Transcript AE 265

AE 265
GATEWAY TO
SPACE
Session 17 – Electrical Power
Subsystem (EPS)
Presented by Leon Searl (ITTC)
EPS Overview
Subsystem Purpose
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
EPS Overview
Requirements - Examples
The Asiasat 3S commercial
communications satellite Power
Subsystem must supply ~10,000W over
15+ years of on-orbit service.
Asiasat 3S
Courtesy of Hughes Space and
Communications Company
The XTE Power Subsystem must
supply 800 watts at 28V  7V
over 5 years of on-orbit service.
XTE spacecraft
Courtesy of NASA
The CGRO Power Subsystem must
supply 3,600 watts at 22 to 35V
over 5 years of on-orbit service.
CGRO spacecraft
Courtesy of NASA
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
Energy Source
Energy
Storage
Power
Generation
Electrical
Power
Subsystem
Power
Regulation,
Distribution
& Control
Payload &
Subsystem
Loads
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
Primary
Pwr A
Primary
Pwr B
Cntl A
Charger
A
Secondary
Pwr A
Charger
B
Secondary
Pwr B
Cntl B
Pwr
Conv A
Pwr
Conv B
Bus B
EPS Overview
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.
• Most Common Types
•Solar Cell
•Generally use inside Mars
orbit
•Radioisotope Thermal Generator
•Generally used beyond Mars
orbit
Terra spacecraft (EOS AM-1)
Courtesy of NASA
EPS Overview
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, pyrofirings, peak loading, and/or contingency operations.
Energy storage is usually done
via batteries.
NiCd Battery
Courtesy 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.
Power Source Considerations
• Desirable Properties
• Solar Cells
– Orbits
• Eclipse
– Distance From Sun
•
•
•
•
•
Radioisotope Thermoelelectric Generators
Fuel Cells
Nuclear Reactor
Batteries
Other
Energy Sources
• 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
Solar Cells
Photovoltaic Solar Cells Comparison
Note: efficiencies are for single cells, not arrays
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
Solar Cells
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.
Negative
Contact
Photons
- - - - - +
+
+
+
+ + +
+
+
+
n-layer
Junction Region
p-layer
Positive Contact
Solar Array Degradations
Angle of Incidence
Plot of Typical Reduction in Cell Output With Increasing Angle of Incidence
Cell Output Power (W)
0.4
0.3
0.2
0.1
10
30
50
Angle of Incidence
70
90
Orbital Considerations
Eclipses
Orbital Considerations
Eclipse Shadowing - LEOs
Earth eclipse periods vary during the year for low-Earth orbit satellites.
E
X
A
M
P
L
E
Eclipse Period/Orbit (Min)
38
36
Vernal Equinox
Autumnal Equinox
34
E
X
A
M
P
L
E
32
30
28
26
24
22
20
18
16
Winter Solstice
Jan
Summer Solstice
June
Winter Solstice
Dec
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).
Eclipse
Period/Orbit
(hr)
Example
1.2
1.1
1.0
0.9
0.8
0.7
0.6
0.5
0.4
0.3
0.2
0.1
0
Autumnal
Equinox
Vernal
Equinox
Jan
Feb
March
April
May
June
July
Aug
Sept
Oct
Nov
Dec
Orbital Considerations
Distance from the Sun
Available Solar Energy (W/m2)
The amount of solar power available falls off with the
square of the distance from the Sun.
3000
Venus
2000
Earth
1000
Mars
Jupiter
100
300
500
Distance from the Sun (kmx106)
700
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
Solar Arrays
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
ACE - 4 fixed arrays
Courtesy of NASA
FAST - body-mounted cells,
approximately 40% illuminated at a given time
Courtesy of NASA
TDRS - articulating arrays
Courtesy of NASA
Solar Arrays
Solar Arrays
Radioisotope Thermoelectric Generators
Overview
RTGs convert thermal energy released during the decay of an isotope (usually
Plutonium-238) into electrical energy.
RTGs are found on virtually
all U.S. deep space missions
destined for Mars orbit or
farther out in the solar
system.
