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The Pennsylvania State University
Department of Aerospace Engineering
Aerospace 401B
Detailed Spacecraft Design
Final Report
SISO Explore Simulation Asteroid Retrieval
April 25th, 2014
Deep Space Express
Zachary Armagost
Adam Johnson
Jacob Lampenfield
Ramon Morales
John Perla
Galen I. Stuski
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Executive Summary
The SISO Explore Simulation Asteroid Retrieval and Utilization Mission is made up of
multiple components. The primary objective of this mission is to design a structure that can
successfully capture an asteroid and return it into a retrograde lunar orbit for further scientific
exploration. This may include describing the elemental composition as well as extracting resources
for refilling fuel. Once the asteroid is in lunar orbit, the utilization process is initiated. Based on
current estimates the mission requires a budget of approximately $1.4 billion for the 2013 fiscal
year. The intended asteroid is 2009 BD. This asteroid is relatively small (4-11 meters in diameter)
and has an orbit near identical to Earth’s. The asteroid also has a low rotation rate which will make
it easier to catch. These characteristics make 2009 BD an ideal candidate for capture. The
expected duration of this unmanned mission ranges from four to eight years. The synodic period
of the asteroid was estimated to be around 71 years so specific dates to optimize capturing
efficiency were considered to be unnecessary. Due to the location of the asteroid relative to Earth,
no specific launch date will produce a more efficient and effective capture mission. This allows
for adequate time for maneuvers to be completed under limited pressure. The average distance of
the asteroid during the duration of the mission is around 0.7 astronomical units. A graphical
timeline is shown in Figure 1. It shows the major milestones of the mission and when they occur.
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Figure 1: Graphical timeline.
The current mission plan takes the spacecraft to a heliocentric orbit to rendezvous with the
asteroid as shown in Figure 2 below. Table 1 describes each scenario within the mission in order.
Figure 2: Mission Scenario.
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Table 1: Mission Profile and timeline.
The main capturing mechanism is similar to a claw design. The claw will encapsulate the
asteroid and attempt to slow down its rotation. The structure of the spacecraft will need to protect
the internal components from radiation because the spacecraft will be operating in deep space. The
structure is going to be a prolate body made out of aluminum. The launch vehicle was chosen to
be the Falcon Heavy. This decision was made based on the price and the payload capacity. Figure
3 shows the spacecraft’s structure when all components are fully deployed.
Mission Profile Month and Year
Operations in Low Earth Orbit
I. Launch Jun-18
II. Payload Separation Jun-18
III. Earth Orbit Jun-18
IV. Lunar Gravity Assist Jul-18
Operations in Heliocentric Orbit
V. Enter Heliocentric Orbit Jun-18
VI. Thrust into Asteroid's Orbital Plane Jun-18
VII. Phase Orbit to Approach Asteroid Aug-19
VIII. Observe Asteroid March-20
IX. Capture Asteroid April-20
X. Phase Orbit back to Earth April-20
Operations in Lunar Orbit
XI. Transfer into Low Earth Orbit Feb-22
XII. Transfer to Lunar Orbit August-22
XIII. Place Asteroid into Lunar Orbit August-22
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Figure 3: Conceptual Structural Design of Spacecraft.
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The main propulsion system is four Hall effect thrusters powered by solar arrays. They will
be arranged in a diamond shape so that two thrusters can be used in times that do not require
maximum amounts of thrust. If all thrusters are firing, the maximum amount of power that could
be used for each is 12.5 kW. If two are firing, 25 kW of power can be used to provide the most
thrust for the spacecraft, which would be on the order of 1900 mN of thrust. The Hall Effect
thrusters while using 25 kW could provide a specific impulse of 3250 s per thruster. It is estimated
that this amount of specific impulse will be more than sufficient, including a reasonable storage of
propellant to achieve the total required Δv for the mission. The attitude control propulsion devices
are similar to vernier thrusters which will provide small impulses. The propellant for the main
thrusters will be xenon gas, which was used on Deep Space 1 and the SMART-1 spacecraft and
numerous other missions with a 100% success rate. The exact value for the spacecraft's Δv is 8.2
km/s. Based on fuel consumption related to Δv for those missions, it is estimated that the amount
of xenon gas required for this mission will be on the magnitude of 240-500 kg.
Ground control will use the Deep Space Network to communicate with the spacecraft and
Mission control, payload operations, and spacecraft operations control center will be set at Jet
Propulsion Laboratory (JPL) in Pasadena, CA. It will be responsible for tracking and periodic
updates since the spacecraft is mostly autonomous. The communication subsystem will include an
error detection and correction code that will be analyzed by the ground control stations. The
communication system will also be responsible for sending ground control updates on position and
state. The data rates of the communication subsystem will be 120,000 bps. The sensors onboard
the spacecraft will scan properties of the asteroid such as rotation speed and topography to
determine if additional maneuvers have to be made. It will need a high uplink frequency to receive
any updates from ground control. The architecture of the computer network will a bus and three
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computers will be used to increase the success rate of the mission as a result from redundancy. The
command and data handling subsystem must be robust and withstand radiation because of the
harsh conditions and be able to process a large amount of data.
Guidance, Navigation, and Control is responsible for a safe rendezvous with the asteroid
and completing a number of thrust burns in deep space. To detect position of the spacecraft
Microcosm Autonomous Navigation System (MANS) and Earth and Star Sensing System will be
used. Laser Doppler velocimetry will measure relative velocity between the spacecraft to the
asteroid. This will ensure that the spacecraft will approach the asteroid at a safe velocity for a
smooth capture because the relative velocity between the two bodies needs to be small so that they
do not cause harm to each other. If the relative velocity between the two bodies are too large they
will either be drifting apart or approaching each other quickly, which could result in a collision.
A system of two moveable disks will be used to help realign the center of mass once the asteroid
is captured. An IMU will also be used to help measure changes in the spacecraft’s inertia to ensure
a properly executed capture.
Two 22 meter disc solar arrays will generate power for the spacecraft. These solar arrays
will generate approximately 53 kW of power. This power will be stored in lithium ion power for
when the spacecraft does not have access to the sunlight. Hughson White Paint A-276-1036 will
coat the spacecraft to help reflect sunlight from heat, keeping the interior temperature about
-24°C. This will be used in combination with radiation fins, a diphasic loop, and heating pads to
keep the internal components at a working temperature of about 0°C for the electronics and 21°C
for the batteries. A claw mechanism will be carried on the spacecraft to capture the asteroid
payload. A backup design has been inserted into the capturing mechanism in case of a malfunction
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from the main device. This is a carbon fiber mesh material that will aid in slowing the rotation of
the asteroid as well as provide extra stability once the asteroid is captured. A harpoon is also used
as a secondary backup system as well. The initial mass is estimated to be 28,000 kg. The solar
power arrays produce about 52 kW of power for the entire spacecraft, which is more than enough
for all the power needs of the spacecraft.
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Table of Contents
1. Introduction ……………………………....………………………………………….1
2. Mission Architecture
a. Overview of Mission ……………………………………………………....……...3
b. Sequence of events: Low Earth orbit ……………………………………………..4
c. Sequence of events: Heliocentric orbit ……………………………………..…….5
d. Sequence of events: Lunar orbit ………………………………………………….7
3. Subsystems
a. Structure ………………………………………………………….………………8
b. Launch Vehicle ……………………………………………………...…...………11
c. Propulsion ……………………………………………………….………………12
d. Ground Control ………………………………………………………………….15
e. Communications …………………………………………………………..…….17
f. Command & Data Handling …………………………………………….………17
g. Guidance, Navigation, and Control ……………………………………………..23
h. Power ……………………………………………………………………………26
i. Thermal …...…………………………………………….……………………….30
j. Scientific Instruments and Other Payloads………………………………………32
4. Summary Tables ………………………………………………….…………...……34
5. Conclusion ……………………..………….………………..………………...…….36
6. References ………….…………………………………………….…………………38
7. Appendices ……………………………………………………….…………...…….40
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1. Introduction
The importance of finding raw materials in the void of space is becoming more pertinent
to missions involving manned or unmanned spacecraft. In the solitude of outer space, supplies
become progressively scarcer as time goes on. As a result, asteroid mining has become a feasible
option in finding raw materials during missions. In order to determine if the asteroids can be
utilized for mining resources, retrieval of a near Earth asteroid is necessary. This idea also supports
future human deep space exploration. This mission also scouts other asteroid candidates that may
be accessible for possible future missions. This would be highly beneficial in a number of various
missions in outer space and will open up a new realm of possibilities for longer sustainability of
interplanetary missions.
