaces for this
mission will be the Sun, the solar energy reflected from the sunlit side of the Earth, and the
infrared from the warm surface of the Earth. The average temperature requirements for
components for this mission was from -10 to 40 degrees celsius with an average of 395 (20%
margin) W power dissipation. This mission is designed to have the satellites reside in LEO, and
this provides additional challenge for solar panels as they can be seriously affected by this rather
extreme temperature change due to the Earth’s infrared emission and rapidly changing eclipse
duration due to a relatively short orbital period (less than 2 hours approximately).
3.1.4 Propulsion
For the propulsion systems on the satellite, we chose to select
the engines by analogy. ICESat-2 used four 22N class MR-106
hydrazine monopropellant engines and eight 5N class MR-111
hydrazine monopropellant engines for their payload, thus we selected
the same class types for the Avalanche Risk Assessment payload. With
the engine types selected, the total propellant required was
determined through the use of the dry mass of the satellite and the
required ∆V for various maneuvers, shown in section 5.2.2. After

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finding the required propellant for the maneuvers, we sized the
propellant fuel tank to 250kg to allow for propellant margin.
3.1.5 Attitude Determination and Control
For this mission, sun sensors, star trackers, and GPS receivers were chosen for attitude
determination. The sun sensors determine the angle to the sun so that the solar panel stay
perpendicular to the sun’s rays. Star trackers determine the satellite’s orientation with respect to
the cosmos. Lastly, the team made an adjustment to include GPS receivers while extracting
gyroscopes to offer a lower cost and weight than the previous design.
For 3-axis control, three reaction wheels control the satellite’s orientation, with a fourth
as backup. The spacecraft can still point as long as three are functional. There are eight 5N class
thrusters that can also be used for attitude adjustment.
3.1.6 Telecommunications
The selected antenna for the payload was an omni-directional antenna, chosen to
preventing pointing interference with the sensors. We selected the ground station of White
Sands, New Mexico for its proximity. These satellites are in the Ka-band frequency range.
The worst case scenario to find required data storage was only one ground station
available to downlink and that the satellite continuously scanned data at 150 kB/s across all of
the Andes Mountains and the Rocky Mountains, a combined length of roughly estimate at
11,748.182 km. The total time spent scanning was found to be 1543.255 seconds. The total
gathered data was determined to be 231.49MB by multiplying data acquisition rate by time over
the mountains. The required downlink rate was determined to be 4.41 Mbps and a chosen
downlink rate to be 5 Mbps.
With the omni-directional antenna, dish diameter, and downlink rates determined, the
receiver antenna gain was found to be a value of 72.98 dB. The energy per bit to noise power
spectral density ratio (EbNo) received at the ground station was calculated to be 12.13 dB.
3.1.7 Navigation
Our mission requires us to take laser altimetry measurements. To reduce laser power
needed, we chose a low-earth orbit with an altitude of five hundred kilometers. To get coverage

16. 15
over the inhabited mountainous regions of the globe, we chose an inclination of fifty-five
degrees. Our launch vehicle injects us directly into our orbit, so there is no need for transfer
orbits, although minor station keeping maneuvers will be performed.
3.1.8 Command and Data Handling
The RAD6000 from Lockheed Martin was chosen as the processor for this mission
because it is very durable and has extensive flight heritage. This processor is sufficient to do the
standard tasks outlined in flight software and do minor data processing prior to downlink.
A 500 MB hard drive has enough storage to hold all necessary raw data and reduced
data, plus a 54% margin. Obviously, using a commercial off-the-shelf component reduces cost,
risk, and time associated with subsystem development, so this is much preferred over designing a
new component. This is true of all subsystems, but especially for C&DH, where the processor
and data storage need to be sized appropriately, but do not need any unique functionality.
However, finding a hard drive this small may be difficult. If one is not commercially available,
the smallest COTS space-rated hard drive above 1 GB should be used.
3.1.9 Flight Software
The ARA flight software will be responsible for navigation data, attitude determination,
laser pointing, fault detection, command and data processing and data uplink/downlink. The
operating system will allocate necessary resources to each component of the software.
Navigation data will consist of GPS coordinates and satellite pointing data. Due to the laser
system being fixed in the spacecraft, data collected from the rate gyro, sun sensor, and star
tracker will determine laser pointing and attitude determination. Fault detection will detect and
correct any faults within the system. The flight software will detect faulty measurements by
filtering out light that arrives back to the satellite too fast. If the light returns too quickly, the
light did not go down far enough to measure snow depth, but instead hit an obstruction. For data
processing, the time difference between the two signals can be multiplied by the speed of light to
obtain snow depth. This reduced data as well as the land slope will be sent down to the ground
station. The satellite will perform simple data processing in order to reduce the amount of data
being sent down. Data uplink and downlink will be processed autonomously in order to reduce
data transfer time

