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![Page 1: 1 Formation Flying Rachel Winters Matt Whitten Kyle Tholen Matt Mueller Shelby Sullivan Eric Weber Shunsuke Hirayama Tsutomu Hasegawa Aziatun Burhan Masao.](https://reader031.fdocuments.net/reader031/viewer/2022013012/56649dab5503460f94a99e26/html5/thumbnails/1.jpg)
1
Formation Flying
Rachel WintersMatt WhittenKyle TholenMatt Mueller
Shelby SullivanEric Weber
Shunsuke HirayamaTsutomu Hasegawa
Aziatun BurhanMasao ShimadaTomo Sugano
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2
Motivation
• Can enable baseline to form large instruments in space
• Escort Flights– Provide detection/protection from threats– Provide visual inspection for damage
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3
Design
• A satellite that will fly escort to the space shuttle
• Satellite provides visual inspection of shuttle exterior for 24 hour period of time
• Satellite will be transported into space on shuttle
• Satellite must meet University Nanosat requirements
![Page 4: 1 Formation Flying Rachel Winters Matt Whitten Kyle Tholen Matt Mueller Shelby Sullivan Eric Weber Shunsuke Hirayama Tsutomu Hasegawa Aziatun Burhan Masao.](https://reader031.fdocuments.net/reader031/viewer/2022013012/56649dab5503460f94a99e26/html5/thumbnails/4.jpg)
4
Systems Integration & Management
Rachel Winters, Matt Whitten• Expendable vs Recoverable spacecraft
(90%)• Recovery method designed (80%)• Determine shuttle-interface
requirements (100%)
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5
Relative Orbit Control & Navigation
Kyle Tholen, Matt Mueller• Determine relative orbit to meet
mission requirements (90%)• Determine major disturbances from
orbit and counteract them (100%)• Single vs Multiple spacecraft (90%)
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6
Configuration & Structural Design
Shelby Sullivan, Eric Weber• Find general hardware (cameras,
thrusters, etc.) (100%)• Design structure (material, shape)
(90%, pending necessary changes)• Solidwork components (60%)
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7
Attitude Determination & Control
Shunsuke Hirayama, Tsutomu Hasegawa• Determine method of attitude control
(80%)• Single vs Multiple cameras (90%)
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8
Power, Thermal & Communications
Aziatun Burhan, Masao Shimada,Tomo Sugano
• Determine power needed by satellite (70%)
• Battery only vs Solar Cell + Battery (70%)• Define thermal environment (outside and
inside sources) (80%)• Determine insulation needed (60%)• Determine transmission method (100%)
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9
Trade Studies
• Expendable vs Recoverable Satellite– method of picture storage– viable method of recovery– reasonable amounts of extra fuel needed
• Single vs Multiple Satellite(s)– amount of extra fuel needed for plane
transfers– ability to “see” entire shuttle with only 1
satellite
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10
• Solar cells + Battery vs Battery only– Amount of power solar cells can provide in 24
hr period– Amount of power needed by satellite
components– Size of battery needed to compliment solar
cells vs size of battery needed with no recharge
• Single vs Multiple camera(s)– Ability to control attitude– Camera size
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11
Other Design Aspects
• Structure: Rectangular satellite with aluminum supports, center of mass designed to be at the center of the prism.
• Navigation: Will be using DGPS for location and velocity information, magnometer and gyro for attitude determination.
• Transmission: Decided to store images on memory stick instead of using live transfer.
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12
Systems Integration and Management
Rachel WintersMatt Whiten
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13
SIM
• Role: Work with all groups to balance workload.
• Tasks: – Research lightband technology– Perform trade study on attitude sensors– Research ARVD– Research, calculate and design recovery
method.
Matthew Whitten
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14
SIM
• Attitude Sensors– Distance requires the camera to have the
most accurate attitude control– Small satellite requires inexpensive and
small equipment
• Recovery Method– Robotic arm’s length must be able to reach
the recovery orbit around the shuttle– Design and format end effect to capture
satellite
Matthew Whitten
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15
Special Requirements
• Transmission restrictions– NASA operates in the S-band of frequencies,
from 1700 - 2300 MHz, the space shuttle is generally contacted at 2106.4 and 2041.9 MHz, and the Orbiter also uses the Ku-band, from 15250 - 17250 MHz.
• Vibration requirements– Vibration tests with NASA are usually done
from 20 - 2000 Hz.
