Spacecraft Power Systems The Generation and Storage of Electrical Power D. B. Kanipe Aero 401 February 9, 2016
Spacecraft Power Systems
The Generation and Storage of Electrical Power
D. B. KanipeAero 401 February 9, 2016
Power Systemso Batteries Solar Cells + Batteries Fuel Cells
RTG Nuclear Reactors ?
o Functions of the Power System n Controls the generation, storage, and efficient use of power
n Provides protection against cascading failures
n Provides redundant paths or components in case of failure
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Power System Design Drivers (½)
o Customer/User requirementso Mission, ConOpso Spacecraft configuration
n Mass constraintsn Dimensional constraintsn Launch Vehicle constraintsn Thermal constraints
o Expected lifetime
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o Attitude control systemn Pointing requirementsn Viewing requirements
o Orbit or trajectoryn With respect to the sun
o Payload requirementsn Voltage, currentn Duty cycle, peak loadn Fault protection
o Mission constraintsn Maneuver ratesn G-loads create inertial loads
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Power System Design Drivers (2/2)
Power System Functional Block Diagram
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Power Source
SourceControl
Power Distribution, Main Bus Control & Main Bus Protection
Power Conditioning Load
Energy Storage
Energy Storage Control
MainBus
- Batteries- Solar- RTG- Fuel Cells- Nuclear- R Dynamic- Solar Dynamic
- Shunt Regulator- Series Regulator- Shorting Switch
Array
- DC-DC conversion- DC-AC conversion- Voltage regulator
- Battery charge control- Voltage regulator
Design Practice (1/2)
o Direct Current Switchingn Switches or relays: positive line to an element with a direct
connection to “ground’ on negative sideo Therefore, element is inert until commanded
o Arc Suppressionn Locate as close to the source of the arc as possiblen Current-carrying elements should not be exposed to the
ambient plasmao Conductive cables, connectors, solar array edges
o Modularityn Simplifies testingn Easier element replacementn Reduces “collateral” damage
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Design Practice (2/2)
o Groundingn Cause of some debate among EEsn Common ground preferable to individual component grounding
o Easier to maintain a common potentialo Less likely to disturb sensitive componentso Can be difficult to do in large spacecraft
n Sometimes it is necessary to completely isolate an element from other spacecraft noise
o Continuityn Avoid buildup of static potential; i.e., any voltage differencen Any shielding must have continuity and a common ground
o Complexityn KISS
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Battery Design Considerations
o Physicaln Size, mass, environmental requirements
o Electricaln Voltage n Current loading n Duty cyclesn Limits on depth-of-dischargen Fault recovery
o Programmatic n Cost, reliability, maintainability, safety
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Batteries: Definition of Terms (1/2)
o Charge Capacity, Cchgn Total electric charge stored in a battery; measured in amp-
hours (e.g., 40A for 1 hour = 40Ah)
o Average Discharge Voltage, Vavgn (Number of cells in series) * (Cell discharge voltage)
o Energy Capacity, Ebatn Total energy stored in a battery;
[Cchg* Vavg] (Joules or watt-hours)
o Depth of Discharge, DODn Percent of battery capacity used in discharge cyclen 75% DOD means 25% remainingn Try to limit DOD to promote longer cycle life
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Batteries: Definition of Terms (2/2)
o Charge Rate, Rchgn Rate at which the battery can accept charge
(amps/unit time)
o Energy Density, ebatn Energy per unit mass stored in batteryn Joules/kg or Watt-hours/kg
o Two categories of batteriesn Primaryn Secondary
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Primary Batteries
o Long storage capability (missile in a silo)o Dry (without electrolyte) until needed
n Activate by introducing electrolyte into dry batteryn Electrolyte may be solid at room temperature
Activate heater to melt electrolyte. (Thermal battery)
o Typically have a fairly large energy densityo Used for early major mission events
n Short durationn May be isolated from major power busn Usually non-rechargeablen Mass penalty
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Secondary Batteries (1/2)
o Lower energy density, but rechargeableo Requires DOD management
n LEO – eclipse is about 40% of the orbitn 12 – 16 discharge cycles per dayn Leads to battery degradation and lifetime reduction
o Maximum allowable DOD:Energy required during eclipse
Stored battery energy
PLtd PLtdCchgVavg Ebat
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DOD =
=
PL = load powertd = discharge timeCchg = charge capacityVavg = average discharge voltageEbat = total battery energy capacity
=
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Battery Type
Silver–zinc (Ag-Zn) ………………………………..
Silver-cadmium (Ag-Cd) ………………………..
Nickel-cadmium (Ni-Cd) …………………………
Nickel-hydrogen (Ni-H2) …………………………
Nickel-metal hydride (Ni-MH) ………………..
