Oracle AERSP 401B
James Crawford
2
Table of Contents
ABBREVIATIONS 2
EXECUTIVE SUMMARY 3 MOTIVATION 3 PRIMARY OBJECTIVES 3 MISSION REQUIREMENTS 4 MISSION ARCHITECTURE 5 TIMELINE 5 MISSION SUMMARIES 8 COST ESTIMATE 12 POWER ESTIMATE 14 LINK BUDGET 15
SUBSYSTEMS 17 STRUCTURES 17 LAUNCH VEHICLE 19 PROPULSION 22 GROUND CONTROL 23 COMMUNICATIONS 25 COMMAND AND DATA HANDLING 27 GUIDANCE, NAVIGATION, AND CONTROL 28 POWER 29 THERMAL CONTROL 31 SCIENTIFIC PAYLOAD 36
CONCLUSION 39
APPENDIX 39
REFERENCES 46
Abbreviations
JPL-‐ Jet Propulsion Laboratory NASA-‐ National Aeronautics and Space Administration GRACE-‐ Gravity Recovery and Climate Experiment GOCE-‐ CSR-‐ Centre of Space Research of the University of Texas DLR-‐ German Space Agency
3
GFZ-‐ Germany’s National Research Center for Geosciences LEO-‐ Low Earth Orbit MECO-‐ Main Engine Cut-‐Off SECO-‐ Secondary Engine Cut-‐Off IBM-‐ International Business Machines BAE-‐ British Aerospace CER-‐ Cost Estimating Relationship MM/OD-‐ Micro Meteoroid and Orbital Debris
Executive Summary
Motivation
The mission of Oracle is to accurately map Earth’s dynamic gravity field
distribution. By observing changes in these measurements, the Oracle team can work to
define Earth’s geoid and come to understand the factors that control the fluidity of
Earth’s mass. The primary factor under investigation is the movement of Earth’s ground
water resources as well as changes to Earth’s solid land ice mass. This mission will
provide valuable insight into the changes in Earth’s water circulation and climate, which
have an impact on these measurements.
Primary Objectives
The primary objective of this mission is to design a satellite system to observe the
global gravity field and Earth’s mass distribution, as well as their variability over time,
such that we can accurately observe the movement of Earth’s water. Assuming water
movement accounts for the majority of the change in Earth’s mass distribution, a time-
varying gravity field measurement would show the large-scale disappearance of arctic
4
ice. Established secondary objectives include defining of the shape of the geoid as well
as improving upon the accuracy of similar mission, such as GRACE and GOCE.
Along with the mission objectives, there is a set of requirements that limit the mission
design decisions. Table 1. depicts the requirement matrix that must be taken into account
in the realization of this mission.
Mission Requirements
Table 1. Mission Requirements Requirements Factors that normally impact the requirements
Functional Requirements
Performance Approximately 210W of power Coverage Coverage in the whole global sphere
Interpretation Uplink Signal: 2051.0 MHz Downlink Signal: 2211.0 MHz
Timelines Less than 10 seconds of lag Secondary Missions
Resolution up to 100 km
Operational Requirements
Commanding Entire system commanded in selected ground station Mission Design Life 4 years System Availability 99% availability
Survivability Space minor debris collision survivable, overheat due to Solar radiation survivable.
Data Distribution Centre of Space Research of the University of Texas (CSR), Jet Propulsion Laboratory, German Space Agency (DLR) and Germany’s National Research Center for Geosciences (GFZ)
Data Content, Form and Format
Microwave Signals
User Equipment Signal Receivers and Processors Constraints
Cost $450 million USD (FY 06) Schedule Operational within 4 years
Risk Battery early degradation Regulations Orbital debris, Data Security Regulations
Political Questionable political relationship with Russia Environment Natural
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Interfaces Usable by researchers in Geodesy, Glaciology, Hydrology, Oceanography and Solid Earth Sciences
Development Constraints
N/A
Mission Architecture
Timeline
The following figure details the order of events for this mission. The launch
vehicle will launch from Vandenberg Air Force Base carrying the two-satellite system
south over the Pacific Ocean. The launch vehicle will deliver the satellite system to LEO
with an altitude of 425.8 kilometers and an inclination of 89 degrees.
Figure 1. Mission Launch Timeline
6
Figure 2. Launch Overview
Once the launch vehicle reaches this altitude, the two satellites will be released
simultaneously. As described in the structural subsystem, the twin-satellites are designed
to fit together to conserve space in the payload fairing. Once released Satellite 1 will
separate from Satellite with a relative velocity of 0.5 m/s. As the satellites separate, on-
board systems will activate and the satellites will orient in their orbit. Immediately
following this, communication-links will be established between Ground Control and the
two satellites. To conserve on-board propellant, the satellites will be allowed to orbit
with their marginal difference in velocity until a separation distance of approximately 200
kilometers is established. At this point, communications will allow for continuous
downloading of the data. The operational period of this satellite system is determined to
be 4 years at a minimum. Following this four-year period the satellite system will
undergo a controlled re-entry and burn up in the atmosphere.
7
Figure 3. Satellite Separation Timeline
As the Oracle satellites orbit Earth and the mission enters it’s observational face,
the scientific payload instruments will begin to record the data points needed to generate
a spherical map of Earth’s gravitational potential governed by the equation:
€
U =GM
r+
GM
r(Re
r)n Pnm (sinϕlat )[Cnm cos(mλ) + Snm
m =0
n
∑n =0
∞
∑ sin(mλ)]
Where R is the equatorial radius of Earth, M is the mass of the Oracle satellite, r is the
radial distance coordinate, m is the order, Pnm are normalized associated Legendre
polynomials, Re is the reference radius of Earth, φlat is the latitude, and λ is the longitude.
The gravity coefficients in this equation are normalized and related to the Kaula
coefficients, Cnm and Snm, where delta is the Kronecker delta:
€
Cnm
Snm
⎛
⎝ ⎜
⎞
⎠ ⎟ =
(n −m)!(2n +1)(2 −δ0m )
(n + m)!
⎡
⎣ ⎢ ⎤
⎦ ⎥
1/ 2Cnm
Snm
⎛
⎝ ⎜
⎞
⎠ ⎟ = fnm
Cnm
Snm
⎛
⎝ ⎜
⎞
⎠ ⎟
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Oracle will account for tides caused by the Moon and Sun using the following equation:
Where Mp is the mass of the perturbing body, rp is the position of the perturbing body,
and k2 is the second degree potential Love number.