NASA’s Cassini satellite includes 3 RTGs used to
produce >630W of power throughout an 11-year
mission to Saturn
Courtesy of NASA and NSSDC
Fuel Cells
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
• 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.
-
+
Anode
Hydrogen
Cathode
Electrolyte
(KOH)
Oxygen
Water
Fuel Cells
Examples
Gemini VII
Courtesy of NASA
Gemini Program
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.
Apollo 13
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.
Fuel Cells
Example: Space Shuttle
• 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 Cell
Courtesy of NASA
Nuclear Reactors
Overview
• Take the heat from nuclear reactions
(~700C) and convert it into electrical
energy.
The TOPAZ II reactor is the current
model being used on Russian
satellites. This one was purchased
by the USAF for testing.
• 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.
Topaz II Reactor
Courtesy of USAF
Nuclear Reactors
Example: Mars Transfer Vehicles
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 Concept
Courtesy of Northrop Grumman
Nuclear power systems
for manned Mars
transfer vehicles may
provide propulsion and
>25kW of electrical
energy.
Mars Transfer Vehicle Concept
Courtesy of NASA
Batteries
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.
Landsat-7 Batteries (circled) being readied for flight
Courtesy of NASA
• 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.
NiCd Battery - Approximate Dimensions:
0.3m x 0.25m x 0.15m
Courtesy of NASA/JPL/Caltech
Batteries
Primary 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
Batteries
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:
1. Nickel Cadmium
The standard satellite battery for the past 30+ years
2. Nickel Hydrogen
Becoming the ‘battery of choice’ for most satellite
applications
3. Lithium Ion
Superior energy density and considered most promising
for long-term; used in electronic devices but not yet
satellite flight-tested.
4. Nickel Metal Hydride
Twice the capacity of equivalently sized NiCad batteries
but more sensitive to overcharging.
5. Sodium Sulfur
Approximately twice the capacity of Lithium Ion – tested
on STS-87
Batteries
Profile of Battery Charge/Discharge Cycles
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.
Batteries
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.
-What happened?
• 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.
Other Energy Storage Devices
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
NASA’s Glenn Research
Center is developing a
flywheel technology
test system for the
International Space
Station launched in
2001.
• 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.
Flywheel Technology for
International Space Station
Courtesy of NASA
Energy Generation & Storage Example
Lunar Base Concept
Solar Array Primary Power
Regenerative Fuel Cell
Cryogenic
Oxygen/Hydrogen
Storage
Photovoltaic/Regenerative Fuel Cell Power System for a Lunar Observatory
Courtesy of NASA
Power Regulation
Overview
• 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 Regulation
Solar Array Output Regulation
– 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
Power Regulation
Regulated Bus Voltage
Regulated Bus - voltage should remain constant
with variations of less than 2%
Bus Voltage (V)
30
29
28
27
3
9
15
Time (hours)
21
Power Regulation
Unregulated Bus Voltage
Unregulated Bus - voltage will vary during the battery charge/discharge cycle
as shown here for the LEO TRMM satellite
Transition from Peak Power
Tracking to 12A Constant
Current battery charge mode
Enter Eclipse
Bus Voltage (V)
33
31
Begin Taper
to Trickle
29
27
Enter
Sunlight
10
30
50
Orbit Time (Minutes)
TRMM data courtesy of NASA
70
Power Distribution
Overview
• 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
Relationship Between Current and
Distribution Cable Mass
Critical Level
Power Control
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
• Two redundant Intel 8085
processors with autofailover
• 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
Reset
Hardware
Interface
Reset SR
Timer
Interface
Timer
SR
EPS Software
(embedded in
8085 EPS
Processors)
Housekeeping
Data
Acquisition
Interface
SR =
Service Request
CT&DH
Cmd
Command
SR
Interface
Command
s
Telemetry
Tlm
SR
Outputs
Output
Interface
ACE EPS Software Context Diagram
Courtesy of NASA
CT&DH
Telemetry
Interface
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 Launch
2/7/2001
MMIII Launch
VAFB 9/19/02
Clementine’s View
of Earth Over Lunar
North Pole Mar.
1994