AE 265

GATEWAY TO SPACE

Session 17 – Electrical Power Subsystem (EPS)

Presented by Leon Searl (ITTC)

The EPS, or more correctly, Power Generation, Storage, & Distribution Subsystem provides power to all spacecraft equipment:

Provides the power required for operation of subsystems and payloads at required current and voltage levels

Turns power to payloads and S/C subsystems on/off on command Ensures that payloads and satellite components are protected from

component power failures (e.g., short circuits) that could cause system- wide damage

Provides voltage, current, temperature measurements via telemetry for power management and status

Ensures that required power is available over the expected mission lifetime

Ensured through redundancy and backup power

Subsystem Purpose

EPS Overview

Requirements - Examples

The CGRO Power Subsystem must supply 3,600 watts at 22 to 35V over 5 years of on-orbit service.

The Asiasat 3S commercial communications satellite Power Subsystem must supply ~10,000W over 15+ years of on-orbit service.

The XTE Power Subsystem must supply 800 watts at 28V 7V

over 5 years of on-orbit service.

CGRO spacecraftCourtesy of NASA

XTE spacecraftCourtesy of NASA

Asiasat 3SCourtesy of Hughes Space and

Communications Company

EPS Overview

Energy Balance

A primary goal in power subsystem management is to maintain the spacecraft in positive energy balance

(Power-Positive).

Power available

Spacecraft load

requirements + losses

EPS Overview

Basic Subsystem Block Diagram

Electrical Power

Subsystem

Payload & Subsystem

Loads

Power Generation

Power Regulation, Distribution & Control

Energy Storage

EPS Overview

Energy Source

EPS Components

• Primary Power Source• Backup Power Source/ Energy Storage

– Primary power may not be available at all times• Power Conversion

– From Primary/Secondary voltage to subsystem required voltages• Redundant power bus

– Power Components can fail• Smart power management

– Automatic switching between Primary and Secondary source– Current limiting (protection from short circuit)– Alerts to CTDH on voltages/loads

• Telemetry– Voltages, loads, switch positions, temperatures

• Radiation Tolerance• Heat Dissipation

EPS Diagram

Bus A Bus BPrimaryPwr A

PrimaryPwr B

ChargerA

ChargerB

SecondaryPwr B

SecondaryPwr A

PwrConv A

PwrConv B

Cntl A Cntl B

Power Generation Function

Power Generation...

• Source of power for supplying the spacecraft with the power required to sustain platform and payload operations.

• Provides excess power that may be stored for later use.

Terra spacecraft (EOS AM-1)Courtesy of NASA

EPS Overview

• Most Common Types•Solar Cell

•Generally use inside Mars orbit

•Radioisotope Thermal Generator•Generally used beyond Mars orbit

Energy Storage Function

Energy Storage...

• Preserves power for use when primary power generation sources are unavailable or insufficient to satisfy spacecraft power requirements - e.g. launch operations, eclipse periods, pyro-firings, peak loading, and/or contingency operations.

Energy storage is usually done via batteries.

NiCd BatteryCourtesy of NASA/JPL/Caltech

EPS Overview

Power Regulation, Distribution & Control Function

• Manages power distribution to the satellite’s loads.

• Ensures that the necessary power is delivered at the correct voltage and current as requested by each payload and subsystem load.

• Accommodates rapid changes in the load requirements as loads are power-cycled or change modes.

Typically, the Power Distribution & Control function includes a processor, various relays, fuses, shunts, etc.

EPS Overview

Power Source Considerations

• Desirable Properties• Solar Cells

– Orbits• Eclipse

– Distance From Sun• Radioisotope Thermoelelectric Generators• Fuel Cells• Nuclear Reactor• Batteries• Other

• Desirable Properties of Spacecraft Power Sources– Safe (nonhazardous to personnel/equipment)– Reliable– Low weight and volume, high power density– Compatible with spacecraft and mission– Available when needed in schedule– Low cost

Energy Sources

Photovoltaic Solar Cells Comparison

Note: efficiencies are for single cells, not arrays

Solar Cells

Hughes PanAmSat-6B uses ~60m2 of Dual-Junction cells (Gallium Arsenide and Gallium Indium

Phosphide) to provide the 10kW of power required to operate in GEO orbit. PanAmSat-6B

Courtesy of Hughes Space and Communications Company

How A Solar Cell Works

A cell consists of a semi-conductor ‘sandwich’ with an electron-rich layer (n) on top and an electron-poor layer (p) on the bottom. (The sandwich could also be created with the p-layer on top, but the n-layer on top design has higher radiation resistance and is more commonly used in spacecraft.)