Modeling and Simulation is a key component in the mission design phase. The purpose of
this project is to design a spacecraft architecture prototype in order to be utilized in missions
regarding asteroid retrieval. The Simulation Interoperability Standards Organization (SISO), will
build a simulation based on the design of the spacecraft that will be created by the Deep Space
Express team. The primary objective of this mission is to design a structure that can successfully
capture a pre-determined near Earth asteroid and return it into distant retrograde lunar orbit for
further scientific exploration and research to be conducted. A lunar orbit was preferred over a low
Earth orbit to optimize the use of propulsion and aid the specimen to avoid the influence of Earth’s
gravity. LEO contains a plethora of space debris, satellites, and other obstacles that may interfere
with the captured asteroid. Returning the spacecraft into lunar orbit also avoids the risk of
impacting the Earth. After the entirety of the mission gets completed, NASA will plan a mission
of its own in order to utilize the asteroid. A final prototype was selected based on efficiency, cost,
sustainability, and re-usability and a final conceptual design review was made. One of the
secondary objectives is to develop the methods and technology capable of completing the mission
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to enable implementation of similar missions in the future. The completion of this mission will
further scientific understanding of celestial bodies and contribute to the success of missions to
come. Having the asteroid accessible for research will provide valuable information of the
elemental composition of asteroids. Once the process of extracting resources is implemented, other
advanced techniques such as commercial resource production or distribution may be considered.
The design level of this report is in a completed conceptual design phase. The
specifications of the subsystems was determined by the most efficient material parameters to create
a well-balanced vehicle for the mission. Each subsystem has also completed its mass, power, and
cost estimates. The subsystems include structure, launch vehicle, propulsion, ground control,
communications, command & data handling, guidance, navigation & control, power, thermal
control systems, and scientific instruments and payloads. The payload for the mission has been
decided based on a variety of factors. The asteroid named 2009 BD was chosen particularly for its
size, which was determined to be 4 – 11 meters. The velocity as well as the rotation rate was also
found to be relatively slower than other potential asteroids. The v∞ and rotation rate for the asteroid
is 1.2 km/s and less than one rotation every 2 hours respectively. These parameters make the
capture of the asteroid much easier and helps for the design of the structure. Upon the departure of
the launch vehicle the distance between the Earth and the asteroid will be around 0.7 Astronomical
Units. The mission is expected to take about 3.5-8 years to complete. (7)
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2. Mission Architecture
2.a Mission overview
The SISO explore simulation Asteroid Retrieval mission is to be completed in three phases.
These include Low Earth orbit, heliocentric orbit, and lunar orbits. The timeline of the mission is
shown in Table 2. The launch date for the mission is scheduled to be in June 2018 and the duration
of the mission is expected to last around four to eight years. This date is of no particular importance
since the asteroid remains relatively close to Earth’s orbit, at a range of 0.5 to 0.7 AU away. After
the synodic period was calculated to be around 71 years, the idea to complete the mission was
proposed to continue since a time limit is of no detriment. Upon launch, the structure will retrieve
the asteroid and return it into lunar orbit with the payload to be further examined. The entire
mission is currently predicted to need approximately 8 km/s of change in velocity. Figure 4
illustrates the mission scenario. A detailed sequence of events is shown in the following pages.
Table 2: Mission Profile and Timeline. These are the earliest estimations that have been
conducted. The actual mission will most likely be 3.5-8 years.
Mission Profile Month and Year
Operations in Low Earth Orbit
I. Launch Jun-18
II. Payload Separation Jun-18
III. Earth Orbit Jun-18
IV. Lunar Gravity Assist Jul-18
Operations in Heliocentric Orbit
V. Enter Heliocentric Orbit Jun-18
VI. Thrust into Asteroid's Orbital Plane Jun-18
VII. Phase Orbit to Approach Asteroid Aug-19
VIII. Observe Asteroid March-20
IX. Capture Asteroid April-20
X. Phase Orbit back to Earth April-20
Operations in Lunar Orbit
XI. Transfer into Low Earth Orbit Feb-22
XII. Transfer to Lunar Orbit August-22
XIII. Place Asteroid into Lunar Orbit August-22
4
2.b Low Earth orbit
The mission is expected to be launched from Cape Canaveral, Florida. The location was
chosen due to the low inclination, which uses the Earth’s rotational speed to its advantage. The
launch vehicle was determined to be the Falcon Heavy rocket, and the structural design of the
system was constructed to fit the specifications of the mission as well as fit into the launch
vehicle. Shortly after low Earth orbit is reached, the spacecraft will orbit for a few days in order
to ensure the systems and other electronics on board are operational. After all systems are assessed
for their functionality, the launch vehicle separates from the structure and begins to make a phase
orbit into the asteroid’s orbit using the five hall effect thrusters as its primary propulsion source as
well as the 16 Vernier thrusters in order to maintain the proper orientation. The estimated time for
the first phase is around 1.5 years.
Figure 4: Mission Scenario.
5
2.c.Heliocentric Orbit
The location of the asteroid is on a slightly more eccentric orbit than Earth's orbit around
the sun. The spacecraft will perform a Hohmann transfer to reach a heliocentric orbit in order to
reach the asteroid. After this maneuver the spacecraft will be approximately 450 ahead of the
asteroid. The spacecraft is to stay on its orbital path and transfer into the asteroid's orbit using a
phasing orbit. It will have to complete a full period before it can reach the target. Sensors and
communication from the spacecraft to the payload operations control center will provide us with
intricate information on the asteroid that will help with a successful mission. The exact size, mass
distribution, topography, and precise rotational speed will be scanned in order to perform the
mission as efficiently as possible. The spacecraft will approach the asteroid from behind,
performing thrust maneuvers to attain a speed capable of catching up with the asteroid's velocity.
Once the spacecraft is within the proper range, the capture phase of the mission will begin. The
spacecraft must match the asteroid’s rotational speed to optimize the claw’s capturing parameters.
A conceptual structural design is illustrated in Figure 5. A more detailed examination of the
structure's specifications is shown in the structure’s subsystems portion of this report.
After the claw mechanism is securely latched onto the asteroid, the retrieval phase of the
mission is initiated. A hydraulics system is used to power the claw’s grip. It is intended to hook
around the asteroid’s geometrically smallest side. A secondary capturing mechanism will be
deployed along with the claw’s grip to reinforce the capturing phase in the form of composite fiber
chords. In the event that the hydraulics system fails and the claw is unable to perform its task, the
chords would be able to latch itself to the target and complete the mission. A third and final
mechanism in the form of four harpoons powered by compressed nitrogen. This will be deployed
in a last resort attempt to complete the mission. If the claw acquisition capturing mechanism is
unable to open or close due to a malfunction in the hydraulics system, the harpoons will be shot
6
into the center of the asteroid. Upon impact, the asteroid will be reeled in and attached to the front
of the spacecraft. Once the capture phase is complete the spacecraft must stabilize itself and align
its orientation towards the correct path. The added payload increases the systems entire mass and
decreases its initial speed. The counter masses and vernier thrusters will be used to maintain the
proper orientation on its return trip. The estimated time for the second phase is about two to three
years. Most of the fuel consumed is during this phase in the mission due to the multiple maneuvers,
capture, and adjustments made to the spacecraft’s position.
Figure 5: Conceptual Structural Design of Spacecraft.