17. 16
3.1.10 Payload Accommodation
Payload accommodation will consist of laser pointing direction and laser power usage.
By using the attitude sensors (rate gyro, star tracker, and sun sensor) for data, the spacecraft
orientation (and therefore laser pointing direction) will be moved accordingly by the four
reaction wheel system and thrusters. In order to increase power efficiency, when the satellite is
passing over areas of little or no risk such as an ocean, the laser will be turned off. Data from the
GPS receiver and SSN will determine when and where to point the laser system.
3.2 Launch Vehicle
3.2.1 Launch Vehicle Selection
Our orbit inclination limits us to launching from one of two launch sites in the United
States: Cape Canaveral (CCAFS) in Florida and Wallops Island (WFF) in Virginia. From the
launch vehicles that launch from these sites, only two are in our approximate weight range of
1000-2000 kg. We chose the Orbital ATK Minotaur IV because it is closer to our satellite size
and does not waste as much fairing space or launch capacity.
3.2.2 Launch Vehicle Performance
The Minotaur IV is a four-stage vehicle used to get mid-sized satellites to low-earth orbit.
From WFF, the Minotaur IV can lift 1400 kg to a 500 km altitude orbit. For orbital insertion, the
Minotaur IV is accurate to within +/- 5 km altitude and +/- 0.1 degrees of inclination.
3.2.3 Flight System Integration with Launch Vehicle
Our satellite in its folded configuration has approximate dimensions of 2200 x 1200 x
1100mm. These dimensions allow our satellite to easily fit within the Minotaur IV fairing. Our
satellite will be attached to the launch vehicle through the payload adapter fitting.

18. 17
4.0 Sensors
4.1 Payload Description
The ARA payload will be a laser altimeter system (LAS). The design of the system will
be completed through analogy of the ATLAS (Advanced Topographic Laser Altimeter System)
from the ICESat-2 mission. The LAS will consist of two separate lasers. One laser will send
down visible light which will reach the surface of the snow. The other laser will send down
microwaves which will penetrate the snow and reach the ground surface of the mountain itself.
Both lasers will collect data at a frequency of 10 kHz. The time difference between the two
signals to return to the satellite will determine snow depth. Along with snow depth, the lasers
will also measure land slope and land elevation. Laser measurements will be carried out using a
Micro-Pulse Multi-Beam approach. The micro-pulse approach will essentially “separate” the
continuous laser beam into low energy repetitive pulses that will allow for more frequent
measurements. Studies show that with the Micro-Pulse Multi Beam approach, the ICESat-2 is
able to take measurements every seventy centimeters. The multi-beam approach will separate
each laser into six beams. The six beams will be paired off allowing three pairs of beams to
sweep across the land. The data returned from the pairs of beams will be averaged in order to
improve accuracy.
For coverage area, the spacing between the points measured will be roughly 20 meters.
One dual-laser can cover roughly 150 square miles per pass. With an access time average of
10.61067 minutes, roughly 6,366,402 measurement locations are available if 10,000
measurements are made per second. If one measurement per 400 meters squared is desired, than
the spacing between measurement points will be 20 meters. This will result in one laser being
able to cover 983.232 square miles. With the laser being split into three beams, this increases
coverage to 2950 square miles per satellite in a given pass.

19. 18
5.0 System Engineering, Risks, and Mission Operation
5.1 Driving Requirements and Trade Studies
5.1.1 Key Requirements and Implementation Approach
The key driving requirements for this mission were identified as downlink rate, coverage,
vertical spatial resolution, horizontal spatial resolution, and survivability. These drivers were
addressed as follows. The data is processed onboard before being downlinked. This way, all data
can be downlinked with a single ground station pass. The 3-satellite architecture ensures that the
mission has excellent coverage. The payload is designed by analogy to one of the most advanced
laser altimetry systems built, so vertical resolution will be about 1.2 cm, good enough to use in
avalanche models. The horizontal spatial resolution is not perfect; the pointing is not precise
enough to measure any exact point we want, but this is accounted for by gathering an abundance
of data. Statistically, one of the millions of data points gathered every minute will be placed
close to any important location that needs to be measured. Rather than pointing perfectly, a wide
net is being casted. The satellite subsystems were designed to last 10 years.
5.1.2 Major System Trade Studies
A Pugh analysis was performed to compare a 3-satellite architecture, 2-satellite
architecture, and 1-satellite architecture.
In addition, an omnidirectional antenna was chosen over a pointed antenna. It was
determined that the pointing and attitude complexity if both the antenna and payload needed to
be pointed was too much to justify the lower power usage of a pointed antenna.

20. 19
5.2 Technical Resources
Table 2. Mass Allocation.
Table 3. The propellant budget.
With these ∆V values above, the Isp of both classes of engines, and the dry
mass of the satellite, the required propellant mass was acquired through the use of the equation
below.
Then, surplus fuel was included for emergency collision avoidance or extra attitude control.
The amount of fuel was selected to fill the rest of a 250kg fuel tank. This gave a surplus of a
35.21kg of fuel which is enough for three altitude change maneuvers of 18.5 km away and back
to the target orbit of 500km.