When solar photons impinge on the junction layer between the two materials, current flows from the top to the bottom layer. This flow of electrical energy is captured and used to supply spacecraft needs.

For space applications, solar cells range in size from ~2x2 cm up to ~6x4 cm.

++ +

++++++

+

-------- n-layer

p-layerJunction Region

Positive Contact

Negative Contact

Photons

Solar Cells

0.1

0.2

0.3

0.4

10 30 50 70 90

Plot of Typical Reduction in Cell Output With Increasing Angle of Incidence

Angle of Incidence

Cel

l O

utp

ut

Po

wer

(W

)Angle of Incidence

Solar Array Degradations

Orbital ConsiderationsEclipses

Eclipse Shadowing - LEOs

Earth eclipse periods vary during the year for low-Earth orbit satellites.

16

18

20

22

24

26

28

30

32

34

36

38

Ec

lip

se

Pe

rio

d/O

rbit

(M

in)

Vernal Equinox Autumnal Equinox

Winter Solstice Summer Solstice Winter Solstice

Jan DecJune

EXAMPLE

EXAMPLE

Orbital Considerations

Eclipse Shadowing - GEOs

GEOs also experience once per-orbit eclipses but only during eclipse “seasons” at vernal and autumnal equinoxes

(1.2 hours maximum eclipse at GEO).

Jan Feb March April May June July Aug Sept Oct Nov Dec

Eclipse Period/Orbit

(hr)1.21.11.00.90.80.70.60.50.40.30.20.10

Vernal Equinox

Autumnal Equinox

Example

Orbital Considerations

Distance from the Sun

Ava

ilab

le S

ola

r E

ner

gy

(W/m

2 )

100 300 500 700Distance from the Sun (kmx106)

1000

2000

3000

Mars

Earth

Venus

Jupiter

The amount of solar power available falls off with the square of the distance from the Sun.

Orbital Considerations

Distance from the Sun: Cassini Example

Prior to deciding to go with an RTG power source, NASA investigated solar arrays as a possible power source for the Cassini mission to Saturn, but 500 m2 of array would have been required to supply the 700 W required to operate the

spacecraft at Saturn.

Solar array scaling for orbits beyond Earth Courtesy of NASA/JPL/Caltech

Orbital Considerations

Planar Arrays vs. Body-Mounting

Body-mounted• limits the quantity of cells that can be used• not dependent on deployment or articulation

components• useful only on spinning satellites

Planar Array-mounted• use deployment mechanisms (launch vehicle

packaging constraints)• may be articulated to track the sun

FAST - body-mounted cells, approximately 40% illuminated at a given time

Courtesy of NASA

ACE - 4 fixed arrays Courtesy of NASA

TDRS - articulating arraysCourtesy of NASA

Solar Arrays

Solar Arrays

Solar Arrays

Radioisotope Thermoelectric Generators

RTGs convert thermal energy released during the decay of an isotope (usually Plutonium-238) into electrical energy.

NASA’s Cassini satellite includes 3 RTGs used to produce >630W of power throughout an 11-year

mission to SaturnCourtesy of NASA and NSSDC

RTGs are found on virtually all U.S. deep space missions

destined for Mars orbit or farther out in the solar

system.

Overview

Overview

• A fuel cell provides a stored energy source - cryogenically stored hydrogen and oxygen - that can be expended as required to support spacecraft power needs. The total amount of energy available is limited by the quantity of the stored fuels.

• Fuel cells have only been used in the U.S. for human spaceflight (Gemini, Apollo, Shuttle).

• Fuel cells are reuseable and re-startable

• Typical fuel cell energy conversion efficiency is ~65%.

Fuel Cells

How Fuel Cells Generate Electricity

+ - Cathode Anode

Electrolyte (KOH)Hydrogen

Oxygen

Water

• A fuel cell consists of two electrodes sandwiched around an electrolyte. As oxygen passes over one electrode and hydrogen over the other, electricity, water and heat are generated.

Fuel Cells

2 days into the Apollo 13 mission, commands to heat and stir a cryogenic oxygen tank used by the fuel cells and for crew respiration resulted in an explosion of the tank. This explosion ruptured the other oxygen tank, leaving the crew with no electrical power.

Later investigation found that the cryogenic oxygen system had been redesigned to run off of 65V ground power as well as the 28V spacecraft power, but the heating component was not replaced to run at this higher voltage. During the last ground test prior to launch, oxygen was ‘boiled’ out of the tank during an 8 hour period, overstressing the heater and causing it to fail during the mission.

