Download Power Sources and Systems of Satellites and Deep Space Probes

Power Sources and Systems of Satellites
and Deep Space Probes
Gabriel Farkas
Space for Education, Education for Space
ESA Contract No. 4000117400/16NL/NDe
Specialized lectures
Power sources and systems of satellites and deep space probes
Space for Education, Education for Space
Spacecraft Power Systems
Intro
 The spacecraft electrical power system (EPS) generates, stores,
conditions, control, and distributes power within specified
voltage band to all bus and payload equipment
 Power System Functions:
Power generation, storage, distribution, and efficient use of
power
Protection against failures
Providing redundant path or components in case of failure
Power sources and systems of satellites and deep space probes
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Spacecraft Power Systems
Design Criteria
 Customer-User specifications
 Basic satellite/space probe/planetary rover parameters 
orbit altitude, orbit inclination, and mission duration  used to determine the:
orbit period, sunlight and eclipse durations, and solar  angle between orbit
plane and the Earth-Sun line
 Load power requirements in all phases of the mission: launch + ascent, transfer
orbit, parking orbit, operational orbit, disposal orbit
 Prior experiences on similar satellites that met requirements in the most
economical and mass effective manner
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Spacecraft Power Systems
Design Criteria
 Spacecraft configuration:
 Mass constrains
 Dimensional constrains
 Launch Vehicle Constrains
 Thermal Considerations
 Expected lifetime
 Major self-derived requirements:
 Solar array EOL power level
 Solar array pointing and rotation for sun orientation
 Battery capacity in Ah
 Battery DOD and charge control
 Bus voltage regulation
 EMI/EMC and ESD
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Spacecraft Power Systems
Design Criteria
 Arc Suppression:
 Locate as close to the source of arc as possible
 Current-carrying elements should not be exposed to the ambient plasma 
conductive cables, connectors, solar array edges
 Modularity:
 Simplifies testing
 Easier element replacement
 Reduced “collateral” damage
 Grounding:
 Cause of some debate among EEs
 Common ground preferable to individual component grounding
• Easier to maintain a common potential
• Less likely to disturb sensitive components
• Difficult to do in large spacecraft
 Sometimes necessary to completely isolate an element from other spacecraft noise
 Continuity
 Avoid buildup of static potential  any voltage difference
 Any shield must have continuity + common ground
Power sources and systems of satellites and deep space probes
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Power System Architecture
Power sources and systems of satellites and deep space probes
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Spacecraft Power Systems
Power System Functional Block
Diagram
Power Source
 Batteries
 Solar Array
 RTG
 Fuel Cell
 Nuclear
 Solar
Dynamic
…
Source
Control
Power Distribution
Main Bus Voltage Control
Main Bus Protection
Energy Storage Control
Power
Conditioning
Load
 DC-DC Conversion
 DC-AC Conversion
 Voltage Regulator
Energy Storage
 Shunt Regulator
 Series Regulator
 Shorting Switch Array
…
 Battery Charge Control
 Voltage Regulator
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Spacecraft Power Systems
Power System Options
Optimum energy sources for various power levels
and mission durations
Final selection of the energy
source  must meet multiple
criteria
Primary criteria are always 
low mass and low life-cycle cost
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Spacecraft Power Systems
Photovoltaic – Battery System
 Photovoltaic conversion  the most common source of electrical power in space
 Array of photovoltaic cells  powers the load and charges a battery during sunlight
 Battery  power the load during an eclipse
 Solar array output voltage is HIGHER  at the beginning of life (BOL) + when the array
is cold for several minutes after each eclipse
 Battery has a LOWER voltage  during discharge than during charge
28V
 Voltage must be regulated within specified
limits  a voltage regulator needed
 PV system  primarily consists of a solar
array + rechargeable battery + power
regulator to control the bus voltage
 Other components  various sensors to
make the array and the battery work
together
www.saftbatteries.com
Credit: NASA
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Spacecraft Power Systems
Solar Array
http://lasp.colorado.edu
 Solar Array (SA)  made of numerous PV cells
stacked in series-parallel connections  to
obtain the desired voltage and current
 Converts the incident photon energy into 
DC voltage
MAVEN Solar Panel
 Solar array works more like  a constant
current source over its normal operating rate
 Array power output  maximum at the kneepoint voltage