21. 20
Table 4. Power Allocations.
Table 5. Link Budget.
5.2.5 Data Volume
After data is acquired, it will immediately be reduced to a 3-byte snow depth value and a
2-byte land slope value. The raw data will be deleted. This lowers the amount of data that needs
to be stored. Since the maximum time of operation for the payload is 40%, and 3 lasers store 5
bytes each at 10 kHz, a maximum of 324 MB need to be stored each orbit before being
downlinked. This is well within the satellite’s storage capability.
5.3 Risk Identification and Mitigation
Reaction wheel failure is a major component failure and would compromise attitude
determination and momentum control. If one reaction wheel fails, the remaining three reaction
wheels as well as the thrusters can be used. If the aid of thrusters is required, the mission lifetime

22. 21
will be much shorter. Orbit insertion and launch vehicle failure were identified as two major
system level risks. The launch success rate was approximated as the general rate of 90% . Three
launch vehicles produce a chance of failure of 27%. If one launch vehicle were to fail, then the
spacing between the satellites will be increased from 120 degrees to 180 degrees. If orbit
insertion failure occurs data will not be collected and fuel will have to be used for correction
orbital maneuvers which will reduce the lifespan of the entire mission. Promptness and orbital
debris were identified as mission level risks. No matter how accurate the data is, if the data
cannot be distributed in time for the authorities to evacuate the area, than failure has occurred. If
orbital debris were to come into contact with the satellite, there is a high possibility that the
satellite will become damaged and therefore unable to collect data. With orbital debris data from
NASA, orbital maneuvers may be made in order to avoid orbital debris.
Table 6. Risk Matrix.
5.4 Mission Operations
Mission operations is a two-phase process for the ARA mission. Launch operators will
handle the launch phase of the mission. Once in orbit, ground station operators will be
responsible for the day-to-day operations phase of the mission, including orbit correction
maneuvers and state of health checks, and emergency collision avoidance. Data movement will
be handled autonomously.

23. 22
6.0 Management, Schedule, and Cost
6.1 Management Plan
JPL engineers will review the development plan set forth by our undergraduate team. The
mission design will be developed further with the assistance of the students. The development of
individual subsystems will be based off of system drivers, which are based on overall mission
driving requirements. Students will assist JPL in the development of subsystems. The systems
engineer will also include students in his/her operations as an educational exercise.
Operations crews are necessary for satellite management. This mission will mainly
operate autonomously. Therefore, large operations crews are not necessary. However, five small
crews will be utilized so that the ground station is constantly monitored, in case emergency
action needs to be taken. In addition, routine station-keeping maneuvers are necessary.
A separate team will be responsible for analyzing the data in order to update and improve
the avalanche prediction model.
6.2 Program Schedule
From the point of funding for this mission, it should take about five years to develop and
ship to the launch site. This estimate is based off of other similarly sized missions. It also
considers that there are three spacecraft, so while the design phase does not change, production
of flight units will be lengthy.
The three satellites will be launched from separate launch vehicles. After the first launch,
the next will be in the next available launch window. Inside the launch window, the next satellite
will be launched at a time that will put it in orbit 120° away from the first. The last satellite will
be launched in the next window, again 120° away from the first. This launch strategy is flexible,
but it is logistically simpler than launching multiple satellites from the same place on the same
day, or transporting the payloads to three different locations.
Until all satellites are in orbit, those already in orbit will remain in safe mode to conserve
so that they have the maximum lifespan working as a constellation.
These satellites have been designed for a lifetime of 10 years, at which time the
remaining fuel will be used to de-orbit. During these 10 years, orbital maneuvers will be made

24. 23
when necessary; it is not possible to schedule these due to the unpredictable nature of orbital
perturbations.
6.3 Cost Estimate
USCM8 estimates a non-recurring cost of the mission, plus a qualification unit, to be
$558,940,625, with a standard error of $126,076,015. The recurring cost of the first satellite is
$200,602,810, with a standard error of $39,427,828. Assuming a 95% learning rate, the total
recurring cost of the three satellites is $554,818,606. The cost of the mission is $1,113,759,231
± $132,097,370. Including the launch cost of 3 rockets at $50 million each, the total is
$1,263,759,231 (Fiscal Year 2010).
After accounting for inflation from 2010 to 2016 dollars, these prices become:
● $1,246,478,092 ± $147,838,485 (USCM 8)
● With launch costs: $1.40 Billion
● 30% cost overrun: $1.77 Billion
Of course, an overrun is not expected, but seeing as overruns are not uncommon in the
space industry, it is worth noting that the cost could approach or even exceed $1.77B.
6.4 Descope Options
● If only two satellites survive launch, they will be repositioned such that their true
anomalies are 180° apart instead of 120°. If only one survives, it will not be repositioned.
● Having each laser split into two beams instead of three would use less power, mass, and
cost less. All supporting subsystems could accordingly be downscaled. This is a
possibility if the mission overruns cost, although obviously the amount of data collected
would be reduced.
● If launch puts the satellites on imperfect orbits, STK will be used to propagate the orbits
into the future. If the different orbits do not compromise mission functionality, the
satellites will not be maneuvered into their intended orbits, as that wastes fuel.
Appendix
Nomenclature
Mp: Mass of propellant
Mf: Dry mass of the satellite

25. 24
∆V: Instantaneous change in velocity
g: Gravitational constant
Isp: Specific impulse
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