104
1 Formation Flying Rachel Winters Matt Whitten Kyle Tholen Matt Mueller Shelby Sullivan Eric Weber Shunsuke Hirayama Tsutomu Hasegawa Aziatun Burhan Masao Shimada Tomo Sugano

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Formation Flying. Rachel Winters Matt Whitten Kyle Tholen Matt Mueller Shelby Sullivan Eric Weber Shunsuke Hirayama Tsutomu Hasegawa Aziatun Burhan Masao Shimada Tomo Sugano. Motivation. Can enable baseline to form large instruments in space Escort Flights - PowerPoint PPT Presentation

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Page 1: Formation Flying

1

Formation Flying

Rachel WintersMatt WhittenKyle TholenMatt Mueller

Shelby SullivanEric Weber

Shunsuke HirayamaTsutomu Hasegawa

Aziatun BurhanMasao ShimadaTomo Sugano

Page 2: Formation Flying

2

Motivation

• Can enable baseline to form large instruments in space

• Escort Flights– Provide detection/protection from threats– Provide visual inspection for damage

Page 3: Formation Flying

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: Formation Flying

4

Systems Integration & Management

Rachel Winters, Matt Whitten• Expendable vs Recoverable spacecraft

(90%)• Recovery method designed (80%)• Determine shuttle-interface

requirements (100%)

Page 5: Formation Flying

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%)

Page 6: Formation Flying

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%)

Page 7: Formation Flying

7

Attitude Determination & Control

Shunsuke Hirayama, Tsutomu Hasegawa• Determine method of attitude control

(80%)• Single vs Multiple cameras (90%)

Page 8: Formation Flying

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%)

Page 9: Formation Flying

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

Page 10: Formation Flying

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

Page 11: Formation Flying

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.

Page 12: Formation Flying

12

Systems Integration and Management

Rachel WintersMatt Whiten

Page 13: Formation Flying

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

Page 14: Formation Flying

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

Page 15: Formation Flying

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.

Page 16: Formation Flying

16

Satellite-Shuttle Interactions

• Capture feasibility case study– MIR Space capsule– SPARTON satellite– SFU Satellite

• Automatic movement near to shuttle– Mini AERCam– STS-87

Page 17: Formation Flying

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

Page 18: Formation Flying

18

Delta V estimation

• Delta V for orbit transfers estimated with Clohessy Wiltshire equations:

}vv(t)]{[}vr(t)]{[)]([

}rv(t)]{[}rr(t)]{[)]([

vorotv

vorotr

Page 19: Formation Flying

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

Page 20: Formation Flying

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.

Page 21: Formation Flying

21

Orbit Determination

• Need two orbits to view shuttle from all angles

• Orbits achieved through small changes in Inclination and Eccentricity

Page 22: Formation Flying

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

Page 23: Formation Flying

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

Page 24: Formation Flying

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

Page 25: Formation Flying

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

Page 26: Formation Flying

26

STK Orbit Simulation

Page 27: Formation Flying

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

Page 29: Formation Flying

29

Structure and Configuration

Page 30: Formation Flying

30

Structure and Design

Gyro

Magnometer

CPU

Transceiver

Thruster

Page 31: Formation Flying

31

Structure and Design

• Future Work– Reconfigure satellite structure to better

accomplish design goals– Model remaining hardware– Place selected hardware to accomplish

design goals

Page 32: Formation Flying

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

Page 33: Formation Flying

33

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

Page 34: Formation Flying

34

Field of View

Page 35: Formation Flying

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)

Page 36: Formation Flying

36

Sample PicturesWith Pixels/Meter

2350

9500

600

390 200

~ 360m from shuttle

~ 700m from shuttle

Page 37: Formation Flying

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°

Page 38: Formation Flying

38

360 Meters from Shuttle

Page 39: Formation Flying

39

Attitude Determination & Control

Shunsuke HirayamaTsutomu Hasegawa

Page 40: Formation Flying

40

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?