€
U = k2
GM p
R
R6
r3rp3
3
2(ˆ r • ˆ r p )2 −
1
2
⎡
⎣ ⎢ ⎤
⎦ ⎥
By considering the effects of these perturbing bodies, Oracle will be able to define
changes in the values of J2, C21, S21, C22, and S22.
Oracle’s gravity field estimations are determined using two data sources, the S-
band link between the satellite system and Earth, and the inter-spacecraft K-band. The
inter-satellite link allows for the precise measurements of the relative movement of the
satellites, which leads to the estimation of Earth’s gravity field. At the same time, GPS
data allows mission control to track the absolute position of each satellite. Using the data
recorded by these instruments, Oracle software will integrate the differential equation:
€
r..
= f (r,v,q) =∇U(r) + f pm + f in− pm + f in−orl + f srp + f alb + f att + f rel + • • •
To obtain the total acceleration of the spacecraft, and after filtering out unnecessary
noise, the gravitational acceleration of the satellite system due to Earth’s gravity
anomalies.
Mission Summaries
At this point, the spacecraft will be equipped with MLI coatings and a set of 8-mil
quartz mirror s as surface finish. These decisions are due to the low absorptivity of the
quartz with respect to other surface finish alternatives. In the circuit board of the satellite,
bonding straps will be located in the batteries and other power generating components to
prevent abrupt changes in power or sparks. The Thermal Control Team determined that
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the frequencies in the Communications Subsystem must not approach the critical
frequencies of 4.5 MHz and 11 MHz in nighttime and daytime instant in orbit. Moreover,
a plasma contactor will be located in strategic regions of the surface to counteract the
electromagnetic effects of plasmas currents. Batteries, computers and other yet
undetermined components will be provided with heat to maintain their operational
temperature ranges with active heaters. For future work, the Thermal Control Team will
evaluate the regions that will contain contactors, bonding straps, and heaters. Also, the
Thermal Control Team will explore more widely the effects of plasmas in polar, regions
of Earth due to the change in orbit.
The Delta IV Medium was chosen to carry the twin satellites to orbit. This choice
was made primarily on the availability and reliability of the Delta IV rocket. The Delta
IV’s payload capacity vastly over exceeds the mass estimation of this mission, this excess
will be offset by selling available space in the rocket’s payload fairing to other missions.
The launch vehicle will launch from the Space Launch Complex 6 at Vandenberg Air
Force base into a polar orbit.
From the primary altitude, attitude and orbit changes conducted throughout the
mission, four small cold gas thrusters will be situated in various places on each satellite.
There will be two thrusters placed on each end of the satellites with one controlling every
axis of the principal axes. The team has changed their selection from the Moog 58-126 to
Moog 58-125 thruster due to the smaller thrust, lower mass and power consumption. The
Moog 58-125 thruster still utilizes gaseous nitrogen as the fuel source. Each satellite will
contain two pressurized gaseous nitrogen tanks containing 35 kilograms each for a total
of 70 kilograms of fuel per satellite.
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The desired location for the mission control and spacecraft operations is in
California, as the Delta IV can be launched into polar orbit at Vandenberg Air Force
Base. This would allow for the use of the Jet Propulsion Laboratory in Pasadena. In the
1970s, JPL began to adapt previously used interplanetary missions sensors to Earth-
observation missions. JPL currently serves at the mission control for numerous
unmanned missions with a similar purpose to Oracle. GRACE, Oracle’s mission
predecessor, was controlled out of JPL as well as other global climate related missions.
The payload operations center will be located in the Neustrelitz Ground Station in
Germany. The Neustrelitz Ground Station is considered because of the facilities that
exist in Neustrelitz used by the first GRACE mission.
The following illustration summarizes the proposed ground control system.
Each satellite will contain one S-band antenna on the bottom of the spacecraft for
uplink and downlink. Assuming that the satellite’s attitude control system can hold the
satellite to a 32.5 degree pointing error, each antenna will require a maximum 28.421
watts of power. The satellite will transmit data at a maximum of 62 Mbps. Each antenna
will be 1.5 m in diameter so it can fit on the bottom of the satellite. The satellite will
downlink all science data to the Neustrelitz Ground Station in Germany using their 7.3 m
antenna as receiver. The communication team is also looking into a laser communication
crosslink for data transmission between the two satellites. The communications team will
also reevaluate the GPS receiver to choose the most viable option. The Viceroy-4 GPS
receiver made by General Dynamics was initially chosen as the GPS receiver however
the previous GRACE mission used a custom GPS receiver developed by JPL.
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For the data collection for the two-satellite system, a RAD750 computer will be
fitted to each satellite. This computer is the predecessor to the RAD6000, which was the
computer the team had selected last December. The new RAD750 was designed by IBM
and manufactured by BAE and contains 10.4 million transistors. Additionally, the
RAD750 can process 266 MIPS, withstand one million rads, is relatively small and only
consumes about 20 watts of power. Lastly, a trade study is displayed in the appendix
comparing the RAD750 to the RAD6000 and other space-hardened computers.
Each satellite will orbit with a circular, polar orbit. This means that the
eccentricity of the orbit will be 0 and the inclination will be 900. A polar orbit will
guarantee that the whole Earth will be surveyed especially the polar regions. A circular
orbit was chosen for simplicity and also because it will ensure equal coverage of the
Earth’s surface. If the orbit were too elliptical the resolution of the data would decrease,
as the satellite got farther away from Earth. The satellites will be launched to an altitude
of 429 km. This altitude was chosen because it is close enough to Earth to provide a high
resolution of the Earth’s gravitational field but far enough away where atmospheric drag
won’t cause the satellite to re enter. Both satellites will have identical orbits but the
satellites will be 220 km apart. For the future our team will be researching the previous
GRACE missions to see why they have inclinations of 890, a starting altitude of 500 km,
and have slight elliptical orbits with average eccentricities of approximately 0.002.
The power system has remained largely unchanged since December. The same
NeXt triple-junction solar cells are being used due to their high efficiency and low
degradation rate. Lithium ion batteries are still being considered due to their low mass
and large capacity to store power. Since the mission was switched from a sun-
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synchronous orbit to a polar orbit, this large capacity plays an important rule to the fact
that the spacecraft will be experiencing more eclipse time. Below is a schematic for the
triple-junction solar cells. The biggest progression in the power subsystem was the sizing
of the solar array. After all of the power requirements were calculated, it was a simple
task to size the panels knowing the orbits time in eclipse and time in daylight.