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16
Satellite-Shuttle Interactions
• Capture feasibility case study– MIR Space capsule– SPARTON satellite– SFU Satellite
• Automatic movement near to shuttle– Mini AERCam– STS-87
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17
Orbital Navigation and Control
Group Members:Kyle Tholen
•Orbit Determination •Delta V Estimation
–GPS Navigation
Matt Mueller•Effects of Earth’s Oblatness•Propulsion Methods•Orbit Modeling in STK
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18
Delta V estimation
• Delta V for orbit transfers estimated with Clohessy Wiltshire equations:
}vv(t)]{[}vr(t)]{[)]([
}rv(t)]{[}rr(t)]{[)]([
vorotv
vorotr
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19
GPS Navigation
• GPS can be used to determine position in orbit
• Two signals are transmitted from GPS satellites– Precise Position Service (PPS)
• Very accurate• Currently restricted to military applications
– Standard Position Service (SPS)• Available for anyone to use• Not as accurate as PPS
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20
GPS Navigation Continued
• Use Differential GPS (DGPS) for a much more accurate position– Need a known fixed reference position with
GPS capabilities– Space Shuttle are GPS certified and position
is known very accurately with ground tracking
• DGPS can potentially be accurate to the centimeter.
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21
Orbit Determination
• Need two orbits to view shuttle from all angles
• Orbits achieved through small changes in Inclination and Eccentricity
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22
Effect Of Earth’s Oblatness
• Causes secular drift in right ascension, argument of perigee and mean anomaly
iae
RJdot cos
)1(2
32/722
2
2sin
2
5
)1(2
3 22/722
2 iae
RJdot
22
222
22
12sin31
/
4
3ei
ea
RJa
a
Mdot
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23
Earth’s Oblatness Continued
• Effect on shuttle and satellite nearly the same over 24 hr period
degdegdeg
• These values will give the change in the relative distance to the shuttle, estimation of deltaV needed to correct orbit.
00015.000463.00038.M
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24
Propulsion Methods
Requirements– Small amount of thrust– Capable of being used numerous times– Small size, light weight– Low price
Possible candidates– Small mono-propellant hydrazine thrusters– Cold gas thrusters– Due to simplicity, ease of handling and
price, cold gas thrusters were chosen as method of propulsion
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25
Orbit Modeling in STK
• Visualization of relative orbit proved difficult without simulation
• Created scale simulation of shuttle orbit as well as satellite orbit
• Useful to visualize relative orbit about shuttle and aid in initial selection of orbit parameters– Use of MATLAB distance function
determined final orbit parameters– Simulation proved orbit provided 100%
visible coverage of shuttle
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26
STK Orbit Simulation
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27
Configuration & Structural Design
Shelby SullivanEric Weber
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28
Structure and Configuration
• Satellite Structure– Cube (60x60x50 cm)– Aluminum
• Low cost and availability• Success on many other satellites• Adequate properties for mission
• Configuration– Keep the moments of inertia near center of cube– Allow space for large camera to see through one face– Allow for proper thermal control
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29
Structure and Configuration
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30
Structure and Design
Gyro
Magnometer
CPU
Transceiver
Thruster
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31
Structure and Design
• Future Work– Reconfigure satellite structure to better
accomplish design goals– Model remaining hardware– Place selected hardware to accomplish
design goals
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32
Camera - MegaPlus II EP1600
16 Megapixel4872 x 3248
Three sensor grades for “demanding applications”Selectable 8, 10, or 12 bits/pixel“Temperature Resistant” construction
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Lens - Nikon Super Telephoto 1000mm
• Angle of view – 2 x 1.4 degrees• Length – 24 cm• Mass – 2 kg• Fixed focal length
– Little to no moving parts– Higher vibration resistance– Higher temperature resistance
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34
Field of View
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35
Distance vs Pixels per meter1000mm focal
0
200
400
600
800
1000
1200
1400
1600
0 200 400 600 800 1000 1200
Distance (m)
Pix
els
per
met
er (P
/m)
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36
Sample PicturesWith Pixels/Meter
2350
9500
600
390 200
~ 360m from shuttle
~ 700m from shuttle
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37
~360m From Shuttle
• ~Cross-sectional are of shuttle – 400 m^2
• Field of View area– 105 m^2
• ~25% of shuttle captured per photo• Accuracy required for view of shuttle
– X angle ~ 2.6°– Y angle ~ 0.86°
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360 Meters from Shuttle
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Attitude Determination & Control
Shunsuke HirayamaTsutomu Hasegawa
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2 Bias+Roll
Bias+Roll/Yaw
Zero momentum
Type
Gravity inclination
Spin
Dual Spin
freeHeat generation
Bias Momentum
Controled bias
momentum
costAttitude Change
Accuracy
Reaction
Wheel
0
1
3
Low
High
None
Components
Nutation
Dumper
Mass WheelReaction
Wheel Momentum
Large
Difficult
Easy
SmallLow
High
Can't change
Other
Weak from disturbance For small
communication Sat.