 Power output gradually degrades  with
increasing temperature, accumulated dose,
and at the end of life (EOL)
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Spacecraft Power Systems
 Battery  made of rechargeable electrochemical
cells connected in a series-parallel  to obtain
desired voltage and current
Battery
www.nasa.gov
 Terminal voltage  depends primarily on the state of
charge (SOC), and operating temperature
 Battery works more like  a constant voltage source
over its normal operating rate
 Average voltage during charge is higher than during
discharge
Nickel-Hydrogen batteries for Hubble
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Spacecraft Power Systems
Power Regulation
https://spaceequipment.airbusdefenceandspace.com
Power Supply Regulator
 Power regulation  primarily by battery charge and discharge converters + shunt
dissipator  to obtain desired U and I, and a mode controller that responds to the bus
voltage error signal
 Shunt dissipator  to control the bus voltage during sunlight
 Mode controller  sets operating mode based on the error signal  difference between
the actual and reference bus voltage
 Error signal value and its polarity (+/-)  mode controller activates  either the shunt
regulator, or the battery charge, resp. discharge regulator
Power sources and systems of satellites and deep space probes
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Spacecraft Power Systems
Power System Architectures
 Solar array, battery, shunt characteristics, load voltage
requirement  extremely important in selecting the power system architectures
1. Direct energy transfer (DET)  solar power transferred to loads with no
components in between except  load switching relays, fuses, solar array drive
 DET  subdivided into two classes:
• Fully Regulated Bus
• Sunlight Regulated Bus
Direct Energy Transfer Architecture
DET
Operating
point
PPT
Peak Power Tracking Architecture
www.ece.colorado.edu
2. Peak Power Tracker (PPT)  solar array output voltage  always set
at the value resulting in the maximum power transfer from the array  load
 Power converters between the array and load  matches the voltage
requirement and the array output voltage
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Spacecraft Power Systems
Fully Regulated Bus
 Fully Regulated DET bus (regulated bus)  the bus voltage is controlled within a few
percent during the entire orbit period   (2 – 5)% of the nominal voltage
 Components 
 Solar array drive (SAD)  slip rings +
motor + motor drive electronics 
continuously orients SA to face the sun
 Shunt Dissipator  during sunlight
(especially in the BOL) dissipates
unwanted power after meeting the
load power and the battery charge
power requirements
 Battery  stores energy to supply
power to the loads during eclipse
periods over the entire mission life
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Spacecraft Power Systems
Fully Regulated Bus
 Components (cont.) 
 Power regulator unit (PRU)  interface
between the solar array bus and the
battery
 Power distribution unit (PDU) 
ensures that all loads, except critical
and essential ones  powered through
switches and fuses
 Bus Voltage Controller  bus voltage
sensor + the reference voltage + error
signal amplifier
ISS Power Systems Status – all channels except 1B 
negative currents, meaning power is flowing out of
the Direct Current Switching Units to power loads.
1B channel is shunted  positive current.
Image: http://isslive.com
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Spacecraft Power Systems
Fully Regulated Bus
www.nasaspaceflight.com
The ISS Sequential Shunt Unit
http://spaceflight101.com
 Mode controller  automatically changes the EPS mode in response to the error signal:
 Shunt mode  during sunlight, if solar power exceeds the load and battery charge
requirements  shunts dissipate excess power, else UBUS will rise above the
allowable limit, during this mode  battery is charged as required
 Charge cut-back mode  when battery is approaching full charge  the charge rate
is cut back to control battery T; moreover  when solar power exceeds the load
requirement, but not enough to supply the required charge current to battery
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Spacecraft Power Systems
Fully Regulated Bus
www.nasaspaceflight.com
 Discharge mode  in the absence of solar power during an eclipse  battery is discharged
to maintain UBUS. Battery voltage  with decreasing state of charge  discharge convertors
must increase the boost ratio accordingly
 PRU bypass mode  in case of fault in any loads  fuse must be blown asap to minimize
UBUS decay. Delay in PRU response  battery is instantly connected to the bus by bypass
diode for quick delivery the battery energy to the fuse in case of a fault
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Spacecraft Power Systems
Fully Regulated Bus
 Components (cont.) 