The decisions taken regarding the structure lead to the following: the width of the
spacecraft will be c.a. 1.73 m and the length will be approximately 2.8 m. The shape of
the spacecraft will be a heptagonal prism. The estimated weight in the heaviest case will
be of 450 kg. The spacecraft will be assembled will an Epoxy resin that will damp the
impact energy in case of MM/OD collision. As previously decided, the main metallic
material will be an Al-Li 8090 alloy and aramid fibers as the non-metallic materials.
The payload will consist of scientific observational equipment and sensor to
determine the changes in the Earth’s gravitational field. Both satellites will use the same
instruments that Grace I used. A combination of GPS configurations, accelerometers and
star cameras will be used to determine the locations of the satellites relative to each
other. The accelerometer is essential for recording change in motion of the satellites and
maintaining proper inertial reference frame changes in the system. The information
collected from the accelerometer is used for every gravitational calculation. Since there is
no “gravity” monitor that can be used to determine the force of gravity, these scientific
instruments are necessary for the gravity calculations and modeling.
Cost Estimate
The following cost estimate was performed using the QuickCost Non-recurring
plus Recurring (T1) CER for Space Vehicles for Unmanned Mission and was adjusted fro
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2010 USD to 2015 USD. The cost of this mission is determined using the following
equation:
€
Y = 2.829 × (DryMass0.457) × (Power0.157) × (2.718(0.171×Data%)) × (2.718(0.00209×Life ))
×(2.718(1.52×New )) × (2.718(0.258×Planetary )) ×1
2.718(0.0145×(Year−1960)) × (2.718(0.467×InstrComp%))
×1
(2.718(0.237×Team )
where DryMass is the mass of the spacecraft bus and instruments in kilograms, Power is
the LEO equivalent beginning of life power, Data% is the data rate percentile, Life is
advertised design life excluding extended operations, New is the new technology
adjustment, Planetary is the planetary adjustment, Year is the operation date adjusted to
1960, InstrComp% is the instrument complexity percentage, and Team is the team
experience.
These values are represented for the Oracle mission in the table below:
Table 2. Cost Drivers Cost Driver Value
Dry Mass 374.9 kg Power 285 W Date% 0.5 Life 60 months New 0.3 Planetary 0
(Earth mission) Year 56 InstrComp% 0.5 Team 3
(Normal experience) Using this estimation process, the cost of this mission is estimated to be a total of
$331.3 million USD.
Mass Estimate
Table 3. Mass Estimates
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Mass Budget
Subsystem Component Estimated Mass (kg)
SuperSTAR (Camera and Processor) 2.2
Laser Retro-Reflector 3
Star Camera Assembly 0.8
GPS 0.156
K-Band Antenna 1.48
Accelerometers 2.364
Science
SUBTOTAL 10
Solar Arrays 15.628
Batteries 5 Power
SUBTOTAL 20.628
RAD 750 Computer 10 Command/Data Handling
SUBTOTAL 10
Patch Heaters 0.05
Plasmas Contactor 10
MLI Coatings 0.1 Thermal
SUBTOTAL 10.15
Chassis 269.64 Structure
SUBTOTAL 269.64
Cold Gas Thrusters
Dry Weight 0.516
Propellant 89.484 Propulsion
SUBTOTAL 90
S-Band Antenna 30
Communications
SUBTOTAL 30
NET ESTIMATED MASS 440.418
Power Estimate
Table 4. Power Budget Estimate
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Subsystem Power Required (Watts) Command and Data
Handling 20
Thermal 100 Communications 90 Scientific Payload 75
Total 285
Link Budget
Downlink Budget
Antenna Property Unit Amount Comments
Dish Size m 0.1 For Main Satellite Antenna
Frequency GHz 2.2 S-‐Band
Beamwidth deg 95.4545
Pointing Error deg 10 Estimated
Pointing Loss dB -‐0.13170
Peak Xmtr Gain dB 4.66208 Efficiency of 0.55
Xmtr Gain dB 4.53038
Line Loss dB -‐1
Imp. Loss dB -‐2
Space Loss dB -‐175.940 Altitude of 425.9 km
Peak Data Rate Mbps 62 For Science Data
Slow Data Rate Mbps 0.5 For Orbital Data
Neutralitz Rcvr G/T dB/K 17 For Science Data
Wilheim Rcvr G/T dB/K 26.810 For Orbital Data using 15 m antenna
Required Eb/No dB 3.9
Estimated Eb/No dB 6.9
Peak Xmtr Power W 69.252 For Science Data
Slow Xmtr Power W 0.0583 For Orbital Data
Uplink Budget
Antenna Property Unit Amount Comments
Dish Size m 15 At Weilheim
Frequency GHz 2.12 S-‐Band
Beamwidth deg 0.6604
Pointing Error deg 0.05 Estimated
Pointing Loss dB -‐0.06879
Peak Xmtr Gain dB 47.8622 Efficiency of 0.55
Xmtr Gain dB 47.7934
16
Line Loss dB -‐1
Imp. Loss dB -‐2
Space Loss dB -‐175.618 Altitude of 425.9 km
Peak Data Rate Mbps 30 Estimated
Satellite Rcvr G dB/K 4.5304 Using S-‐Band antenna
Required Eb/No dB 3.9
Estimated Eb/No dB 23.9
Peak Xmtr Power W 801.030
Crosslink Budget
Antenna Property Unit Amount Comments
Horn Circumference m 0.31416 From Sat. 1 to Sat. 2
Frequency GHz 32 Ka-‐Band
Wavelength m 0.00938
Beamwidth deg 6.7143
Pointing Error deg 0.05 Estimated
Pointing Loss dB -‐0.00067
Peak Xmtr Gain dB 24.8636 Efficiency of 0.52
Sat. 1 Xmtr Gain dB 24.8629
Line Loss dB -‐1
Imp. Loss dB -‐2
Space Loss dB -‐169.401 Max Distance of 220 Km
Peak Data Rate Mbps 10 Estimated
Sat. 2 Rcvr G dB/K 22.3645
Required Eb/No dB 3.9
Estimated Eb/No dB 6.9
Peak Xmtr Power W 4.116
Antenna Property Unit Amount Comments
Horn Circumference m 0.31416 From Sat. 2 t oSat. 1
Frequency GHz 24 Ka-‐Band
Wavelength m 0.01250
Beamwidth deg 8.9525
Pointing Error deg 0.05 Estimated
Pointing Loss dB -‐0.00037
Peak Xmtr Gain dB 22.3648 Efficiency of 0.52
Sat. 2 Xmtr Gain dB 22.3645
Line Loss dB -‐1
Imp. Loss dB -‐2
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Space Loss dB -‐166.903 Max Distance of 220 Km
Peak Data Rate Mbps 10 Estimated
Sat. 1 Rcvr G dB/K 24.86294
Required Eb/No dB 3.9
Estimated Eb/No dB 6.9
Peak Xmtr Power W 2.316
Subsystems
Structures
For the structures subsystem, the most likely structural threatens that the system might
face are the collision with micrometeoroids and orbital debris (MM/OD) at high speed
and the physical damage to the circuit board and the spacecraft’s surface that radiation
can cause. Regardless of the size of the micrometeoroids, the effects of a collision with
the structure can be devastating. Near Earth (at less than 450 km), the flux of meteoroids
is about twice as much as in deep space. The flux of man-made orbital debris is even
higher. The effects of radiation over the spacecraft’s structure and the circuitry depend
on the solar activity cycle and many other factors. Radiation can break down the lattices
of materials in the circuitry as well as in the surface. This is why a surface finish with
high reflectivity and low absorptivity are required.