Strong from disturbance For Large Sat. LEO
is OK
Weak from disturbance. For communication Sat. Don't use LEO
Why Zero momentum?
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2
2
2
2
2
2
32122
21 )(
h
h
b
b
a
a
dddJ
)(12
]1212
[ 2233
baabh
habba
Moment of inertia of a*b*h cube sat.
h
ba
From Nihon Univ. Text book
][3)6.06.0(12
1][50 222 mkgkg
Once we get angular acceleration, we can get the Moment.
Tsutomu and Shunsuke
Where, is body frame exm
based momentmex = Jώ + ω x (Jω)
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42
Attitude determination
Front View Side Viewx y
z
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43
Aerodynamic torque
Altitude 326-346km
0.3DC
for worst case
311300 /10418.2 mkg
S = 0.4243 m2
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44
Gravity-Gradient Torque
n3 = μ = 398600 km3/s2
R3 3263 km3
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45
Solar Radiation Pressure Torque
24243.0 mSA
Our surface material is Aluminum
0.02 K 0.04 (surface reflectivity)
Is = 1358 w/m2 at 1 AU
smc /109979.2 8
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Choosing reaction wheel• Using Matlab we calculated required
torque to change attitude with disturbances. The result is below:
Rise Time: 14.178531Settling Time: 1.322471Overshoot: 32.247096 %Max Torque: 0.024617
Max torque is 24.6mNm so that we use reaction wheel produced by Sunspace whose max torque is 50mNm.
There is error so that we should work on matlab again.
For Y axis
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Problem about simulation
• Disturbance torque is:
Required torques is:
mNmTYreq 4.26
YY reqdis TT We should figure out what is wrong and fix it.
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48
Requirement for reaction wheel
The rotation speed of satellite should be:
360º/90min = 0.06667deg/s = 0.0698 rad/min - 1.163x10-3 rad/s
It takes 90 min to go around the orbit.
360º/90min
We use 0.1 rad/s as a rotation speed in matlab
sradsrad /10163.1/1.0 3
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49
Future work
• calculate a disturbance from magnetic torque.
• work on matlab with all disturbances.
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50
Communications
Tomo Sugano
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Tasks done so far:
•Communication/CPU selection
•In-flight Delta V estimation of the mission
•Atmospheric Drag Analysis
•Orbital Decay Life
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FCS and COMM
• FCS – Flight Control System• COMM – Communications (camera is
assumed to be part of COMM)• Satellite needs to handle both FCS and
COMM systems• Use of COTS (Consumer Off-the-Shelf)
computer(s) aimed• COMM utilizes a low-cost COTS
transceiver radio
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CPU selection for the Nanosat
• Arcom VIPER 400 MHz CPU recommended• VIPER is suitable because of its
- Light weight, 96 grams- Operable temperature range, -40 C to + 85 C- Windows Embedded feature, easy to program- Computation speed, 400 MHz- Memory capacity, up to 64MB of SDRAM- Embedded audio I/O, necessary for COMM with voice radio
• Redundancy can be implemented.