Lithium Thionyl Chloride (Li-SOCl2) ……….
Lithium Vanadium Pentoxide (Li-V2O5) …. Lithium Sulfur Dioxide (Li-SO2) ……………..
Energy Density
120 – 130 (W-hr)/kg
60 – 70 (W-hr)/kg
20 – 30 (W-hr)/kg
60 – 70 (W-hr)/kg
120 – 130 (W-hr)/kg
650 (W-hr)/kg
250 (W-hr)/kg
50 – 80 (W-hr)/kg
Secondary Batteries (2/2)
DOD Management
o Typically, a LEO spacecraft spends 40% of its time discharging and 60% charging
o DOD during eclipse is limited by the rate at which its batteries can be restored in sunlight by solar arrays
o All expended energy must be restored or net draino Driver is the charge rate, Rchg
n DOD limited to 7-8% per orbitn Battery temperature can affect charge rate
o Battery generally must be charged at a voltage > Vavg(~20% higher) to restore full charge
o This is a driver in the solar array design
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Solar Arrays
o Photoelectric Effectn Electrons are emitted from matter as a result of absorption of
short wavelength electromagnetic radiation such as visible light.
o Originally limited to spacecraft skin acreageo Deployable panels more flexible, but more complexo Solar Cell Characteristics
n 1st order: V decreases as T increases (and vice versa)n 2nd order: I increases as T increases, BUT
o Only about 10% relative to the voltage dropn Therefore, overall power output is reduced as temperature
increases. P = I*Vo May need radiators to remove excess heat
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Solar Cell Capability
o Delivered electrical power:Pe = ФeA(1- I) Ф = solar flux (W/m2)
e = cell efficiency (≈15% for silicon)A = area
I = parasitic losses (≈10%)
o Nominal solar flux density at earth: 1353 W/m2 at 1 AUФ = W(a/d)2cos(ө) W = Nominal solar flux
a = Mean earth-sun distanced = actual earth-sun distanceө = panel inclination
o Cell efficiency (t)eEOL = eBOLe-0.043T T = time in orbit years
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Maximum Power Point (MPP)
o Desirable to operate at the MPP if possiblen Minimize mass and maximize efficiency
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I
V
MPP
Maximum area rectangle under the IV curve
Sun Tracking
o Ideal situation: sun normal to the arrayo Cosine rule applies – up to a point
Optimum
Up to 45-60˚cosine functionworks, then falls off rapidly
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Beta and Alpha
o Beta, ßn Angle between a line from the sun to the
center of the earth, and spacecraft orbit
o Alpha, αn Apparent rotation of sun angle from
spacecraft pov during its orbit. α = 0-36019
Solar-to-Electric
o Efficiency of solar cellsn Gallium arsenide solar cells (Ga-As)
o More efficient (20%) and radiation toleranto More expensive
n Crystalline Silicon Cellso 11-16%, 18-20%, >20%?
n Multi-junction (multi-layer cells)o Top layer converts light in the visible rangeo Bottom layer(s) optimized for infraredo Up to 30% efficiencyo Not surprisingly, very expensive
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Radioisotope Thermoelectric Generator (RTG)
o Converts heat energy generated by radioisotope decay into DC energy via thermoelectric effectn Plutonium 238, 238Pun Strontium 90, 90Sr
o Complicated ground handlingo RTG radiation
n Alpha raysn Detrimental to spacecraft electronicsn Clothing (or paper) will stop alpha raysn Don’t inhale 238Pu dustn 238Pu pellets are a ceramic form – no dust if exposed
o Expensive but effective and reliable21
Fuel Cells
o Direct conversion of chemical energy into electricity
o More efficient than batterieso Oxidizer and fuel fed into a cello Electricity generated from oxidation
reaction in the cello Space applications use oxygen/hydrogeno By product: watero ~ 35% efficiency
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Power Conditioning & Control
o Voltage from power source, especially solar arrays, may fluctuate
o Power conditioning functionsn Control solar array outputn Control battery charge/discharge cyclen Regulate voltage supplied to spacecraft
systems
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Additional Power Sourceso Nuclear Reactorso Dynamic Isotope Systemso Alkali Metal Thermal-to-Electric
Conversion (AMTEC)o Solar Dynamic
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Backup
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Dissipative Systems
o Simplero Not in series with array output
Dissipates current in excess of instantaneous load requirement
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Solar Array
Shunt Battery Charge
Controller
Spacecraft Loads
Battery
Non-dissipative Systems (PPT)
o In series regulation of solar power
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o Usually reserved for large spacecraft
Solar Array
Battery Charge
Controller
Spacecraft Loads
Battery
Battery Discharge Controller
PPT