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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.

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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

⎛

⎝ ⎜

⎞

⎠ ⎟

8

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


Dish Size m 15 At Weilheim

Frequency GHz 2.12 S-‐Band


Pointing Error deg 0.05 Estimated



Xmtr Gain dB 47.7934

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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



Peak Xmtr Power W 801.030

Crosslink Budget


Horn Circumference m 0.31416 From Sat. 1 to Sat. 2

Frequency GHz 32 Ka-‐Band

Wavelength m 0.00938





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


Sat. 2 Rcvr G dB/K 22.3645





Horn Circumference m 0.31416 From Sat. 2 t oSat. 1

Frequency GHz 24 Ka-‐Band

Wavelength m 0.01250





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


Sat. 1 Rcvr G dB/K 24.86294




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

23

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.

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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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