ExamplesFuel Cells

Gemini VIICourtesy of NASA

Long duration (8 & 14 day) Gemini missions used fuel cells instead of batteries to provide

the 5,500 amp-hours of energy to support loads of up to 2kW at 25 +/-2V.

Gemini Program

Apollo 13

• Each shuttle carries three fuel cells to meet orbiter power requirements:

• Bus Voltage: 28 +/-4V• ~14kW average subsystem load• ~7 kW payload load

• Fuel cell specifications:• Dimensions: 90cm (height), 97cm

(width), 258cm (length)• Mass: 116 kg• Power Range: (2kW @ 32.5V &

61.5A) through (12 kW @ 27.5V & 436A)

• Each fuel cell = 96 individual cells separated into 3 bays

• Maximum Output Power: 21 kW• Short Duration (15 min.) Maximum: 36 KW

STS Fuel CellCourtesy of NASA

Example: Space ShuttleFuel Cells

Overview

• Take the heat from nuclear reactions (~700C) and convert it into electrical energy.

• Reactors can provide kilowatts to megawatts of power.

• Numerous Russian Cosmos satellites have been outfitted with these reactors since the 1970s.

• Safety issues have stopped the U.S. from ever flying a nuclear reactor.

• Future human missions to Mars and large lunar bases may require nuclear reactors to provide the megawatt power levels required in the most mass/cost effective manner.

The TOPAZ II reactor is the current model being used on Russian

satellites. This one was purchased by the USAF for testing.

Topaz II ReactorCourtesy of USAF

Nuclear Reactors

Example: Mars Transfer Vehicles

Nuclear power systems for manned Mars

transfer vehicles may provide propulsion and

>25kW of electrical energy.

Mars Transfer Vehicle ConceptCourtesy of NASA

Nuclear Reactors

Prometheus 1 will have a nuclear fission reactor

powered electric propulsion system enabling to orbit 3

Jovian moons during a single mission

Prometheus 1 Spacecraft ConceptCourtesy of Northrop Grumman

Overview

• Over the past 30+ years, energy storage on satellites has been performed by batteries.

• Although battery performance is usually fairly stable over the first ~3 years on-orbit, gradual performance degradation and component failures have been responsible for many satellites being taken out of service.

• Current battery technology efforts are focused on improving the energy storage efficiency of batteries and improving their longevity.

• The problem is that these advanced technologies have little or no flight heritage at this time, so, although the theoretical gains are great, so are the risks.

Landsat-7 Batteries (circled) being readied for flightCourtesy of NASA

Batteries

NiCd Battery - Approximate Dimensions:0.3m x 0.25m x 0.15m

Courtesy of NASA/JPL/Caltech

Primary Batteries

Batteries

Primary batteries are non-rechargeable batteries used for short missions (especially suborbital), single-use purposes (such firing of pyrotechnic devices) or for infrequent use of a high-power component or subsystem on a longer mission. Most common type used is Silver Zinc.Desirable attributes for primary batteries include:• long shelf life• high energy density• nonhazardous• wide range of operating temperatures

Secondary Batteries

• Rechargeable batteries for use on satellites have greatly increased in efficiency and energy storage capability over the past 30 years.

• The primary batteries that have been used - or are planned - for satellite applications are:

The standard satellite battery for the past 30+ years

Becoming the ‘battery of choice’ for most satellite applications

Superior energy density and considered most promising for long-term; used in electronic devices but not yet satellite flight-tested.

Twice the capacity of equivalently sized NiCad batteries but more sensitive to overcharging.

Approximately twice the capacity of Lithium Ion – tested on STS-87

1. Nickel Cadmium

2. Nickel Hydrogen

3. Lithium Ion

4. Nickel Metal Hydride

5. Sodium Sulfur

Batteries

Profile of Battery Charge/Discharge CyclesBatteries

The average LEO satellite goes through ~6,000 charge/discharge cycles each year with each charge cycle lasting ~60 minutes and discharge cycles of ~30 minutes.

Charging: Skipper Example

• In December 1995, a joint U.S. Defense Department/Russian military satellite named Skipper was launched from the Baikonur cosmodrome.

• The 250 kg, 150 cm diameter satellite was scheduled for a 3-day mission during which it would try to detect and identify incoming missiles.

• Within the first day of the mission, the satellite was lost.

• Investigators quickly identified the cause of the problem: the wiring between the solar arrays and the NiCd battery was installed backwards, so instead of charging the battery during daylight, the battery discharged.

• By the time that ground controllers identified what had happened, the battery had been completely drained of energy and the satellite was dead.

Batteries

-What happened?