7
2.d Lunar Orbit
The spacecraft, along with its new payload, will orbit in its path until it is in a location
close enough for another Hohmann transfer into Earth's orbit. A prograde burn will commence in
effort to maintain the proper speed for the maneuvers. As it travels in the terrestrial orbit, the
spacecraft's velocity will be adjusted to make its final transfer into lunar orbit. The velocity is
adjusted for the purpose of future missions, which may include rendezvous in order to gather
resources from the captured asteroid. This particular phase was estimated to take about one to two
years. The added payload mass will alter the speed of the spacecraft so fuel levels must be
conserved during this portion of the mission in case of any unforeseen circumstances. Occasional
controlled burns may be applied to maintain lunar orbit. Based on future mission requirements, the
spacecraft may remain in lunar orbit anywhere from several months to several years. This parking
orbit will serve as a convenient spot to begin the extraction of resources and scientific exploration.
The entire structure may or may not be kept for use in future asteroid retrieval missions based on
its condition and success in completing the mission requirements. If it is deemed to be in working
condition the spacecraft can be refueled and used in other asteroid redirect missions. However, if
the spacecraft is in poor condition it will be separated into its modular parts and then be burnt up
in the Earth’s atmosphere.
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3. Subsystems
3.a Structures
The structure is important for encasing the subsystems to form a cohesive spacecraft. It
must endure the stresses and vibrations from launch, as well as the capture and retrieval of the
asteroid. The spacecraft will have a diameter of 4 m and a length of about 12 m. The spacecraft
will have a prolate shape, with all of the onboard computers and circuitry in the lower end near the
main propulsion system and the capturing mechanism near the top. The structure is responsible for
protecting all of the onboard equipment from the harsh environment it will encounter in space.
This environment includes hazardous radiation, large temperature fluctuations, and space
debris. Radiation is a bigger concern to the integrity of the system because the spacecraft is
traveling far from the Earth’s magnetic protection. Many of the onboard electrical systems are in
danger of being damaged by the intensity and duration of radiation exposure. In order to prevent
complications with the electronic systems, all on board technology will include radiation hardened
components. Special materials will also be selected to help shield the vital electronics in the interior
of the spacecraft.
The structure must be designed to contain all of the electronics and essential equipment in
order to carry out the mission. All of the onboard equipment will add to the mass of the spacecraft
and is estimated in Table 5. With the total mass of the spacecraft around 28,386 kg, the amount of
fuel and power needed was calculated in order to stay within the parameters stated in the mission
overview. The structure must also be able to withstand the forces and vibrations from the onboard
propulsion system which was estimated to be around 1900 mN. During launch, the launch vehicle
will transmit vibrations and stresses to the structure. The material that will be used to create a
robust yet versatile mainframe for the spacecraft was determined to be aluminum. The lightweight
material will decrease the mass of the spacecraft and the strength of the material will serve as a
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strong base for the rest of the subsystems. The geometry of the beams were selected carefully so
the spacecraft can survive the forces and vibrations applied to it.
When the asteroid is captured, there are many additional factors that must be accounted
for. One of the primary concerns is the added mass to the system. The additional mass will also
reduce the speed of the spacecraft, so therefore there must be adequate room allocated towards the
consumption of fuel. This space would take approximately 4.189 cubic meters. The remaining
space consists of the computer systems, capturing device, and communication systems. The
structure must be able to handle any forces and moments applied to it by the capturing of the
asteroid, while maintaining a hold of it. The attitude thrusters are positioned in four different
locations along the spacecraft. The vernier thrusters will be closer to the lower portion of the
spacecraft near the electrical systems and main propulsion system. This was decided in order to
avoid contamination of the capturing mechanism from the fuel ejection. The asteroid must remain
sterile to properly exploit the resources. Another important factor is that the center of mass of the
spacecraft must be in line with the propulsion system. The attitude thrusters will also help to
maintain the proper orientation by applying occasional controlled burns. Another method to keep
control of the spacecraft’s attitude are two nacelles of the structure that can rotate to any position
to get the center of mass back in line with the propulsion. The mass in these two nacelles will be
off center, allowing for adjustments to the center of mass of the spacecraft once the asteroid is
captured. Another way the spacecraft will acquire energy is through the use of solar panels. These
were also designed to fold into a compacted form as well as maintain symmetric shape and reduce
a dramatic change of the center of mass. The solar arrays contain a 22m diameter and are located
at the center of the vehicle as shown in Figure 6.
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Figure 6: Spacecraft Concept Design.
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3.b Launch Vehicle
The choice of launch vehicle has been narrowed down to the Falcon Heavy. If needed, the
spacecraft can be split and launched in two separate launches using either the Delta IV Heavy or
Atlas V. For simplicity and cost reduction, the Falcon Heavy is the best choice.
The mass of the spacecraft is about 33700 kg which is well within the Falcon Heavy’s
53000 kg limit. This also gives a mass buffer in case the spacecraft exceeds the estimated mass.
The Falcon Heavy has a payload fairing with an outer diameter of 5.2 m and a length of 12 m.
This will accommodate the spacecraft, which has a stowed diameter of 5 m and a length of 12 m.
(5)
The Delta IV Heavy is the largest of the Delta IV family of rockets. They were designed
by Boeing and are being built by United Launch Alliance. The Delta IV Heavy has the capability
to lift 27569 kg into Low Earth Orbit (LEO). There have been seven launches of the Delta IV
Heavy, with its first being a partial failure. This failure was due to a premature cut off of the
booster rockets. This partial failure is the only issue that the Delta IV family of rockets has ever
had. With a cost of $254 million, this is an expensive option, and 2 will be needed to lift the
spacecraft into orbit. This option would also complicate the mission by adding another
rendezvous. (2)
The Atlas V is a rocket designed by Lockheed Martin and built by United Launch
Alliance. It has the ability to lift 29400 kg to LEO. The Atlas V has had 39 successful launches
and only one partial failure, where the second stage shut down too early. The Atlas V has a payload
fairing of either 4 or 5 meters allowing for larger payloads to be launched. There have also been
proposals to allow payloads of up to 7 meters in diameter, expanding its capabilities. At a cost of
12
$223 million, the Atlas V is not as affordable as the Falcon Heavy, and two must be used to launch
the vehicle in its current configuration.
The Falcon Heavy is the primary launch vehicle because of its heavy lift capacity and the
ability to launch the entire spacecraft in one vehicle and avoid the complexity of a rendezvous. It
also has a low cost at $135 million, which is much cheaper than one of either the Delta IV Heavy
or Atlas V. Although the Falcon Heavy has not been completed yet, SpaceX has a promising
record with their Falcon 9 rocket, which is a smaller version of the Falcon Heavy. (5)
3.c Propulsion
The original propulsion system that was considered was Magnetoplasmadynamic
Thrusters. The MPD thrusters use Lorentz forces and have high specific impulses on the order of
4000-6000 s and high thrust capabilities on the order of 26,000-88,500 mN. The issue with the
MPDT is, that it is still highly experimental and has not flown a technology validation mission due
to the power requirements, which are on the order of 1,500-7,500 kW of power. This is why a
specially designed nuclear reactor was considered for the mission that would provide 5-6 MW of
power based on initial thermodynamic calculations. It was ultimately decided that a propulsion
system with more flight heritage should be used for this mission. (15)
The propulsion system that will be used is an arrangement of five Hall Effect Thrusters,
with a maximum power usage of 10 kW per thruster. They will be arranged in a diamond shape so
that two thrusters can be used rather than four to provide maximum thrust if needed. If all thrusters
are firing, the maximum amount of power that could be used for each is 12.5 kW. If two are firing,
25 kW of power can be used to provide the most thrust for the spacecraft, which would be on the
order of 1900 mN of thrust. The Hall Effect thrusters while using 25 kW could provide a specific
impulse of 3250 s per thruster. It is estimated that this amount of specific impulse will be more
13
than sufficient, including a reasonable storage of propellant to achieve the total required Δv for the
mission. The propellant for the main thrusters will be xenon gas, which was used on Deep Space
1 and the SMART-1 spacecraft and numerous other missions with a 100% success rate. Based on
fuel consumption related to Δv for those missions, it is estimated that the amount of xenon gas
required for this mission will be on the magnitude of 11,000 kg. However, many more calculations
and the consideration of the amount of mass must be accounted for, which could potentially alter
this estimate dramatically. Figure 7 below demonstrates how a Hall effect thruster functions and
Figure 8 below shows an example of a xenon gas HET in operation using 2 kW of power. (15)
Figure 7: Hall Effect Thruster Operation Method (21).