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Arcom VIPER 400 MHz embedded controller
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Radio selection for the Nanosat
• Kenwood Free Talk XL 2W transceiver recommended
• Kenwood Free Talk XL is suitable because of its- COTS nature, low cost- 2W of transmission power, more than enough for non-obstructed space communication, but higher wattage than FRS 500 mW radio- Ability to use both GMRS and FRS frequencies- FRS frequencies recommended because by international treaty FRS (Family walkie talkie) is restricted to 500 mW- 500 mW is too weak to penetrate into space- MilSpec cetified
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Kenwood Free Talk XL 2W FRS/GMRS Transceiver
• 15 UHF channels (7 FRS and 8 GMRS)• 2W output for both categories• DC 7.2 V (600mAh)• Circuit board weighs only 60 grams • Speaker/Microphone/Encapsulation
Removed
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Scheme of FCS/COMM Integration
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Detailed Scheme of integration
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Presence of Atmospheric Drag in LEO orbit
• Atmospheric density is largest at perigee• Largest drag is experienced at perigee• Atmospheric drag shall be considered if orbit
perigee height is <1000 km• Atmospheric drag acceleration (D):
• 1/(ACD/m) is the ballistic coefficient, a measure of resistance to fluid
A (projected area normal to flight path) m (mass of spacecraft) f (latitude correction coefficient)
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Effect of Atmospheric Drag to Orbit Profile
• Atmospheric drag tends to circularise the probe’s orbit
• Drag effect greatest at perigee• Apogee height consequently reduced• Overall altitude is lost unless orbit correction is
done• Determinant of satellite decay time
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Drag Coefficient of STS and other LEO probes
• STS Orbiter (aka the Space Shuttle)
• STS has a CD of 2.0 at typical mission altitudes in LEO
• Above 200 km of orbit altitude, use 2.2 < CD < 3.0
• Cylindrical probes have larger CD than those of spherical probes
• Exact CD is hard to predict as LEO environment is not fully understood
• Currently best determined by actual flight test
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Consideration of Drag in Formation Flying
• FF mission is required to last at least 24 hours
• STS orbiter (primary) typically performs a trim burn once a day
• Trim burns correct orbit altitude and ascending node
• Drag differentials present between primary and satellite(s)
• Possible consideration of LEO drag in our mission
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Orbital Decay
• Perturbation in LEO is mainly due to atmospheric drag
• Orbital decay of space probes (e.g. Space Shuttle, ISS, satellites)
• Altitude correction “trim burns” necessary to keep probes in orbit
• Orbit will decay in the absence of trim burns
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Orbit Lifetime Estimation
• Estimation of the orbit lifetime of our satellite after mission
• Consider atmospheric drag effect only• Mission orbit is assumed virtually
circular for simplicity
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Orbit Lifetime Equation
• Circular Orbit Lifetime Equation (Approximation)
a0 = initial altitude
S = projected area of the space probe m = space probe mass
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Exponential Atmospheric Model
• Scale height, H, obtained from tabulated data
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Assumptions set forth for our lifetime computation
• Assumptions: (Made for worst case or shortest decay)m = 50 kg (maximum); S = 0.385m2 (spherical correction of max volume)CD = 3.0 (upper bound value in LEO probes)
a0 = 6400 + 300 km (typical altitude for STS or ISS)
Δ = 150 – 300 = - 150 km (typical re-entry altitude, note the minus sign)f = 1 (ignore latitude effect; not significant (<10%))ρ0 = 2.418x10-11 kg/m3 (Table, 300 km base altitude)
• Unavoidable uncertainty Scale height, H- Not constant between orbit and re-entry altitude- Take H = 30 km, so β = 1 / (30 km)
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Computation Result
• Based on the assumptions we made- T = tau_0 * 189.565- T = (approx. 1.5 hr of initial orbit period)*(190) = 12 days
• LEO Nanosat at 300 km of altitude will take 12 days to decay.
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Conclusion
• Our Nanosat does not decease for 12 days• Retroburn delta-V input to decelerate the
Nanosat for faster decay will be costly without a compelling space debris concern(?)
• Unless allowed to dispose of the Nanosat in space, retrieval is rather recommended(?)
• Retrieval may be attained fairly easily by using robot arm of STS perhaps equipped with capture net(?)
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Drag Differential Compensation
• Different ballistic coefficients between the orbiter and the Nonosat
• Consequent difference in drag forces exerted during mission
• Ballistic Coeff. of STS >> Ballistic Coeff. of Nanosat
• Nanosat must expend Delta-V to keep up with STS orbiter
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Computations• Atmospheric drag acceleration (Da):
• Drag (acceleration) difference between the two spacecraft:
STS: S = 64.1 m2, CD = 2.0, m = 104,000 kg (orbiter average)
sat: S = 0.385 m2 (nominal), CD = 3.0 (worst case), m = 50 kg
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Computations (cont’d)
• Orbiter speed (assuming circular orbit)
• Definition of Delta-V (or specific impulse)
• Mass expenditure of propellant (i.e. GN2 cold gas)
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Results• Using Isp = 65 sec; assume 50 kg for satellite
weight
• Conclusions- At the typical 300 km LEO, Delta-V for 1 day mission is 1.36 m/s- Satellite will need at least 107 grams of GN2 to compensate drag- Besides this Delta-V requirement, we have orbit transfer Delta-V (currently estimated at 1.17 m/s) and ADCS Delta-V.