 Battery Bus  a tap point directly off the
battery  during the launch and ascent
phases  PV array not deployed  battery
meets all the energy needs, for example all
electro-explosive devices (EED)
 Power and Energy Management Software (PEMS)  dedicated to the EPS performance,
health monitoring, control, and protection. In emergency cases, or planned operations 
sheds loads in a preset sequence if /when battery cannot support all loads
 Loads  payloads (transmitters, receivers, science instruments, etc.) + bus system loads;
most loads in satellites  constant power loads
 Ground Power Cord  to preserve the battery during pre-launch testing and final checks
 the on-board system uses external ground power via an umbilical cord
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Spacecraft Power Systems
Bus Voltage Control
 Regulated bus  in normal operation
maintains the bus voltage  between
specified upper and lower limit
 Voltage control scheme  for fully
regulated 120 V LEO bus with 3000 W
load (UBUS regulated within 5 V 
115-125 V):
 Discharge mode  when UBUS falls below the specified limit (a)
 Dead band mode (do nothing band)  when UBUS is between dead band limits (a–b)
 Charge mode  when UBUS rises above dead-band limit (b)
 Shunt mode  when battery is fully charged and solar array output power exceed the
load power requirements (c)
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Spacecraft Power Systems
Sun Regulated Bus
 If design objective  to minimize complexity
 to distribute power from both sources: solar
array + battery directly to the load
 Sun Regulated Bus (partially regulated bus, or
unregulated bus)  UBUS regulated by shunt
control during sunlight, and unregulated only
during an eclipse
 Difference between sun and fully regulated bus
 only in the power regulator unit
 Sun-regulated bus  has battery charge regulator, but no battery discharge converter 
discharges directly to the bus during eclipse through  battery discharge diode
 d(4)  only allows discharge from the battery, but blocks any uncontrolled charge current
 Most beneficial architecture in multiple battery systems in GEO  sunlight duration long
and the eclipse short
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Spacecraft Power Systems
Sun Regulated Bus
 Voltage regulation during sunlight  by
shunt control circuits
 Battery directly connected to the bus
without discharge converter 
UBUS = UBAT  bus voltage falls as the
battery discharges during eclipse, resp.
rises at recharging during sunlight
 Nominal 28 V bus voltage  typically
varies from 22-35 V during orbit period
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Spacecraft Power Systems
Peak Power Tracking Bus
 Solar array generates  power at higher
voltage at BOL and when cold coming out of
an eclipse  managed by shunt control
circuits
 Peak power tracking  activated only when
battery needs charging or the load demands
exceeds the solar array output
 Otherwise  excess power is left on the
array raising the array temperature
 In LEO  battery must be charged in a short
period  PPT allows maximum power to be
captured for several minutes after each
eclipse when the array is cold
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Spacecraft Power Systems
Peak Power Tracking Bus
 Main advantages of PPT  maximizes the SA output power all the time (reduce required
solar array area and mass)  no need for shunt and battery charge regulator
 Disadvantages  poor system efficiency due to power loss in PPT converter  loss
dissipated inside the spacecraft body  negative impact on thermal system
www.swri.org
Southwest Research Institute’s
space grade PPT
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Spacecraft Power Systems
Architecture Trades
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Spacecraft Power Systems
 The ISS  the largest and most
complex space structure ever built
The ISS Power System
 The ISS power system generates
105 kW using solar array 
 US modules (wings) 76 kW
 Russian modules  29 kW
 US solar modules  crystalline
silicon PV cells + coverglass against
space charged particles
https://spaceflight.nasa.gov
US part of power system
 Sequential shunt units operate at 20 kHz
 The seasonal sun-pointing  done by  gimbals
 The orbit sun-following by  drive and roll rings
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The ISS Power System
https://spaceflight.nasa.gov
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Spacecraft Power Systems
 During an eclipse  powered by
batteries via bi-directional battery
charge-discharge units
The ISS Power System
 Two interconnected power system:
 160/120 V US system
 120/28 V dual voltage RU system
 US solar array output voltage  160 V
(practically the highest U for LEO)
 160 V stepped down to 120 V using
DC-DC converter unit
US part of power system
 The two systems  independent, but interconnected via DC converters ARCU
RACU
http://electrical-pdf-articles.blogspot.com
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Spacecraft Power Systems
 The battery  made of 48 batt. packs,
each with 81 Ah Li-ion/NiH2 cells
The ISS Power System
 Two packs connected in series  batt.