14
Figure 8: Hall Effect Thruster in Operation with 2 kW of power (21).
In addition to the Hall Effect Main Thrusters, there are several attitude control thrusters
that can be used to rotate the spacecraft or change its direction for orbital maneuvers and to
properly match the asteroid’s velocity and rotation rate to make capture possible. These thrusters
are small rocket burst fire engines that are based on the vernier rocket engines used on missile
attitude control systems. The spacecraft will have eight of these thrusters in the rear structure and
four in the front structure that all will have the ability to rotate 90˚ for multi-direction impulse
maneuvers. Two helium tanks supply gaseous helium pressure to the oxidizer and fuel tanks. The
fuel for these thrusters will be monomethyl-hydrazine and the oxidizer will be nitrogen tetroxide,
both supplied under a helium gas pressurization system. (15)
Figure 9: Vernier Thruster used on Space Shuttle (20).
15
Figure 10: Attitude Control Thrusters Firing (23).
3.d Ground Control
Due to the autonomy of the spacecraft, ground control will be mostly responsible for
tracking the spacecraft. The spacecraft will communicate to Earth via the Deep Space Network
(DSN). (18) The communication system will include an error detection and correction code which
will have to be analyzed by ground control. This error code ensures that the signals received are
less corrupted by noise and power loss. (4) Through the communication and GNC subsystems,
ground control will receive updates on position and state. Based on this information, ground
control can decide if an additional thrust burn is needed to maintain the trajectory. The spacecraft
is to follow the asteroid to measure its characteristics such as rotational speed and
topography. Ground control will run simulations to confirm a safe capture is possible. If
simulations do not have a positive mission outcome, a failsafe method will be executed.
Along with the data from the spacecraft, NASA’s Near Earth Objects (NEO) and space
debris programs will help track any potential obstacles in the spacecraft’s orbit. If any obstacles
appear, ground control will be able to perform immediate thrust burns to avoid collision.
Mission, payload operations, and spacecraft operations control centers will be set at Jet Propulsion
Laboratory (JPL) in Pasadena, CA. JPL is currently mission control for over 100 satellite missions.
16
(9) JPL is also home of NASA NEO and other useful resources making it an ideal location to set
mission control. Their expertise and heritage in past space exploration missions would aid to the
success of the mission. Since project is a mission proposed by NASA, the operations would be
easily integrated into their facilities. This saves on the cost of renting or constructing control
centers from scratch. The mission may also be given a certain level of priority, which will allow
the mission to be executed on its scheduled time.
If the spacecraft cannot successfully capture the asteroid, ground control will decide
between possible outcomes. If the asteroid cannot be captured due to some physical property such
as rotating too fast, the spacecraft may try to rendezvous with another asteroid. Possible near Earth
asteroids candidates for back up include 2011 MD or 2013 EC20. There are two main problems
with this solution however. The first is if the claw can handle different sized asteroids. These
asteroids are similar in size to 2009 BD, but 2011 MD may be a little larger which could pose an
issue. The backup harpoon capturing mechanism will be the main capturing mechanism since it
does not rely on size. The other issue is whether the spacecraft will have enough fuel left to execute
the rendezvous and arrive back to the moon. If there will not be enough fuel, the spacecraft could
arrive back to Earth for a potential refueling, but this is a difficult task. However enough fuel is
onboard to accomplish three similar sized missions with a safety margin of 7 km/s Δv.
If there is a mechanical failure like the claw or harpoon not able to deploy, the mission will
have to be considered a failure. There is no easily reasonable solutions to help correct for failure
of both capture mechanisms.
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3.e Communications
The spacecraft will send communications to the Deep Space Network. Based on the
planned orbit of the spacecraft, there should not be interference with the communication system to
ground control due to a foreign body like the sun. The only time Earth will not be within sight of
the vehicle is when it is orbiting the moon for release of the asteroid. While in lunar orbit relay
satellites can be used for crucial messages.
Another way to decrease power consumption on the spacecraft is to add error detecting and
correcting code techniques. (4) Using error detection and correction code will help improve
ground control’s accuracy to analyze the spacecraft’s signal.
A high gain/high gain will be used to transmit the data between the spacecraft and
Earth. Potential frequency bands are, Ka-Band Deep Space and X-Band Earth Science. Ka Band
allows for faster links, which will be important for mission critical messages, but will be more
susceptible to power loss. This frequency band will be used for uplink to earth communications
which have a frequency of 34.20-34.70 GHz. (12) The downlink system will used X-Band
frequencies due to their history of reliable spacecraft to earth communications. The downlink
antenna will operate at frequencies between 8.4-8.45 GHz. (12) A third antenna will be added to
the spacecraft for redundancy and near Earth communications (less than 2 million kilometers from
Earth). It will operate in the X-band at 7.190-7.235 GHz for uplink signals and 8.45-8.50 GHz for
downlink.
3.f Command & Data Handling
The primary function of the command and data handling subsystem is to process and
distribute commands, as well as process, store, and format data. This subsystem is related to the
communications and ground control subsystems. Ground control uses the communications
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subsystems to send procedures to the on board computer and vice versa. (18) The on board
computer will primarily perform autonomously and likely receive very few commands from
ground control.
The command processing aspect of this subsystem will need to execute orbit and attitude
corrections and changes throughout the mission. These commands will come from ground control,
during the time the spacecraft is on its way from Earth to the asteroid. These commands are low
in complexity and do not take up a lot of processing. These commands also do not need to be
performed in real-time. The position of the spacecraft can be estimated and the command can be
sent early, so the transmission delay can be ignored. These commands could also be predetermined
and stored on the computer.
The telemetry command processing aspect will be a highly important in this
mission. Whenever the spacecraft arrives at the asteroid, many attitude commands will need to be
processed in real time. This needs to be done autonomously as the transmission delay will cause
the command to be executed while the asteroid is in a different position. A variety of sensors to
track the position, relative velocity, and center of the asteroid will be used. The computer needs
to be able to process all of this data quickly and accurately.
There are three different architectures that can be used for the command and data handling
subsystem. The first that can be used is the centralized architecture. This type has one central
processor connected to each instrument or subsystem. This is a highly reliable
architecture. However, it works best with only a few subsystems. Another problem with this
architecture is that there is a large amount of wiring that needs to be done. Whenever a new
component is added, more wiring needs to be implemented as well as new software. (10)(18)
19
The next architecture is the ring. The central processor and each component are connected
in series. This allows for each component to have the same information as well as limits the
amount of wiring. However, if one component fails, the other components can no longer receive
information. (18)
The last configuration is the BUS architecture. The central processor and all components
are all connected via a network BUS. This results in a reliable system that can implement fault
tolerance and redundancy with limited impact to physical constraints such as size, weight, and
power. (18) The disadvantage is that the system’s throughput is bottlenecked by the size of the
BUS network. The BUS architecture is the most fitting candidate because the amount of data to
compute is large, as well as multiple computers will be used. (10)
Table 3 illustrates a trade study on possible computers that have all been on previous space
flight missions. The key aspects in determining if the correct computer for the mission include the
word length, memory, performance, radiation hardness, connectivity, heritage, and price. Each of
these characteristics were given a weight based on the conditions of the mission. A higher weight
indicates a bigger importance for that characteristic.
20
Sup
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er
ISA
Wo
rd L
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(b
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Me
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(M
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Rad
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# o
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Pri
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do
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s)
Ho
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300
032
4 gb
2010
0000
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34.
60E+
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L-M
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SC17
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1616
mb
210
0000
03
11.
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L-M
Rad
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16 G
B20
1000
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54.
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SWR
I SC
-580
c386
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3232
0 K
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610
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1.32
E+01
SWR
I SC
-7TI
320C
3264
0 K
B12
1000
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13.