Altitude [km] Density [kg/m3] Dsat-Dsts [m/s2] Thrust Req [N] m_dot [kg/s] mass in 24hr [kg] Delta V for 24hr [m/s]300 2.4180E-11 1.5780E-05 7.8901E-04 1.2386E-06 0.1070 1.363350 9.1580E-12 5.9322E-06 2.9661E-04 4.6564E-07 0.0402 0.513400 3.7250E-12 2.3951E-06 1.1976E-04 1.8800E-07 0.0162 0.207
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Thermal Control Subsystem
Masao Shimada
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• Qs : Direct radiation from the Sun• Qe : Radiation from the Earth• Qa : Solar radiation reflected back by the earth (Albedo)• Qi : Heat generation• Qps : Radiation to Space• Qpe : Radiation to the Earth
Qa
Qe
Qs
Qi
Qpe
Qps
pepsiaesp QQQQQQdt
dTCm
Space Thermal Environment
Earth pic: http://palimpsest.typepad.com/frogsandravens/pictures/earth.jpg
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Orbit Model
kmPHAH
H ISSISS 3352
min22.912
3)(
HR
MGT
s
Approximated ISS Circular orbit:
Period (T) :
radHR
HRHs 04.157.59
cos)(
2arccos
2
Atitude (H) :
Shadow time (Ts) :
Shadow angle ( ) :
min19.3033.0360
2
TTT ss
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o
Orbit Model
984.51ISS
Sunlight
Shadow
s
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1. Steady-State Approximation
peearthpsspaceislnpeenn
pepsiaesp
FTTAFTTAQSFaAFIASA
QQQQQQdt
dTCm
)()(0 4444
Assumptions:1) Steady State: dT/dt=02) Spherical satellite with thermal surface area A= 2.16m^2 so An=0.54m^2 3) Surface characteristic: 4) Heat generration: 50W5) View Factors:
6) Direct Solar flux: 0 (Cold), 1399w/m^s (Hot)
)(5.0),(0
5.01
5.0)/(112
1 22
HotColdF
FF
HRRF
sl
peps
pe
8.0,6.0 s
Results:1) Worst-case HOT: Tmax= 316.0 K2) Worst-case COLD: Tmin = 219.3 K
Tmax-Tmin=96.7 K
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2. Node Analysis
isliipeeiiiiii
pi QSFAaFIASAdt
dTCm
ii
)()( 441
11jiiji
n
jjiji
n
jij TTFATTK
ii peearthiiipsspaceiii FTTAFTTA )()( 4444
QaQs Qe
QpeQps
Thermal Equilibrium Equation
Conduction between Node i and Node j Radiation between Node i and Node j
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Satellite Model for Node Analysis
Assume no width for each surface Surface 2 always look downward.
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Ex: Direct Solar Flux (Worst-case Hot)
rads 04.1
0
0
0
876521 ,,,,,
4
3
0.616
0.788
2 3
ii SS
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Worst Cases
s
984.51ISS
984.6
Worst-case HotWorst-case Cold
s
00
Earth pic: http://www.bc.edu/schools/cas/geo/meta-elements/jpg/new_earth.gif
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Surface characteristics
•Inside of the satellite is painted with L-300 (Black)•Conductivity between surfaces : K=0.06
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Simulink model (Node Analysis)
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Simulation (Worst-case COLD)
Temperatures [K]
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Simulation (Worst-case HOT)
Temperatures [K]
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Results (Node Analysis)
minmax TTT
][: CUnit
•High temperature differences on surface 4 and 5•Use MLI to make thermal disturbances from outside smaller.•Need to consider thermal control methods to make temperature higher.
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Future works
• Thermal Control by using Thermal Control Elements so that Design Temperature range fits Permissible temperature range of components.
• More-nodes analysis for accurate simulation
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EPS Design
• Trade study of PV-battery vs battery as power source•Preliminary analysis (solar array sizing & battery sizing)•Power Load Profile•Overview of other power susbsystems design - power distribution - power regulation
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Trade study
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Trade StudyMission Constraint & Requirement:• Length of mission: 1 day• Mass <= 30 kg• Size: 60 cmx60 cmx50 cm
ATITUDE CONTROL• Conformal solar array - required spinner to radiate excess heat . Cells not always
oriented to the sun , thus reducing power output - for 3 axis stabilized satellite that does not employ active
tracking, array’s reduction in output power per total surface area would be approximately 4. Not all surfaces are in the sun.