Due to the needs of the spacecraft to function despite the given environment, the set
of requirements for the structural subsystem goes as follows:
• Safety factor between 1.2 and 1.4
• Ability to fit in the launch vehicle’s fairing (less than 2.5 m of diameter)
• Surface finish with absorptivity less than 0.3 and emissivity higher than 0.7 and
low thermal conductivity.
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The spacecraft’s structure will consist of 3 main pieces for the chassis that holds the
rest of the components in each spacecraft. Eight faces of the structure will be covered
with solar panels of approximately 0.8 m2 each one, adding up to about 6.7 m^2 of area
covered by solar arrays. To counteract the effects of radiation and MM/OD in the
structure of the spacecraft, a surface finish consisting of 8-mil quartz mirrors, which has
an emissivity of 0.8, an absorptivity of 0.07 (which meet the requirements set), and also
has low outgassing rates. The main components of the chassis will be assembled using a
space rated EPOXY resin able to dissipate the incoming kinetic energy of an MM/OD
object. The metallic materials in the structure will be of Al-Li 8090 and the composites
used will be of aramid fiber. The shape and sizing of the spacecraft is described by Figure
4.
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Figure 4. Satellite Casing Drawing
Launch Vehicle
The requirements for this mission’s launch vehicle are dependent upon the
designed payload and the specific orbit utilized in this mission. The available launch
sites, performance, and launch success were also considered in a more limited capacity.
The launch vehicle must be able to launch a payload of 1,000 kilograms, to be launched
into polar, low Earth orbit, and must have a launch success percentage greater than 0.9.
Based on the above criteria and the determined requirements, multiple launch
vehicles were reviewed in a trade study. Launch vehicles were rated first upon their
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payload capacity. An ideal capacity of 1,000 and 2,000 kilograms was determined based
on previous missions of the same heritage. Performance to low Earth orbit was
considered next, and rockets were evaluated on a normalized scale. Available launch
sites were considered but are not present in the trade survey. Based on the
aforementioned requirements and the Delta IV M+ was chosen.
Figure 5. Delta IV M+ with 4-m Payload Fairing
The Delta IV is a two stage rocket capable of carrying a payload 13,140 kg
payload to LEO. The Delta IV also features two 1.5-meter diameter solid rocket strap-on
GEM-60s and delivers its payload in a 4-meter diameter payload fairing, but can also
utilize a larger 5-meter diameter fairing, illustrated in the figure below.
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Figure 6. Depiction of the launch vehicle payload fairings.
Available launch sites for the Delta IV include the Space Launch Complex 37 at
Cape Canaveral and the Space Launch Complex 6 at the Vandenberg Air Force Base.
Space Launch Complex 6 was chosen as the designated launch site due to the ability to
launch to a polar orbit.
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Figure 7. Map of Launch Trajectories out of Vandenberg Air Force Base.
The payload capacity of the Delta IV far exceeds the necessary capacity for this mission
alone. A large launch vehicle like the Delta IV can risk the mission going over budget.
After the cost analysis was completed, it was clear the Delta IV would not go over the
given cost requirements. In order to offset some of the launch costs and wasted space,
the rest of the payload fairing will be sold. Missions with similar orbit requirements were
researched. The high inclination of this mission allows for the inclusion of many types
reconnaissance, weather, and other earth observation satellites in the payload fairing.
Propulsion
For the second Gravity Recovery and Climate Experiment (GRACE) to be
successful, the orbital maneuvers must be conducted regularly and accurately. To fulfill
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this requirement, the Oracle team has switched from the Moog 58-126 to the Moog 58-
125 cold gas thruster. Like the previous selection, the Moog 58-125 utilizes gaseous
nitrogen as the fuel source. The switch was made to the Moog 58-125 thruster for
various reasons. For example, the team greatly overestimated the necessary thrust for
each satellite, which led to an inaccurate trade study. After consulting various sources,
the decision was made that such a large thrust was not needed and the trade study was
redone, which can be seen below. The Moog 58-125 has much better specifications for
the second GRACE mission. For instance, the Moog 58-125 has a vacuum thrust of
0.0045 Newtons and vacuum specific impulse of 65 seconds. Additionally, the nitrogen
cold gas thruster has a mass of 0.00734 kilograms and a low input power of just 2.4
watts. Five Moog 58-125 engines will be placed in various locations on each satellite.
Two engines will be placed on each of the ends of the satellites and one below. These
locations ensure that all attitude, altitude and orbital maneuvers can be completed
accurately and effectively. To fuel these thrusters, two 43-kilogram tanks of gaseous
nitrogen will be placed on each of the satellites for a total of 86 kilograms of fuel per
satellite. This should provide more than enough fuel for the desired duration and some
extra in case the satellites are still operational after the mission timeline. Overall, the five
Moog 58-125 engines are very efficient and will be able to perform all the necessary
propulsion maneuvers for the GRACE II satellites.
Ground Control
For this mission, the ground control system must handle all mission data. This requires
quick communication with multiple sources. Ground control must communicate with the
satellites to receive the distance changes between them. These distance changes can them
24
be used to determine gravity anomalies. Ground control must also communicate with
Global Positioning System instruments to determine the precise locations of the satellites.
This allows scientists to create an accurate map of Earth’s gravity anomalies.
A spacecraft operations, payload operations, and mission control center are
necessary for this mission. The spacecraft operations control center must monitor and
command the launch vehicle. The payload operations control center must analyze the
mission data from onboard payload instruments as well as issue commands to those
instruments. The mission control center must plan and operate the mission. Space for the
required control centers will be rented if available, if unavailable, space will be built.