15E+
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SWR
I SC
-9R
S 60
0032
128
MB
2030
000
11
4.91
E+01
San
de
rs S
tar
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1050
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2.83
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Ch
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Rad
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55
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23
Co
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31
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31
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igth
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s5
35
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126
153
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3
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204
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2525
205
2010
15
66
62
22
2
62
82
22
2
820
416
124
12
Tota
l Sco
re82
7078
3760
5251
Table 3: Trade Study of Computer.
Candidates
21
There is a risk of radiation causing interference with the on board computer for this
mission. Radiation hardness needs to be taken heavily into account when choosing the on board
computer. The spacecraft will not have any protection from radiation due to not having the Earth’s
protection. The radiation hardness was the largest portion that was taken into consideration for
choosing the on board computer. Another important requirement is the performance of the
computer. The computer will likely have to process many commands at the same time. The
computer will be receiving a large variety of data from many different sensors. The computer will
have to process all of the data quickly in order to act in time. If the computer takes too long to
process the received data would be irrelevant. The memory of the computer is also important
whenever many commands need to be handled at the same time. It needs a large amount of random
access memory in order to process and use all the data. Price is always a large contributor for any
mission. A cheaper product was given a better score. The connectivity, heritage, and word length
were not considered to be as important so their weights were smaller than the other
characteristics. They weren’t considered as important because they are not directly related to the
main portion of the mission conditions. A trade study was conducted with a focus on these
attributes and the Honeywell RH32 currently the best candidate. (10)
The number of computers used on the spacecraft will be three. There needs to be at least
two computers so that there is enough processing power to handle all the data. There also needs
to be redundancy so that there are no failure points during the critical parts of the mission. If there
were a failure during the capture of the asteroid, the mission would be compromised and the time,
effort, and money put into it would be a waste. The added computer will help increase the success
rate of the mission. There is no human life aboard the spacecraft, which means that a backup to
the backup is not mandatory. (10)
22
An estimation of the number of lines of code needed for the software was conducted using
the methods in Space Mission Engineering: The New SMAD. Table 20-9 was used to determine
the amount of source lines of code for each aspect of the spacecraft. The components present in
the spacecraft where selected from the table and compiled onto Table 4. (18)
Table 4: Estimation of Source Lines of Code Needed for Spacecraft Software based on Space
Mission Engineering: The New SMAD method.
Computer Software Component Source Lines of Code
Executive 1000
Communication Command Processing 1000 Telemetry Processing 1000
Attitude/Orbit Sensor Processing Rate Gyro 800 Sun Sensor 500 Earth Sensor 1200 Star Tracker (output quaternion) 2000
Attitude Determination and Control Kalman Filter 7000 Error Determination 800 Precision Control 3500
Attitude Actuator Processing Thruster Control 1200 CMG (Rotating Disks) 1500
Fault Detection Monitor 4000 Fault Identification 2500 Fault Correction 5000
Utilities Basic Mathematics 800 Transcendental Mathematics 1500 Matrix Mathematics 2000 Time Management and Conversion 700 Coordinate Conversion 2500
Other Functions Momentum Management 3000 Power Management 1200 Thermal Control 800
Total Lines of Code 45500
23
The total amount of lines of code that needs to be developed is 45,500. The time it will
take depends on the number of people working on the code, the amount of funding, and the skill
level of the software engineers. The code should be developed to be working by 2017 in order for
a testing period of one year. Having the code completed a year in advance leaves time for
correction any problems in the code. (18)
3.g Guidance, Navigation, and Control
To capture an asteroid, the spacecraft will have to undergo a series of burns to reach the
asteroid and place it into lunar orbit. After being launched, the vehicle will continually spiral
outwards from Earth to reach an orbit similar to that of Earth’s. Over roughly two years, the
spacecraft will alter its orbit before it can rendezvous with the asteroid. This will be done using a
phase orbit. During this time, very few adjustments in regards to orbit and attitude control would
be needed. The capturing of the asteroid requires many or even constant attitude adjustments in
order to ensure a successful capture. After capturing the asteroid, the spacecraft will enter another
phasing orbit and make its way to the moon. Once the spacecraft arrives at the moon it will
continue to complete a series of thrusts to place the asteroid into a retrograde lunar orbit. Lunar
flybys will assist in propelling the spacecraft to the asteroid. Control is essential when
rendezvousing with the asteroid and returning it.
The vehicle will be semi-autonomous. It will be able to perform all major decisions and
thrusts on its own. Ground control will be responsible for minor adjustments as needed and to
repair any bugs. Ground control will receive periodic updates of the spacecraft and the success of
each task. The spacecraft will receive confirmation from ground control that it is safe to capture
the asteroid based on simulations on measured characteristics of the asteroid.
24
The spacecraft’s location must always be tracked. The spacecraft and asteroid location are
essential to completing the task. NASA has Near-Earth Object Program (NEO) which has tracked
various asteroids, including the intended target. To track the spacecraft, Microcosm Autonomous
Navigation System (MANS) and Earth and Star Sensing systems can be used. (8) Both systems
allow for autonomy and planetary orbits. These systems are more accurate compared to other
systems that could be used. Accurate tracking of the spacecraft is a key requirement for a
successful mission, especially whenever the asteroid is being captured. Both MANS and Earth
and Star Sensing systems will be used to ensure redundancy in case of a failure. These will be
used to make sure the spacecraft can be accurately tracked at all times. The Earth sensor will also
be used to guide the antennas to point accurately to ground control. Sun sensors will also be used
to determine the spacecraft’s orientation relative to the sun. The sun sensors will also be used to
determine which direction to face the solar panel arrays.
Relative position and distance between the asteroid and spacecraft is vital for proper
control. The spacecraft will need to know how to alter its velocity upon arrival. One way to detect
this is using laser Doppler velocimetry. The centers of the spacecraft and asteroid need to align
for a proper capture. Image processing or active sonar will aid in alignment. This device will also
determine the asteroid’s spin rate and spin axis. It is important to know these two characteristics
so that proper measures can be taken to capture the asteroid. The capturing mechanism needs to
approach the asteroid along the asteroid’s spin axis so that it will be easier to despin the asteroid. If
the spin rate is known, it is also easier to estimate how much action and fuel is needed in order to
despin the asteroid. The gathered information from this device will also be sent to ground control,
so that ground control can simulate the capturing action to make sure there are no
25
problems. Simulations are helpful to predict how the spacecraft will react to this key moment in
the mission.
Attitude control is another key system while capturing the asteroid. The spacecraft must
keep its heading and not rotate while approaching the asteroid. While in the process of capturing
the asteroid, it is important for the spacecraft to be aligned with the center of the asteroid. The
attitude control system will have to keep the spacecraft aligned with the center during this phase.
The IMU will be used to detect changes in attitude and vernier thrusters will correct for the
changes. The equipment must be able to provide large torques to counteract any forces caused by
the asteroid. A system of moveable mass will also be needed to correct the inertia the asteroid adds
to the spacecraft asteroid system. The asteroid’s mass is not perfectly distributed throughout it,
which would cause a change in the spacecraft’s inertia matrix. Two of the cells will have moveable
masses. For the beginning of the mission these will be equal and opposite to keep the center of
mass stationary. These will react to the asteroid to keep the center of mass aligned for the thrusters.
The IMU being used on the spacecraft is the LN-200S from Northrop Grumman. The LN-
200S has a long heritage of use on spacecraft. It was also has high accuracy while being light
weight. There will be three IMU’s on board. (26) Ball Aerospace’s FSC-701 will be used for star
tracking. Ball has a large history of successful sensors making it a reliable choice. This sensor
has a wider range of possible uses so software can be updated to suit later needs if issues arise.