• Primary battery - does not affect the choice of attitude control
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Trade studyOPERATING ENVIRONMENT• LEO orbit: Worst cold ~-80°C, worst hot ~ 100°C• Solar flux variation • Radiation
PERFORMANCE• Conformal solar array - less power output due to cosine loss - single cell efficiencies : 14.8 % (Si), 18.5% (GaAs) - assembled solar array is less efficient than single cell due to inherent degradation, Id ( design efficiencies, temperature, shadowing). Nominal value of Id at 0.77 - life degradation -> ≈1 for short mission (days) - peak power point depends on the array’s operating temperature - required energy storage -> provide power during eclipse
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• Battery - cell voltage decays with Ah discharge - small range of operating temperature -> require thermal
control
THERMAL CONTROL• Both require thermal control, but could be complex for solar
array
COST• Solar array : $800-3000/W . GaAs costs 3 times more than Si $5-$13 per cell• Rechargeable battery: $8/cell (NiMh) - $30 (Li Ion)• Primary Battery: Lithium type (~ N/A )
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RISK & SAFETY• Solar array - shadowing of one cell results in the loss of entire
string. Low risk (with bypass diodes) - minimal safety analysis reporting• Primary Battery - limited space qualified battery , safety concerns CONCLUSION• Choice of power source depends on power load profile• Analysis need to be done to make sure power source
meet the mass & area constraint.
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Primary Battery
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Solar Cells
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Rechargeable Battery
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ANALYSIS• Preliminary Solar Array Sizing (according to Space Mission Analysis &
Design textbook)
Assumption: - Only 2 surfaces will be used to mount solar array. Therefore, optimum available area is 0.72m sq. Maximum number of
cells =900 - average power: 50 W - lifetime: 1 day - PPT regulation scheme: Xe=0.6, Xd=0.8
Input: Orbital parameter (ISS orbit) h=300km, inclination = 52 degrees, assume circular orbit => eclipse duration ~36 min, orbital period ~ 91 min
Equation: Psa = [( PeTe/Xe) + (PdTd/Xd) P BOL = Po* Id* cosθ P EOL = P BOL * Ld P BOL because Ld 1 for 1 day mission Asa = Psa / P EOL
Ma = 0.1 P (with specific 100W/kg) Solar array area , Asa = 0.86m^2Mass of solar array = 1.18 kg
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• Preliminary array sizing ( according to AEM4332 textbook, pg 495 )
- A array = 1.68 m^2
- N cell = 2100 cells
• Energy storage sizing ( textbook pg. 485) Mass of battery = 1.55 kg Number of NiCd cells = 22 cells
Total mass for solar array + battery = 1.18 kg + 1.55 kg = 2.73 kg
• Primary battery sizing (lithium sulfur dioxide) Number of cells= 10
Total mass of battery= 6.65 kg (22% )
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Power Load Profile
0.2 *P + 0.05P wiring & cable
0.07*P 2 PPT, Power (dissipation)
Thermal
35-4528 dc1 GPSOrbit determination
9 each 3 to 12 Cold gas ThrusterPropulsion
5 (estimate) Onboard computer
4.327.2 dc Radio
27.2dc Receiver power output
27.2dc Transmit power outputCommunications
2 W (max) per each11 - 16 dc3 Reaction wheel
312 dc1 Earth sensor
3 / 0.00524-32 dc / 5 -15 dc1 Sun sensor
8.528 dc1 Star sensorAttitude control
1512 dc1 CameraPayload
Power Consumption (W)Operating Voltage: V (1 unit)
Quantity
WeightComponentSUBSYSTEM
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Power Subsystem
• General Layout
Power source
Power distribution
Dc-Dc converterLoad
Energy
storage
Payload
Comm.
ADCS
Propulsion
Thermal
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Power Distribution & Regulation
• Main tasks: - power the satellite operation directly - control bus voltage on EPS - control power generated by solar arrays** - charge the secondary battery**• Centralize control• 28 Vdc bus voltage (regulated)• **PPT : extract exact power a satellite require up to
array’s peak power • Distribution subsystem consist of cabling, fault
protection, switching gear, converters (dc-dc)• **Battery charging system: Parallel / individual charging
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Future work• Power duty cycle (application profile) - continuous / noncontinuous operation• Detail solar array design
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Thanks,
Derek Surka,Joe Mueller