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Electrical Power System

ByAziatun Burhan

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Overview

Design goal requirements throughout mission operation:

• Energy source generates enough electrical power • Energy storage stores electrical power• Power distribution distributes electrical power • Power regulation controls electrical power

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Power BudgetSubsystem Power consumption (W)

Total Power (inc. 20% safety margin)

Continuous operation  

23.7     

ADCS 4

Orbit Control 3.74

OBC 2

Power 5

Communication 3.05

Camera (idle) 2

   

Non-continuous operation  

66    

Thermal control ** 10

Propulsion ** 30

Camera ** 15

   

Total   89.7

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Mission Power Profile

Power Profile

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30

40

50

60

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80

0 1000 2000 3000 4000 5000 6000

time (s)

po

wer

(W

)

daylight

eclipse

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Trade studies summary• This trade study examines the option of using only non-

rechargeable battery vs solar cells+rechargeable battery system as satellite’s power source. To meet NANOsat requirements, a power source that has low mass and small size is desirable for this mission.

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Solar cells vs Battery• Factors that effect trade study: - Total power consumption and power profile during the

mission - Mission life of satellite : 24 hours - Mass and area constraints that come from NANOsat

requirement * Solar cells can produce lots of power with little increase in total mass * If power consumption is large, the mass and size of non-rechargeable battery could be greater. - Choice of orbit and type of attitude control - Operating environment

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Solar cells vs Battery• We choose solar cells as the main power source with

rechargeable battery to store energy and to provide power during eclipse

- Solar cells+ battery has little increase in mass for larger

increase in power consumption - Fewer non rechargeable battery that is qualified for

space application

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Solar cell vs Battery

Solar cell+battery Non rechargeable battery

Mass Less More

Power/mass High Low

Mission length Weeks-months Few hours-one/two days

Attitude/Orbit control

Optional / Dependant

Independant

Effect of Failure No power during eclipse

No power throughout mission

Thermal control Less dependant Extremely dependant

Cost Do you have cost estimates?

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Solar cells•Ultra triple junction Gallium Arsenide solar cell

• 28.0 % BOL efficiency

•2.31 V, 16.3 mA/cm² ( ~0.96 W/ cell)

•2.3gCustomized size:

3.69 cm

6.85cm

Area per cell: ~ 25cm²

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Solar cells layout & assemblies

Sides (A)

• 5 identical solar panels• Power source: Direct sunlight, at 45 ° angle from normal direction of plane• 91 UTJ GaAs solar cells per side - 13 solar cells per string : 30 V - 7 strings : 2.85 A• 40.68 Watt minimum per side during daylight • Dimension : 48 cm x 48 cm• Area: 2275 cm²

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Assumption:

•2 sides are exposed in direct sunlight at one time.

•Solar flux is at constant value of 1353 W/m^2

•Worst case hot temperature was used to find thermal efficiency for a solar cell , therefore the calculated power output from the a solar panel is the minimum value.

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Rechargeable Batteries

• Saft MPS 176065 Lithium-ion cells

• 8 cells in series in a battery box

• Capacity: 5.8 Ah

• Mean voltage: 3.6 V

• Battery mass: 1.4 kg (including casing)

• Maximum DOD: 70% for <500 cycles

• Charging method: Constant Voltage-constant current + balancing

• Space qualified

• 2 battery boxes (for redundant operation with one unit failed)

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• Energy storage requirements:

Peak power load: ~90 W Discharge time: 36 min (maximum) Charging time: 0.9 to 1 hour Charge/Discharge cycle / day: 16 Required battery capacity: 2.6 Ah for 75% DOD

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Demonstrationpower gererated and power stored per orbit

0

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300

0 0.2 0.4 0.6 0.8 1 1.2 1.4 1.6

time (hr)

po

we

rge

ne

rate

d (

W)

/ c

ap

ac

ity

(w

.h)

ideal power from solar panel

battery capacity in W.h

power discharge at 1.5C

*will recheck the values this weekend

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Power Management & Distribution

•SmallSat power management electronics

•28V unregulated; MPPT; Modular & Scalable from 30W to 300W

•Consist of 3 main elements:

- Battery Charge Regulator (BCR)

- Power Conditioning Module (PCM)

- Power Distribution Module (PDM)

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Power Distribution Design

Power Bus

(~28 V)

DC-DC step down converter

DC-DC Step downconverter

ADCS

FCS

COMM

OBC

Payload (Camera)

Thermalcontrol

Propulsion Power

5V supply line

12 V supply line

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Power Distribution

** non continuous operation

•Continuous power per one orbit period: 23.7 W

•Maximum power per one orbit period: 89.7 W

•Average power per one orbit period: 47 W

•Average power during daylight: 42.5 W

•Average power during eclipse: 53 W

ADCS 5%

Orbit Control (FCS)5%

OBC3%

Power management7%

Communication4%

Camera (idle)3%

Thermal control **13%

Propulsion **40%

Camera **20%