The spacecraft will use 2 star trackers. (25) Fine Sun Sensor (FSS) from Jenaoptronik will be the
sun sensor. This is a very small sensor and passively powered so it will not tax the system to find
the direction of the sun. There will 6 sun sensors on the spacecraft. (24) The earth sensor on board
will be the Earth Sensor Assembly by NEC Toshiba Space Systems. (27)
26
3.h Power
Numerous different power supplies have been discussed, including: Radioisotope
Thermoelectric Generators (RTGs), Nuclear Fission Reactors, Solar Panel Arrays, and other
experimental technologies have been explored, including Low Energy Nuclear Reactors (LENR)
and Kinetic Fluid Stabilized Nuclear Reactor (KIFSNR) which is a modified version of a ground
based fission reactor with systems enabling operation in outer space. (3) The reason for so many
power sources have been considered is due to the power requirements for different propulsion
systems. Taking into account for almost any space mission using ion propulsion, the propulsion
system consumes the majority of the power. After much discussion on what propulsion system will
best suite this mission, it has currently been determined that Hall Effect thrusters will serve as the
spacecraft’s main propulsion system. The specific type of Hall Effect thruster that is being
considered uses up to 25 kW of power per thruster, which presented an enormous road block when
it comes to the power system. (3)
The best technology to supply that amount of electrical power is currently solar panels. For
this mission, a solar panel array has been designed that will provide an estimated 26.26 kW of
power per array. Two arrays are currently a part of the spacecraft design, which would supply an
estimated total of 52.52 kW of electrical power for the Hall Effect thrusters and other sub-systems
alike. These calculations are approximations, but have been calculated based on the type of solar
panels which were used in the Deep Space 1 technology validation mission and have shown the
capability to provide this amount of power for the mission’s solar panel area. The design of the
solar panel array (SPA), is intended to maximize the area, but still allow the panels to be stowed
in order to minimize the space they take up during the launch of the vehicle. The current design is
shown in Figure 11. There are eight rectangular panels with six individual solar panels per disk
27
shaped array. There are also two triangular panels hinged on each rectangular panel that fold onto
each side for launch, for a total of 16 triangular panels per SPA. Multiple high capacity Lithium
Ion Batteries will store the energy generated. The SPA design is presented in Figure 11. (14)
Figure 11: Solar Array. Based on SCARLET Solar array used on Deep Space 1. Using the
NASA Technology Validation Report it has been determined the solar array could produce 26.26
kW of power. There would be two of these arrays on the spacecraft that will have that ability to
rotate completely around the spacecraft using a special direct drive electric motor. (14).
28
Figure 12: Lithium Ion Battery.
The power supply system was designed mostly around the propulsion system as previously
mentioned. The SPA will power four Hall Effect Thrusters for the main propulsion system each
thruster could use 12.5 kW of power at a time if all 4 are being fired, or two thrusters could be
fired at a maximum of 25 kW each, which is the limit of power that can be used.
29
The SPA can be rotated 180˚ in the x-direction (about the z-axis) from the origin of the SPA
coordinate frame, using motor-1. The entire coordinate frame can be shifted 180˚ about the y-axis
relative to motor-2’s coordinate frame, allowing the array to be moved ‘up or down’. In addition
to the motions described, both arrays can be shifted 360˚ about the entire spacecraft’s x-axis, (can
rotate around the exterior of the spacecraft’s hull). Enabling the SPA’s to move in all of these
directions was designed to ensure that the panels always receive the maximum amount of solar
radiation possible given the spacecraft’s relative position to the sun. All of the motors that move
the SPAs are direct drive motors similar to the design shown in Figure 13. Using a modified direct
drive motor to move the SPAs will proved a significant advantage over traditional electric DC or
AC motors. The most important advantage is an increased efficiency, due to the removal of power
waste from friction.
Figure 13: Example of Direct Drive Electric Motor (22).
1 kW of power will be allocated for the electrical motor use to move the SPA’s. Allocating
1 kW of power should be more than sufficient to move the SPA into an adequate position, because
the motors do not need to move fast and they will not require a constant supply of power. 1520 W
30
have been allocated for computer systems, communications, claw operation, sensors, scanners, and
any other control systems that require electrical power. Once the power requirement calculations
have been evaluated more thoroughly these allocations may change
3.i Thermal
The ability for the satellite to perform in a wide range of temperature differences is
extremely important for the success of the mission. The ability for the rocket to also maintain
proper stability during launch is also dependent on the rocket itself to withstand the temperatures
produced with the main thrusters active.
Once the rocket has escaped the atmosphere, the outside temperature will drop dramatically
which was taken into consideration. On the low end of the extreme temperature spectrum, circuits
and other parts of the electronics that were based on semiconductors were shown to still be
operative even at temps as low as -55 ⁰C. Due to this constraint, a coating of Hughson White Paint
A-276-1036 will be used on the outside of the spacecraft (13). This will create an ambient
temperature inside the spacecraft of roughly -24⁰C (18). The ambient temperature inside the craft
does not take into account the heat output of the internal systems on board. However, through
many calculations and estimations, all electronics on board contain internal temperature that will
be close to the freezing temperature of water, 0⁰C.
The ambient temperature of the spacecraft without the electronics on will allow for the
circuits and transistors of the computers to run correctly, the batteries however, would be best
preserved at a temperature of around 21⁰C. This is due to the self-discharge rate of the batter
increasing with an increase in temperature. The self-discharge rate is the amount of energy that is
lost naturally because of the reaction occurring the battery. At 21⁰C, the discharge rate is roughly
8% per month (16). This is an optimal level so that the battery loses the least amount of charge,
31
while still being in an environment that will allow it to function. In order to maintain this
temperature, heating pads will be placed on the batteries to keep them at this minimal temperature.
The pads will be made of kapton with nichrome wires running through them. A battery heat sensor
will be used to turn the heat pad on and off to maintain this minimum temperature (3).
The electronics, the propulsion system, as well as the satellites position in the orbit with
respect to the sun, can greatly increase the internal temperature of the satellite which may cause
problems which have been taken into consideration. In the current design of the spacecraft, there
is sufficient space between the thrusters and electronics to create a buffer from the heat produced
by the propulsion system. The nacelle that contains the electronics also requires a type of
refrigeration to keep the temperature inside the satellite at a certain temperature range due to the
energy produced by the sun.
Two energy systems were discussed, a radioisotope thermoelectric generator and a solar
panel system utilizing lithium ion batteries. Both energy systems release heat while in operation
and will need to utilize a cooling system in order to maintain stability. With the number of heat
sources on the spacecraft, a diphasic loop will need to be implemented as the cooling system (1).
The coolant that will be utilized is ammonia due to its intrinsic properties. The cooling system will
consist of copper tubing running the length of craft containing the electronics and ion batteries, as
well as a compressor and a pump. Heat sensors will be placed on the computers on board as well
as other electronic sensors to autonomously maintain an ambient temperature within the hull of
roughly 21⁰C to preserve an optimal temperature for all the onboard systems.
32
3.j Scientific Instruments and Other Payloads
There are two additional payloads that the spacecraft has to carry besides its own
components. These two payloads are the asteroid capturing claw-harpoon mechanism and the
asteroid 2009 BD. Each of these creates forces and torques that affect the spacecraft.
The claw is one of the most vital parts for this mission to succeed. If the claw fails the
asteroid cannot be captured making the mission a failure. The claw will be a four pronged
mechanism that will encapsulate the asteroid. It will be designed with several joints so that a
secure fit around the asteroid can be achieved. Each prong of the claw will also be connected to
its surrounding prongs by a tether to ensure that the asteroid cannot slip out from in between
them. It was designed to be 35 meters in length and made of titanium. This was chosen because
it is the same length as the maximum circumference that the asteroid can possess. The asteroid is
not a perfect sphere, but these factors of safety have to be used so that the mission can succeed. If
the spacecraft arrived at the asteroid and could not fully wrap around the asteroid the mission
would be a failure. The claw is also designed to be compact so that the total length of the spacecraft
system will not be too long and so it can fit onto the launch vehicle comfortably. A claw was
chosen to be used because it can be used to capture a variety of different small asteroids as well as
different shaped asteroids. The claw will also be able to keep the asteroid safe, tethered, and locked
in place because the claw is rigid.
Other possible capture methods that were considered include an asteroid trash bag, drill,
and ram. The asteroid bag was not chosen because it would be harder to access the asteroid to
study once the spacecraft is in orbit around the moon. The drill was not chosen because it is not
certain if the drill would be able to penetrate the asteroid because the substance of the asteroid is
not known. The ram was not chosen because it would require a high impulsive velocity change
33
that would require a large amount of fuel. The propulsion system would not be able to create an
impulsive maneuver of that magnitude.