The desired location for the mission control and spacecraft operations is in
California, as the Delta IV can be launched into polar orbit at Vandenberg Air Force
Base. This would allow for the use of the Jet Propulsion Laboratory in Pasadena. In the
1970s, JPL began to adapt previously used interplanetary missions sensors to Earth-
observation missions. JPL currently serves at the mission control for numerous
unmanned missions with a similar purpose to Oracle. GRACE, Oracle’s mission
predecessor, was controlled out of JPL as well as other global climate related missions.
The payload operations center will be located in the Neustrelitz Ground Station in
Wilhelm, Germany. The Neustrelitz Ground Station was chosen because of the GRACE
mission legacy. The previous GRACE mission used Neustrelitz Ground Station as its
payload operations center and all mission data was down-linked to its facilities. Wilhelm
will handle all mission sensitive information before sending it to JPL for analysis.
25
Communications
Figure 8. Communication Architecture
Figure 8 shows the communication architecture for Team Oracle’s mission. This
architecture is based off of the architecture of the first GRACE mission. Connections 1
and 2 are the satellites link to GPS. This connection will be how the two satellites
communicate with GPS satellites so each satellite gets an accurate position. This
connection will also be used as a backup method for determining the range between the
two satellites.
Connection 3 is the cross link between the two satellites. This link uses a laser
communication system to provide the link and a K-band horn antenna. This laser system
was chosen as an upgrade over the past GRACE mission which just used a K-band
microwave horn antenna to provide cross link communications. The K-Band unit is also
26
included to test the accuracy of the laser ranging system and also as a redundancy. The
upcoming GRACE-FO mission will also use both a laser ranging system and a K-band
horn. The two satellites will operate at slightly different frequencies to prevent the two
antennas from interfering with each other. The cross link will allow the two satellites to
communicate with each other so the satellites can back-up each other’s data and ensure
that a satellite experienced a gravitational anomaly. This cross-link will also double as the
satellites laser ranging system. A laser was chosen because it has a narrow beam width
and can send more data than a microwave antenna. The narrow beam width shouldn’t
cause problems because the satellites will always be facing each other.
Connections 4 and 6 are each of the satellite’s connections to the Weilheim
Ground Station’s 15 m antenna. This connection will be primarily used for sending orbit
and attitude data. The 15 m antenna was chosen because it was the biggest S-band
antenna at Weilheim besides the 30 m antenna, which is primarily used for deep space
missions. Weilheim was also chosen because the prior GRACE mission would downlink
orbital data there since the GRACE mission was a joint mission between NASA and
DLR. This downlink will use the S-band antenna located on the bottom of each satellite
or the backup patch antennas.
Connections 5 and 7 are each of the satellite’s connection to the Neustrelitz Ground
Station. This connection will be used for the downlink of scientific data. This is also a
legacy of the first GRACE mission. The downlink of orbital data and scientific data was
split up because the scientific data downlink will require more power since it is sending
larger amounts of data to Earth. The scientific downlink will also be done less frequently.
This connection will also use the S-band antenna on the bottom of the satellite.
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Command and Data Handling
The GRACE II mission will be collecting multitudes of data in order to generate
spherical harmonic representation of the gravitational potential of Earth. This data will
be collect from scientific instruments including but not limited to accelerometers, star
cameras, K-band ranging systems and GPS configurations. To collect and process the
data an onboard computer must be implemented into both of the twin satellites.
Previously, the Oracle team selected the RAD6000 computer. After research and
consultation, the team decided to switch to the RAD6000’s successor, the RAD750
computer. A trade study, which is shown below, was conducted to ensure the RAD750
computers would in fact be better for the GRACE II twin satellite system. IBM designed
the RAD750 while BAE Systems Electronics manufactured the radiation-hardened single
board computer. Furthermore, the RAD750 can withstand up 100,000 rads, which is
more than enough to prevent damage to computer when the satellites are unprotected by
Earth’s magnetic field at the poles. Additionally, the computer and motherboard of the
RAD750 can withstand temperatures from -55 degrees Celsius to 70 degrees Celsius.
The RAD750 also has almost ten times more transistors than its predecessor with a total
of 10.4 million transistors. Another major benefit of the RAD750 includes the fact that
the computer and motherboard only require about 10 watts of power at any given time
while still being able to perform 266 million instructions per second. The RAD750 has
an excellent heritage with over ten successful missions including Mars Reconnaissance
Orbiter and the Kepler Space Telescope. As a whole, the RAD750 is the optimal choice
for the GRACE II satellites with the ability to effectively process and communicate the
data while still remaining stable and efficient for the duration of the mission.
28
Guidance, Navigation, and Control
Orbital Parameter Value
Eccentricity, e 0
Altitude 425.9 km
Inclination, i 90 deg
Velocity, v 7.660 km/s
Period, P 5572.3 s
Energy, ε -29.338 (km/s)2
Team Oracle has chosen a polar orbit with a 90o inclination. This will ensure that
the satellite system will cover the entire Earth. It will also travel over the polar regions of
the Earth, which is one of the mission objectives. Each satellite’s orbit will be circular
because the satellite is set to measure accelerations and by having a circular orbit the
spacecrafts won’t experience accelerations from the orbit’s eccentricity. The spacecrafts
will orbit at an altitude of 425.9 km. This puts the spacecrafts close enough to the Earth
to get an accurate resolution for the measurements needed and also not too close the
Earth where atmospheric drag will cause the spacecraft to prematurely crash into the
Earth. The period of each orbit is 5572.3 seconds which is 92.87 minutes. The spacecrafts
will orbit the Earth 15.51 times each day. It will take about one month for the two
satellites to scan the entire Earth.
Both satellites will have almost identical orbits but they will be separated by an
average of 220 km. The maximum separation that is allowed to maintain the accuracy of
29
the scientific instruments is 270 km. The satellites will be need to facing each other at all
times in order for the K-band horn and the laser range meter to accurately measure the
changes in the satellites gap. In case the satellites become misaligned each satellite will
have cold gas thrusters to realign the two spacecraft.
Power
The power subsystem is responsible to providing power to the spacecraft so it can
perform its mission and function. The power subsystem is split into a primary power
source consisting of solar panels and a secondary power source consisting of batteries.
These two power systems together will distribute power to the instruments that need it,
such as the on-board computers and scientific instruments.
Solar panels were an easy choice for this mission. Many satellites in LEO utilize
them for their mission due to the lower cost and high efficiency at LEO’s altitude. The
solar cell that was chosen was the NeXt triple junction gallium arsenide solar cell due to
its high efficiency (29.5%) and low degradation rate. These two parameters will assist in
prolonging the spacecraft lives. A schematic of the solar cell can be seen in Figure 9
below.