Page 41: Formation Flying

41

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)

Page 42: Formation Flying

42

Attitude determination

Front View Side Viewx y

z

Page 43: Formation Flying

43

Aerodynamic torque

Altitude 326-346km

0.3DC

for worst case

311300 /10418.2 mkg

S = 0.4243 m2

Page 44: Formation Flying

44

Gravity-Gradient Torque

n3 = μ = 398600 km3/s2

R3 3263 km3

Page 45: Formation Flying

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

Page 46: Formation Flying

46

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

Page 47: Formation Flying

47

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.

Page 48: Formation Flying

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

Page 49: Formation Flying

49

Future work

• calculate a disturbance from magnetic torque.

• work on matlab with all disturbances.

Page 50: Formation Flying

50

Communications

Tomo Sugano

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51

Tasks done so far:

•Communication/CPU selection

•In-flight Delta V estimation of the mission

•Atmospheric Drag Analysis

•Orbital Decay Life

Page 52: Formation Flying

52

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

Page 53: Formation Flying

53

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.

Page 54: Formation Flying

54

Arcom VIPER 400 MHz embedded controller

Page 55: Formation Flying

55

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

Page 56: Formation Flying

56

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

Page 57: Formation Flying

57

Scheme of FCS/COMM Integration

Page 58: Formation Flying

58

Detailed Scheme of integration

Page 59: Formation Flying

59

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)

Page 60: Formation Flying

60

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

Page 61: Formation Flying

61

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

Page 62: Formation Flying

62

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

Page 63: Formation Flying

63

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

Page 64: Formation Flying

64

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

Page 65: Formation Flying

65

Orbit Lifetime Equation

• Circular Orbit Lifetime Equation (Approximation)

a0 = initial altitude

S = projected area of the space probe m = space probe mass

Page 66: Formation Flying

66

Exponential Atmospheric Model

• Scale height, H, obtained from tabulated data

Page 67: Formation Flying

67

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)

Page 68: Formation Flying

68

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.

Page 69: Formation Flying

69

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(?)

Page 70: Formation Flying

70

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

Page 71: Formation Flying

71

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

Page 72: Formation Flying

72

Computations (cont’d)

• Orbiter speed (assuming circular orbit)

• Definition of Delta-V (or specific impulse)

• Mass expenditure of propellant (i.e. GN2 cold gas)

Page 73: Formation Flying

73

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

Page 74: Formation Flying

74

Thermal Control Subsystem

Masao Shimada

Page 75: Formation Flying

75

• 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

Page 76: Formation Flying

76

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

Page 77: Formation Flying

77

o

Orbit Model

984.51ISS

Sunlight

Shadow

s

Page 78: Formation Flying

78

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

Page 79: Formation Flying

79

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

Page 80: Formation Flying

80

Satellite Model for Node Analysis

Assume no width for each surface Surface 2 always look downward.

Page 81: Formation Flying

81

Ex: Direct Solar Flux (Worst-case Hot)

rads 04.1

0

0

0

876521 ,,,,,

4

3

0.616

0.788

2 3

ii SS

Page 82: Formation Flying

82

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

Page 83: Formation Flying

83

Surface characteristics

•Inside of the satellite is painted with L-300 (Black)•Conductivity between surfaces : K=0.06

Page 84: Formation Flying

84

Simulink model (Node Analysis)

Page 85: Formation Flying

85

Simulation (Worst-case COLD)

Temperatures [K]

Page 86: Formation Flying

86

Simulation (Worst-case HOT)

Temperatures [K]

Page 87: Formation Flying

87

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.

Page 88: Formation Flying

88

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

Page 89: Formation Flying

89

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

Page 90: Formation Flying

90

Trade study

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91

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

Page 92: Formation Flying

92

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

Page 93: Formation Flying

93

• 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 )

Page 94: Formation Flying

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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.

Page 95: Formation Flying

95

Primary Battery

Page 96: Formation Flying

96

Solar Cells

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Rechargeable Battery

Page 98: Formation Flying

98

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

Page 99: Formation Flying

99

• 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% )

Page 100: Formation Flying

100

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

Page 101: Formation Flying

101

Power Subsystem

• General Layout

Power source

Power distribution

Dc-Dc converterLoad

Energy

storage

Payload

Comm.

ADCS

Propulsion

Thermal

Page 102: Formation Flying

102

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