The claw will also have a backup system in the event that it could not successfully capture
the target asteroid. There will be a net like structure made of a composite fiber woven between
the prongs of the claw. If the claw cannot sturdily hold the asteroid in place the net will still be
able to confine the asteroid to a controlled area. This backup system is a form of redundancy and
fixes the major failure point of the mission.
Another backup system was determined to be used. A harpoon system will be utilized if
the claw mechanism fails. Harpoons will make up the tips of the claw prongs. These harpoon tips
will be able to shoot from the prongs into the asteroid. The tips will be made of titanium so that it
can easily pierce the porous asteroid. If for some reason the claw mechanism opened but can no
longer close, a harpoon located at the center of the claw can be used to fire at the asteroid. This
harpoon will also have a titanium tip. The harpoons will not be used unless absolutely necessary.
The asteroid that is going to be captured is asteroid 2009 BD. This asteroid was chosen
due to many characteristics that make it an optimal choice. The asteroid has a diameter between
4 and 11 meters. This is not a large asteroid so it makes it easier to capture. It also has a relatively
low excess hyperbolic velocity compared to other asteroids of its size at only 1.2 km/s. The
asteroid also has a density that is slightly less than water. (7) This suggests that the asteroid is
made of a very porous material similar to pumice. The low density would mean it has a lower
mass than other asteroids of its size, which would lessen the loads it exerts onto the spacecraft
system. The asteroid also has a similar orbital path to Earth. The semi major access of the
asteroid’s orbit is 1.0089 Au which is very similar to Earth. The eccentricity of its orbit is only
0.041 which is not that much different than the 0.0167 of Earth’s. The inclination difference
34
between the ecliptic plane and the asteroid is only 0.3847 degrees, which makes it easy to approach
the asteroid. These characteristics makes it so that there is not a large delta V needed to get to and
from the asteroid when compared to other candidate asteroids. (9)
If asteroid 2009 BD cannot be captured for any reason, the spacecraft can move to a
different near Earth asteroid. Other possible candidates include 2011 MD, 2008 HU4, and 2013
EC20. These asteroids are all relatively close to Earth and 2009 BD. They also range from a
diameter of three to seventeen meters, which is similar to the size of 2009 BD. Any of these
candidates could be captured and if one had to be chosen it would be based on whichever was
closest and requires the least amount of propellant expended at the current time of the mission.
4. Table Summaries
Table 5: Mass Budget for the Mission.
Mass Budget
% of dry mass
mass (kg)
Payload 19.06% 3300
Structure 31.76% 5500
Thermal 0.29% 50
Power 26.68% 4620
TT&C 0.26% 45
Computer 0.27% 46
GNC 0.38% 65
Propulsion 13.51% 2340
Propellant 63.53% 11000
Other 7.80% 1350
Total 163.53% 28316
Margin 24684
The total mass of spacecraft was estimated to be approximately 28,000 kg and each
subsystem’s mass can be seen in Table 5. Of that weight, propellant was estimated to be around
60% of the total mass. The rest of the subsystems were estimated using the methods in Space
35
Mission Engineering: The New SMAD. The estimates were either based on typical subsystem
weights used in similar missions or estimations using the size of each subsystem. A margin of
about 25,000 kg was found based on the capabilities of the launch vehicle. This gives plenty of
room for error or unseen changes that would be needed during the building process.
Table 6: Power Budget for Mission.
System
Power
%
Total Power
(W)
Propulsion 0.8 42016
Electromechanical 0.1 5252
C&DH 0.02 1050
GNC 0.03 1576
Comm. 0.05 2626
Total 1 52520
The solar power arrays produce 52 kW of power. Of this power, the propulsion system is
predicted to use the majority of the power. The electromechanical systems include mechanisms to
control the claw, movable mass, and reaction wheels. The remaining power is either used between
C&DH, GNC, and communications or is excess power. Other than propulsion, all subsystems are
over estimated by a significant amount to ensure a safe margin of operability. The power budget
is further detailed in Table 6.
Table 7: Cost Budget Estimations for Mission.
Payload Non-recurring
Cost($K) Recurring($K)
Structure 250800 4674
Thermal 2280 912
Power 235620 6400
TT&C 7425 3712
Computer 8997.6 4500
GNC 21577.4 10784
Propulsion 206388 3308
Launch Vehicle 135000
36
Propellant 26400 Developmental
Cost 464389
Total Cost 1358877 34290
Combined Cost 1393167 2013 FY
Table 7 provides the estimated cost of the mission and the costs for each subsystem. The
total cost of the mission was estimated to be about $1.4 billion. The estimations were calculated
based on the weight of the subsystem and the average cost of the subsystem per kilogram. NASA
has a budget of approximately $2.6 billion which shows these estimations are valid.
5. Conclusion
The work done has ended in creating a spacecraft that will be able to travel to an asteroid
and return it to a lunar orbit. The overall structure has been further defined in shape and size as
well as the allocation of power required by the various systems. The launch vehicle was decided
on estimates of the total payload.
The MCC, POCC, and SOCC will all be stationed at JPL in Pasadena, CA due to its
resources available. The X and K bands are the frequencies that the communication system will
use. The Deep Space Network will receive the communication signals and relay the messages to
ground control. The bus architecture was chosen due to its quicker response time and its ability to
add multiple computers which is desired for this mission.
Thruster type and location were determined to maximize control of the vehicle. An IMU,
star sensor, Earth sensor, sun sensor, and laser Doppler velocimetry will be used to ensure the
asteroid is in the correct position and can capture the asteroid correctly. It was determined that a
rechargeable power system by means of solar panel arrays would be the optimal choice due to the
37
needs of the spacecraft because the duration and nature of the mission. The capturing mechanism
was designed to ensure a successful capture regardless of the asteroid’s dimensions.
Much thought has been put into the type of thermal control system that will be needed for
the spacecraft. The decision to include both a type of heating and cooling system would be best
to ensure an optimal temperature inside the spacecraft for all onboard electronics to function
properly by means of a radiation fins, heating pads, and a diphasic loop.
The design will optimize the process of creating a spacecraft that will be able go into deep
space and return celestial bodies of varying sizes. This goal has been a driving force in the decision
making. This line of thought has brought about original ideas such as movable masses inside the
spacecraft to move the center of gravity as well the claw capturing mechanism.
The spacecraft can possibly be used in future missions depending on its condition after it
has completed its first mission. If the components of the asteroid are still intact and operable, a
decision can be made if it needs to be refueled and/or if it is worth it to refuel. However, enough
propellant has been placed on board to safely accomplish three similar sized missions with a
margin of about 75% of a single mission. The spacecraft can then go on and capture a different
candidate near Earth asteroid. If it is not in operable conditions, the spacecraft can be separated
into its modular parts and then be burnt up in the Earth’s atmosphere.
38
6. References
[1] Bensaada, M., R. Roubache, A. Bellar, and L. Boukhris. "Satellite Thermal Control: Cooling
by a Diphasic Loop." Engineering and Technology 289th ser. 59.59 (2011): 1524-526.
Waset.org. World Academy of Science. Web. <http://www.waset.org/journals/waset/v59/v59-
289.pdf>.
[2] "Delta IV." United Launch Alliance, LLC. N.p., n.d. Web. 20 Oct.
2013.<http://www.ulalaunch.com/site/pages/Products_DeltaIV.shtml>.
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40
7. Appendices
7.a Spacecraft SolidWorks Models
Figure 14: Side View of Spacecraft.
41
Figure 15: Back Side of the Spacecraft.
42
Figure 16: Cross section of the Hull of the Spacecraft.