30
Figure 9. Schematic of NeXt Triple Junction Solar Cell made by Spectrolab, Inc.
As with any satellite with solar panels for its primary power source, batteries
complement as the secondary power source. Due to the orbit the spacecraft will be on,
they will experience times of eclipse where the solar panels will not be able to provide
power to the spacecraft. Since our mission requires that no part of the Earth be missed
when analyzing, the scientific instruments (and spacecraft) must be powered even during
the eclipse times. The solar panels will charge the batteries during times of illumination
so the batteries will be able to distribute power during the times of eclipse. Lithium-ion
batteries were chosen for their lower mass and high efficiency, which lower cost and
prolong life subsequently. The reason lithium ion batteries are so efficient is that they can
discharge their power many times without losing their ability to store maximum power.
31
Each satellite will need about 285 W of electricity to power all the subsystems.
Table 1 shows how the power payloads breakdown for each subsystem. Based on the
power requirements the solar array was calculated to be about 6.7 m2 for each satellite.
Each satellite will house two 8V, 12 A-h Lithium Ion batteries for power storage.
Thermal Control
Using STK data for 5 years to find the angle between the Sun and the orbital
plane (β), we determined an average change of β angle with respect to the number of days
of the year.
Figure 10. Average variation of the β angle with respect to the days of the year.
This information was later used to obtain the angles of incidence of the heat fluxes in the
thermal model of each of the spacecraft. The resulting temperature range in case of a
Solar Maximum is depicted in Figure 11.
32
Figure 11. Results from the thermal mode applied to the spacecraft in Worst Case Hot (WCH) configuration.
From this model, the worst case hot (WCH) temperature resulted to be around 388K.
In our analysis, free molecular heating (2.237×10^-24 W/m^2, worst case, 429
km altitude) at moments other than the first 30 minutes in orbit was approximated to 0.
The altitude of our orbit corresponds to the layer F2 in the ionosphere. This means that
the number density of the electrons that bombard the surface (coming from UV and
extreme- UV) is between 2×10^11 and 2×1012 m^-3, a significant difference between
electron and neutral ions produce current motion outside of the spacecraft and may cause
spacecraft charging. The variation of the incidence angle over the orbital plane of the
spacecraft is between -75 and 75 degrees. This leads to a rough estimation of 15% as the
percentage of time in eclipse in the orbit. At 300 km of altitude, the spacecraft can face
atomic oxygen (AO) attack, which can lead to sputtering of the surface. This is why it is
important to prevent the satellite from decreasing its orbital radius due to drag forces. The
33
heat flux coming from IR radiation is on average 260 W/m^2 for the F2 region of the
ionosphere.
The orbit of the vehicles might be affected by free molecular drag. This force will
try to change the altitude of the spacecraft. To have an idea of how much the drag
coefficients will vary with respect to the angles of attack of a vehicle assumed to be a flat
plate, we plotted the following figure.
Figure 12. Changes in the drag coefficient for planar surfaces at different angles of attack in the environment encountered.
The satellites will be located in the F2 region of the ionosphere. For this region, the
number density of electrons is ne=2×1011 m-3 at the daytime and at nighttime it is
ne=1×1012 m-3, the critical frequencies can be calculated in the following manner:
€
fcr,F 2]day =1
2πnee
2
ε 0me
≈11MHz
€
fcr,F 2]night =1
2πnee
2
ε 0me
≈ 4.5MHz
34
This means that the downlink frequency of the spacecraft cannot approach these values.
Otherwise, the signal will “bounce” from lower levels of the ionosphere.
The Thermal Control Subsystem has set new requirements since the last updates. The
Thermal Control Subsystem must:
• Prevent the spacecraft from shifting the electromagnetic potential con its surface
drastically and for prolonged durations
• Maintain the established temperature range (-15ºC to 50ºC) adding our estimated
worst case drag coefficient (CD = 3.79) and free molecular heating (in the first
instants in orbit) to our analysis
• Prevent the spacecraft from suffering significant parasitic power drains, sputtering
and arcing in the surface
• Withstand auroral plasmas interactions over Earth’s magnetic poles, which is the
most challenging aspect for the Thermal Control Subsystem
Given the encountered space environment, the Thermal Control Team decided to
implement the following design decisions in the system:
• Use of MLI coatings in the exposed metallic surfaces of the spacecraft to provide
a surface protection to radiation and increase the overall reflectivity
• Location of heaters in the batteries and central processing unit because these
components require special minimum temperatures
• Location of a negative ground in the negative pole of the solar arrays to induce
the direction of the current to one that allows us to buy standard junctions and
components
35
• A light-weight plasma contactor will be located in the surface of each spacecraft
to counteract the effects of plasmas over the system. This component will only be
switched on in the eventuality of plasmas particles entering in contact with the
surface of any of the spacecraft. In this mode, the circuitry will enter a phase
where the circuit is not affected by the plasmas. To do this, plasmas physics
students from Princeton University will support the mission.
• The above mentioned surface finish of 8-mil quartz mirrors to prevent high
material loss through outgassing.
The following figure shows the circuit design suggestions made by the Thermal Control
Subsystem to counteract plasmas and radiation effects.
Figure 13. Thermal Control Team’s suggestion to avoid current surge in any of the components in the eventuality of plasmas contact.
36
Scientific Payload
The payload on board the satellites will be in the form of observational equipment
and data sensors that can properly determine the fluctuations in Earth’s gravitational
distribution. Due to space constraints inside the craft it was important that the team
narrowed its selection of equipment to only what was deemed absolutely necessary. The
secondary power supply has been determined and accounted for in the spacecraft payload
as well. The Lithium-ion batteries will be located in both satellites; however their smaller
size will be beneficial by leaving more room for data collection payload. The total mass
of each of the scientific payloads adds up to be 4.926 kilograms.
Communication between the two-satellites will operate with a ranging system that
utilizes newer laser technologies and the microwave ranging technology seen in GRACE
I. The introduction of the lasers is thought to provide a more accurate reading as well as
provide further redundancy in the system.
Grace II will have the same scientific instruments as the original Grace mission.
This decision was made because Grace I used instruments that developed gravity models
an order of magnitude better than all previous spacecrafts. Grace II will use a
combination of accelerometers, GPS configurations, and star cameras to accurately
determine the location of each satellite and their distance relative to each other. Based on
these values, the on-board computer can determine changes in gravitational pull and
accurately map those areas accordingly.
The decision was made to select the exact instrument models for Grace II as the
original Grace mission. A SuperSTAR accelerometer will be used to measure the non-
gravitational accelerations, which are caused by air drag and attitude control impulses.