Flywheels

• Two counter-rotating flywheel modules will be placed in the same orbital replacement unit (ORU) slots as the current Nickel Hydrogen batteries

• Total of 2.4kW of energy storage capability

• Maximum speed: 60,000 rpm

• Magnetic bearings will be used to minimize friction.

• Composite construction will be used to minimize mass.

• The system will also test the capability of this system to generate differential torques that can be used for attitude control.

NASA’s Glenn Research Center is developing a

flywheel technology test system for the International Space Station launched in

2001.

Flywheel Technology for International Space Station

Courtesy of NASA

Other Energy Storage Devices

Lunar Base Concept

Solar Array Primary PowerRegenerative Fuel Cell

Cryogenic Oxygen/Hydrogen

Storage

Photovoltaic/Regenerative Fuel Cell Power System for a Lunar ObservatoryCourtesy of NASA

Energy Generation & Storage Example

• Power regulation method depends on energy source

• Power regulation has three main functions:– regulate and control the energy source output– regulate bus voltage– charge the energy storage (covered in Control)

Power RegulationOverview

– Power generated must be controlled to prevent battery overcharging and undesired spacecraft heating

– Two main control techniques (covered in detail later):

• Peak-Power Tracker (PPT) is a nondissipative subsystem because it extracts the exact power a spacecraft requires up to array’s peak power\

– Example: DC-DC converter, Switching Regulator• Direct-Energy Transfer (DET) is a dissipative

subsystem because it dissipates power not used by the loads - commonly uses shunt regulation to maintain bus voltage at predetermined level

– Example: Linear Regulator

Solar Array Output Regulation

Power Regulation

Regulated Bus VoltageB

us

Vo

ltag

e (V

)

3 9 15

27

28

29

Time (hours)21

30

Regulated Bus - voltage should remain constant with variations of less than 2%

Power Regulation

Unregulated Bus VoltageB

us

Vo

ltag

e (V

)

Orbit Time (Minutes)

Unregulated Bus - voltage will vary during the battery charge/discharge cycle as shown here for the LEO TRMM satellite

Enter Sunlight

Transition from Peak Power Tracking to 12A Constant

Current battery charge mode

10 30 50

27

29

31

70

33

Begin Taper to Trickle

Enter Eclipse

TRMM data courtesy of NASA

Power Regulation

• Spacecraft power distribution subsystem consists of:– electrical bus– fault protection– cabling– switching gear to turn power on and off to spacecraft loads– command decoders to command specific load relays

• Design of PDS strives to minimize power losses and mass while maintaining survivability, cost, reliability, and power quality

• Power switches are normally mechanical relays because of their reliability, flight history, and low power dissipation (solid-state relays also used)

• Power systems normally DC because s/c generates DC. Conversion to AC requires more electronics & mass

Power Distribution

Overview

Relationship Between Current and Distribution Cable Mass

Critical LevelCritical Level

Power Distribution

EPS Processor Functions

The Power Subsystem Processor performs the following functions:

• generates power subsystem health and status telemetry

• processes commands for the power subsystem• controls battery charging/discharging• controls energy transfer from solar arrays• controls the bus voltage• controls power switching of loads• contains automatic load shedding capability to

safe the spacecraft if a power-negative situation exists

Power Control

Power Subsystem Processor: ACE

Timer Interface CT&DH

CommandInterface

CT&DHTelemetryInterfaceOutput

Interface

Data Acquisition Interface

Reset Hardware Interface

EPS Software (embedded in

8085 EPS Processors)

Timer SR

Cmd SR

Reset SR

HousekeepingOutputs

Tlm SR

Commands

Telemetry

SR = Service Request

• Two redundant Intel 8085 processors with auto-failover

• Receive and execute commands from CT&DH subsystem; collect, format, and transmit power system telemetry to the CT&DH subsystem

• Regulate the main bus

• Control the battery charging

• Control power switching of loads

ACE EPS Software Context DiagramCourtesy of NASA

Power Control

EPS Heat Handling

• Dissipation– Inefficiencies in the power system generates heat– Large amounts of heat dissipated by:

• Heat Pipe to external heat radiator• Heat Plate to external heat radiator

– Small amounts of heat dissipated by black body radiation– Heat may be transferred to parts of S/C that are cold

• Heaters– Power System may supply power for heaters in subsystems or

payloads that have critical low end operating temperatures

STS-98 LaunchSTS-98 Launch2/7/20012/7/2001

Clementine’s View Clementine’s View of Earth Over Lunar of Earth Over Lunar

North Pole Mar. North Pole Mar. 19941994

MMIII LaunchMMIII LaunchVAFB 9/19/02VAFB 9/19/02