43
7.b Δv Calculation Code
%Characteristics rearth = 1*149597871; rast = 1.008995*149597871; ah = (rearth + rast)/2; r1 = 1.1*6378; musun = 1.327e11; muearth = 3.986e5;
Eh = -musun/(2*ah); % Energy of Hohmann Transfer
deltaV1 = sqrt(2*(-musun/(rearth + rast) + musun/rearth)) -
sqrt(musun/rearth); %Delta V Calculation of Hohmann Transfer deltaV2 = sqrt(musun/rast) - sqrt(2*(-musun/(rearth+rast)+musun/rast));
%Delta V Calculation of Hohmann Transfer
T = pi*sqrt(ah^3/musun); %Hohmann Trasfer Period EarthPos = sqrt(musun/rearth^3)*T; %true anomaly of Earth theta = acos((1^2 + 1.008^2 - .771^2)/(2*1*1.008)); %angle between Earth and
Asteroid
thetaSCS = theta - (EarthPos - 3.14); %Spacecraft is at Pi true anomaly
hast = sqrt(musun*rast); %Angular Momentum for the Asteroids Orbit
fdot = hast/(rast^2);
Tphase = (thetaSCS + 2*pi)/fdot; %Time of the Phase Orbit
aphase = (Tphase*sqrt(musun)/(2*pi))^(2/3); %semi major axis of the Phase
Orbit rphase = 2*aphase - rast; %Radius of the phase orbit
hphase = sqrt(2*musun)*sqrt(rast*rphase/(rast+rphase)); %Angular momentum of
the phase orbit
Vphase = hphase/rast; %Velocity of the phase orbit
Vast = sqrt(musun/rast); %Velocity of the asteroids orbit deltaV3 = Vphase - Vast; % Delta V to get on the phase orbit deltaV4 = deltaV3; %Delta V to get off the phase orbit
Vesc = sqrt(2*muearth/(1.1*6378)); %Escape Velocity Vcirc = sqrt(muearth/(1.1*6378)); %Velocity of Spacecraft in LEO
deltaV5 = Vesc - Vcirc; % Delta V to enter heliocentric orbit totdeltaV = deltaV1 + deltaV2 + deltaV3 + deltaV4 + deltaV5; %Total DeltaV
%To get the Delta V for the return orbit and Phase orbit replace the %apporpriate values in the phase orbit and Hohmann transfer Sections then %add to the total Delta V
44
7.c Link Budget
Table 8: Link Budget Estimations for Mission.
FSS Forward Link Cases* Main Antenna Backup Antenna Units
Uplink Frequency 34.400 8.500 GHz
Gateway Terminal Type Tracking Tracking
Diameter 70.000 70.000 m
Beamwidth 0.009 0.035 deg
Antenna Efficiency 0.550 0.550 %
Gain 85.437 73.294 dBi
Transmit Power 200.000 200.000 W
Backoff and Line Loss -6.000 -4.000 dB
EIRP, Gateway 102.447 92.304 dBW
EIRP per user 87.006 76.864 dBW
Propagation Range 150000000.000 150000000.000 km
Space Loss -286.680 -274.537 dB
Atmospheric Losses -10.000 -10.000 dB
Net Path Loss -296.680 -284.537 dB
Satellite Antenna, Type
Coverage Area 13.300 13.300 deg2
Antenna Efficiency 0.700 0.700 %
Gain 33.367 33.367 dBi
Line Loss on Satellite -2.000 -2.000 dB
Received Carrier Power Per User, C -178.306 -176.306 dBW
System Noise Temperature 27.200 27.200 dB-K
G/T 6.167 6.167 dB/K
Receiver C/No 23.094 25.094 dB-Hz
Data rate per user 80.792 80.792 dB-Hz
Available Eb/No, Uplink -57.698 -55.698 dB
Downlink Frequency 8.750 7.200 GHz
Satellite Antenna, Type 3.5x2.4-m shaped
Coverage Area 13.300 13.300 deg2
Antenna Efficiency 0.700 0.700 %
Antenna Gain 33.367 33.367 dBi
Satellite Tx Power 20.000 20.000 W
Backoff and Line Loss -5.500 -4.500 dB
EIRP per Line Loss 40.877 41.877 dBW
EIRP per user 25.437 26.437 dBW
Propagation Range 150000000.000 150000000.000 km
Space Loss -274.789 -273.095 dB
Atmospheric Losses -7.000 -7.000 dB
Net Path Loss -281.789 -280.095 dB
User Terminal Type
Diameter 1.500 1.500 m
Beamwidth 1.600 1.944 deg
45
Antenna Efficiency 0.550 0.550 %
Gain, G 40.166 38.472 dBi
Line Loss -2.000 -2.000 dB
Receive Carrier Power Per User, C -218.186 -217.186 dBW
System Noise Temperature 27.000 27.000 dB-K
G/T 13.166 11.472 dB/K
Receiver C/No -16.586 -15.586 dBW
Data Rate per user 80.792 80.792 dB-Hz
Available Eb/No, Downlink -97.378 -96.378 dB-Hz
End-to-End Eb/No 97.379 96.379 dB
Modem Implementation Loss -1.200 -1.200 dB
Required Eb/No 0.730 0.730 dB
Link Margin 3.000 3.100 dB
Channel Bandwidth 36.000 36.000 MHz
Number of Channels 24.000 24.000
Number of Users/Channel 35.000 35.000
Single User Data Rate 120.000 120.000 Mbps
Code Rate, ρ 0.400 0.400
Single User Bandwidth 201.000 201.000 MHz
Bandwidth Used/Channel 7035.000 7035.000 MHz
Total Capacity 100800.000 100800.000 Mbps
The entries in the link budget in Table 8 were determined by either antenna
characteristics, known distances, suggested by Space Mission Engineering: The New SMAD, or
were the default values already inside the table.
46
7.d Battery Trade Study
Table 9: Battery Trade Study. Chosen battery is colored blue.
Cell
Type
Supp
lier
Mod
elCa
paci
ty(A
h)En
ergy
(Wh)
Ener
gy D
ensi
ty (W
h/L)
Spec
ific
Ene
rgy
(Wh/
kg)
Li-I
onG
S Yu
asa
LSE
Gen
III
134
496
349
155
Li-I
onG
S Yu
asa
LFC
4044
163
244
102
Li-I
onQ
ualli
onQ
L075
KA72
269.
3634
114
8
Li-I
onQ
ualli
on12
V-Q
LD01
96-M
T19
621
5639
717
3
Li-I
on (C
oO)
Eagl
e-Pi
cher
SAR-
1019
720
066
6024
010
4.9
Supp
lier
Nom
inal
Vol
tage
(V)
Max
. Con
t, D
isch
arge
Cur
rent
(A)
Cell
Wid
th (m
m)
Cell
Thic
knes
s (m
m)
Cell
Hei
ght (
mm
)Ce
ll W
eigh
t (kg
)Te
mpe
ratu
re R
ange
(°C
)
GS
Yuas
a3.
713
413
050
271
3.53
-10
to 3
5
GS
Yuas
a3.
710
013
731
165
1.6
-5
to 3
0
Qua
llion
3.6
100
8156
.217
3.7
1.82
15
to 3
5
Qua
llion
1190
024
1.3
254
203.
225
.8
-1
8 to
60
Eagl
e-Pi
cher
33.3
154
269.
2474
9.3
264.
1663
.5
-1
5 to
40
47
7.e Propulsion System Capabilities
Delta-V that can be produced by 5 Hall Effect Thrusters:
∆𝑣 = 𝑣𝑒 ln (𝑚0
𝑚1) ; 𝑣𝑒 = 𝐼𝑠𝑝 ∗ 𝑔
𝑚0 = 28000 𝑘𝑔 , 𝑚1 = 17000 𝑘𝑔
𝐼𝑠𝑝 = 6500 𝑠 , 𝑔 = 9.81 𝑚/𝑠
∆𝑣 = 6500 𝑠(9.81 𝑚/𝑠) ∗ ln (28000𝑘𝑔
1700 𝑘𝑔) = 31818.2 𝑚/𝑠
∆𝑣 = 31.8 𝑘𝑚/𝑠
The Delta-V available for our mission based on the propellant mass and the main thrusters
specific impulse capabilities in 31.8 km/s. This is an exceptionally high Delta-V, however it was
agreed to over compensate so that incase asteroid 2009AD is not able to be captured, the mission
will not be a failure, the systems have the capability to produce an additional 23.8 km/s for
another 2 maneuvers to potentially capture back up asteroids in a similar orbit.