37
The accelerometer measures the linear and angular accelerations along three axes.
Acceleration outputs are obtained from the measurements and used to keep the proof
mass of the SuperSTAR at the center of mass creating accurate data recordings.
Figure 14. SuperSTAR accelerometer
To determine the precise orbit, a Laser Retro-Reflector will be onboard to detect
near-infrared signals sent by the ground station. In addition, the Laser Retro-Reflector
will be used to calibrate the GPS. Three GPS TurboRogue Space Receivers will be used
to determine the orbit, coarse positioning, and detect the other twin spacecraft. The GPS
provides digital signal processing and measures the distance change of each satellite
relative to the GPS constellation data.
Figure 15. GPS Tracking System
38
A Star Camera Assembly will be used in addition to the GPS system to determine
precise orientation. The assembly consists of two cameras and a data processing unit that
measure the spacecraft’s attitude by detecting star constellations3. The Star Camera
Assembly measures the attitude with an accuracy of <0.3 milliradians by the autonomous
constellation detection of the onboard processing units.
Figure 16. Laser Retro-Reflector and Star Camera Assembly
In addition to the major scientific instruments onboard, the spacecrafts will have a Coarse
Earth and Sun Sensor and a Center of Mass Trim Assembly. The Coarse Sensor will be
used initially to orient the satellites and the Trim Assembly will adjust any offsets with
the center of mass to a step size of 10 micrometers. The Coarse Earth and Sun Sensor will
track the locations of the Sun and Earth relative to the spacecrafts using omni-directional
and robust tracking. Tracking will take place during the initial stages after launch and
anytime the satellites’ locations are not accurate with the GPS. The Center of Mass Trim
Assembly measures the differences between the spacecraft’s center of mass and the proof
mass of the accelerometer. The Trim Assembly makes the adjustments needed to correct
the difference. The scientific instruments work together to track the acceleration of the
spacecrafts and measure any changes between the distances of the spacecrafts.
39
Conclusion Using the system defined in this report, Oracle will work to produce a
gravitational potential map of the Earth in spherical harmonic form. This
information will prove valuable for the scientific community at large, as it will
provide an understanding of Earth water resources and humanity’s impact on these
resources. Ultimately, the Oracle mission will measure changes in the polar ice caps
as well as land water resources. Oracle will provide and understanding of shallow
and deep ocean current transport and atmosphere-‐ocean mass exchange. Through
an advanced understanding of these principles and the critical input of this mission
into oceanography, hydrology, geology, and related disciplines, Oracle hopes to
produce wide-‐reaching benefits for society and the world’s population.
Appendix Launch Vehicle:
Base
Payload
Weight (kg)
Orbit Launch Success
Performance
Weighting 4 3 1 2
Atlas V 9,800 LEO, GTO, GSO, Escape
1.00 18.5
Delta II 2,700 LEO, GTO, Escape 0.988 5.5
Delta IV 9,190 LEO, GTO, Escape 1.000 22.6
Falcon 1 670 LEO 0.400 1
40
Falcon 9 v1.0
10,450 LEO, GTO 1.000 10.4
Minotaur I 580 LEO 1.000 0.6
Minotaur IV 1735 LEO 1.000 1.7
Pegasus XL 443 LEO 0.925 0.45
Taurus 1,320 LEO 0.750 1.4
Ariane 5 G 16,000 LEO, GTO 0.945 20
PSLV 3,250 LEO, GTO 0.944 1.6
Payload
Weight (kg)
Orbit Launch Success Performance
Weighting 4 3 1 2 Total
Atlas V 0.612 1 1.00 0.818 8.51
Delta II 0.168 1 0.988 0.243 4.55
Delta IV 0.574 1 1.000 1 8.87
Falcon 1 0 1 0.400 0.044 1.93
Falcon 9 0.653 1 1.000 0.460 7.65
Minotaur I 0 1 1.000 0.026 3.08
Minotaur IV 0.108 1 1.000 0.075 3.77
Pegasus XL 0 1 0.925 0.019 2.91
Taurus 0.083 1 0.750 0.061 3.10
Ariane 5 G 1 1 0.945 0.884 7.24
PSLV 0.203 1 0.944 0.071 3.32
Command & DH Trade Study:
41
Multiplier Honeywell L-M
GVSC
L-M RAD 6000
SWRI SC-5
SWRI SC-7
SWRI SC-9
Sanders STAR-
RH
Memory 4 0.25 0.1 1 0.02 0.04 0.008 0.00025
Performance 6 1 0.1 1 0.03 0.6 1 0.5
Radiation Hardness 5 1 1 0.1 0.01 0.1 0.03 0.05
Connectivity 2 1 1 0 0.5 0.5 0 0.5
Heritage 3 0.6 0.2 1 0.4 0.2 0.2 0.2
ISA 1 1 0 1 0 0 1 1
Total 16.8 8.6 17.5 2.51 5.86 7.782 5.851
Power Trade Studies:
Weighting Si GaAs
(Single Junction)
GaAs (Triple Junction)
BOL efficiency (%) 3 14 18.5 28
Degradation Rate (%/yr) 1 3.75 2.75 0.5
Life Degradation (%) 2 85.8 89.4 98.0
Total 2.25 3.07 4.87
Weighting Ni-Cd Ni-H2 Li-Ion
Specific Energy (Whr/kg) 4 30 60 125
Energy Density(Whr/L) 3 50 100 250
Efficiency(%) 1 72 70 98
Temp Range 2 1 1 0
Total 4.29 5.83 8
Propulsion Trade Studies:
Weighting Ion Thruster Bipropellant Cold Gas Monopropellant
Thrust 6 0.5 1 0.85 0.95
Fuel 1 0.333 0.02 1 0.4
42
Specific Impulse 5 1 0.5 0.2 0.4
Mass Propellant 3 0.04525 0.0362 1 0.54
Duty Cycle 4 0 0.5 1 0.9
Power 2 0.01 0.8 0.9 1
Total 8.48875 12.2286 15.9 15.32
Thrust Specific Impulse
Cycle Life
Engine Mass
Inlet Pressure Input Power Voltage Range
Weight 7 6 5 4 3 2 1 Total
Bradford 0.0 1 0.04 0.0312 0.1299 0.0333 0 1.68
AMPAC-ISP 0.0001 0.909 0.04 0.0245 0.1299 0.0333 1 7.11
Moog 58-125
0.0001 0.844 0.04 1 0.1299 0.4167 0 10.8
Marotta 0.001 0.844 0.04 0.1049 0.8961 1 0 10.5
DASA CGT1
0.001 0.87 0.04 0.0612 0.9091 0.0333 0 7.59
Sterer 0.045 0.883 1 0.0422 0.4545 0.2 0.5 12.7
Moog 58-102
0.012 0.844 0.04 0.4893 0.9805 0.0333 0.5 9.86
Moog 58-112
0.012 0.844 0.04 0.4893 0.7987 0.0333 0.5 9.50
Moog 58-115
0.012 0.844 0.04 0.5646 0.1299 0.0333 0 7.96
Moog 58-113
0.013 0.844 0.04 0.4893 0.9805 0.0333 0.5 9.87
Moog 58-103
0.021 0.844 0.04 0.4893 0.9805 0.0333 0.5 9.92
Moog 58-673
0.167 0.844 0.02 0.0318 1 0.1667 0.5 9.46
43
Moog 50-820 0.195 0.844 0.04 0.0318 0.1299 0.1667 0 7.52
Moog 58-126
1 0.844 0.04 0.0406 1 0.0333 0.5 15.0
Table [...]. Weighting process for the polymeric and metallic component materials made by the Structures Subsystem (material properties retrieved from http://products.asminternational.org.ezaccess.libraries.psu.edu, the ASM International Database)
Category
Material
Stiffness
Stabilit
y
Ult. Tensile
Stress
(MPa)
Averg. Compressive Strength (MPa)
Ult. Yield Stress (MPa)
Flexural
Strength
(PSI)
Manufactur
er Cost per
Sheet
Density
(g/cc)
Thermal
Conductivity (W/ mK)
Thermal
Expansio
n (micom/mC)
Young's Modulus (MPa)
Corrosio
n Resistance
(MPY)
Avg. Fracture
Toughnes
s (Ku= MPa*sqrt(m))
Ease of
fabricatio
n (Weldabili
ty)
L-Elongation %
(Ductility
)
Weldalite-049-T8
(TM) High
Good
710
779.45 690 "High" 2.6 88.15 23.6 76 4.82 29.9 Good 5
Al 2090-T83
High
Good
550 562 515
$70.00 2.59 91.5 24
79.28 0.36 43.9
Good.
Superior
than 2024-
T3 3
Metals
Al-Li Hig Go 48 678 400 $45.0 2.7 90 20 82 2 75 Medi 4
44
8090 h od 0 0 um
Melting Point
degC
Specific
Gravity
Ult, Tensile
Stress
(MPa)
Water Absortion (%)
Flammabil
ity Ratin
g Rating(in)
Flexural
Strength
(GPa)
Flexural
Modulus
MPa
Maximum Usag
e Temperature (C)
Coefficient
of Thermal
Expansion (1/C)
Dry Hardness (Izod Test) kJ/m
^2
Dry Hardness
Density
(g/cc)
Tensile
Modulus
(GPa)
Fiberglas
s (Polyamid
e Alloy
) 222 1.23
120.66 2.25
W.B.(0.031
) 158.6
5,102.00 160
5.00E-05 85 83 1.8 100
Kevlar 49 (TM)
>482
0.052
3600 3.5 2.5
675-
700 70-73 177
-2.70E
-06 70 76 1.15 83
Carbon-
Carbon
Composit
e TYPE A
502-
638 1.45
15168.5 2.2 2
86.2
131.00 371 1 151.2 137 2.78 68.95
Poly-
Matrix Comp.
CFRP
>300
1.5-1.6
900-2500
0.01-0.2
70-130
850-1400
>1000
2.15E-06
90-240 56 1.6 70
Table [...]. Weighting process for the active and passive heat dissipation and transport mechanisms (mechanism properties retrieved from MINCO Thermal Solutions®. (2014). Retrieved from http://www.minco.com/Heaters and THERMACORE™ product database, retrieved from: http://www.thermacore.com/products).
Material
Stiffness
Stability
Ult. Tensile Stress (MPa)
Averg. Compressive Strength (MPa)
Ult. Yield Stress (MPa)
Manufacturer Cost
per Sheet
Density
(g/cc)
Thermal
Conductivity
(W/ mK)
Thermal
Expansion
(micom/mC)
Young's Modulus (MPa)
Corrosion
Resistance
(MPY)
Avg. Fracture Toughn
ess (Ku=
MPa*sqrt(m))
Ease of
fabrication
(Weldability)
L-Elongatio
n % (Ductility) Total
Metals
Weldalite-
049-T8 (TM
) 1 1 1 1 1 0.2
0.8666666667 1
0.8666666667
0.6666666667 1
0.39866666
67
0.9333333333 1
86.6573333
3
45
Al 2090-
T83
0.8666666667
0.8666666667
0.7746666667
0.7213333333
0.746 0.56 1
0.7333333333
0.6666666667
0.8666666667
0.0746666666
7
0.58533333
33 1 0.6
73.5033333
3
Al-Li
8090
0.9333333333
0.9333333333
0.6760666667
0.8703333333
0.5797066667 1.00
0.7333333333
0.8666666667 1 1
0.4149333333 1
0.6666666667
0.8
85.5574
Multiplier 2 11 3 4 5 12 14 6 13 1 10 7 9 8
Materials
Melting
Point
degC
Specific Grav
ity
Ult, Tensi
le Stres
s (MPa)
Flexural Strength (GPa
)
Flexural Modulus MPa
Maximum
Usage
Temperature (C)
Coefficient of Thermal Expansion
(1/C)
Dry Hardness (Izod Test
) kJ/m
^2
Dry Hardness
Density
(g/cc)
Axial Young's
Modulus (GPa
) Total
s
Fiberglass (Polya
mide
Alloy)
0.3461538462
0.3846153846
0.00792307692
3
0.6153846154
1.0000000000
0
0.5384615385
0.5384615385
0.5384615385
0.9230769231
0.6153846154 1
52.2545384
6
Kevlar 49
(TM)
0.7553846154 1
0.2376923077 1
0.3846153846
2
0.6153846154
0.9230769231
0.3846153846
0.7692307692 1
0.7692307692
60.4546153
8
Carbon
-Carbon Composite
TYPE A 1
0.6923076923 1
0.5384615385
0.6923076923
1
0.6923076923
0.3076923077
0.7692307692 1
0.3076923077
0.4615384615
53.3846153
8
CFRP
0.4707692308
0.6153846154
0.1648461538
0.4615384615
0.7692307692
3 1 1 1
0.3846153846
0.5384615385
0.6153846154
57.5899230
8
Poly-Matrix Comp.
Multiplier 6 2 3 4 1 5 7 8 9 11 10
46
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