voltage varies: 95-115 V
 Batt. replacement  every 5 years
(40000 charge/discharge cycles)
 Switching and fault protection  by
solid-state remote power controllers  trip
at over-current, over-voltage, under-voltage
US part of power system
https://web.archive.org/
 The ISS PS  controlled by a hierarchy of
redundant computers  sun tracking, battery
energy storage, thermal control, etc.
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Spacecraft Power Systems
The ISS Power System
 Static electricity  can be dangerous to
electronics and personnel  to put
everything in the ISS at the same potential
 grounding all structures and
components to a common point
www.grc.nasa.gov
 Great solar array area + operating 160 V
level  the nature of the single point
ground in high plasma in LEO  poses an
arcing problem
 Although the ISS maintains all components
at a common potential  it may differ
from the surrounding space environment
potential
 Precluding arcing  by two Plasma
Contactor Units (PCU) located on the Z1
PCUs installed on the ISS
truss  emit excess electrons into space
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Spacecraft Power Systems
The ISS Power System
 It creates  a plume of ionized xenon
gas  acts as a conductive bridge
between the ISS structure and ambient
space plasma
 Hollow cathode assembly  superheats
a xenon gas, producing its own plasma
 provides a path for free electrons 
reducing voltage to a safe level
 Xenon gas is a consumable  expected
to support 1,5 – 2 years of continuous
operation
 Protects  the array + other conduct.
surfaces from arcing, pitting, erosion
http://pages.erau.edu
The Plasma Contactor Unit
functional diagram
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Spacecraft Power Systems
Spacecraft Charging
 For a conducting spacecraft  the charges are on
the surfaces  surface charging
 The potential of ambient space plasma is
traditionally defined as zero  p = 0
 In the field of spacecraft charging  spacecraft
potential is relative to the space plasma potential,
which is defined as zero
P =! 0
 The spacecraft potential is floating relative to the
ambient plasma potential
 When a spacecraft potential: s is nonzero relative
to that of the ambient plasma  the spacecraft is
charged: s  0
Floating potential of a spacecraft  a potential
sheath is formed around the spacecraft
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Spacecraft Power Systems
Spacecraft Charging
 A uniformly charged spacecraft  has only one potential: s  uni-form/absolute charging
 A spacecraft composed of electrically separated surfaces  the potentials may be different
on different surfaces depending on the surface properties, and on the environment 
differential charging
 A spacecraft covered with connected conducting surfaces (i.e., spacecraft ground/frame)
and some unconnected or nonconducting surfaces, the charging of the frame  frame
charging
 Very energetic (MeV or ) electrons and ions  can penetrate deep into dielectrics 
deep dielectric/bulk charging
 Conductors  surface charging can occur but deep conductor charging no
 Dielectrics (insulators)  both surface and deep dielectric charging can occur depending on
to incoming electrons energy (surf. ch. below 70-100 keV, deep ch. at higher energy)
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Spacecraft Power Systems
Mechanism of Charging
 Surface charging  caused by the interaction of spacecraft surfaces with  the plasma
environment, solar radiation, high-energy electrons and magnetic fields
 Space system reaches electrical equilibrium with the space plasma by acquiring surface
charges  net current to the whole system and to the individual insulating surfaces is
zero  equilibrium condition - determines spacecraft’s surface potential V relative to the
surrounding plasma
act as a negative current
act as positive currents to the 𝑆
𝐼𝑁𝐸𝑇 𝑉 = 𝐼𝐸 𝑉 −[𝐼𝐼 𝑉 + 𝐼𝑆𝐸 𝑉 + 𝐼𝑆𝐼 𝑉 + 𝐼𝐵𝑆𝐸 𝑉 + 𝐼𝑃𝐻 𝑉 ] = 0
Total current
from the
spacecraft
surface
Incident
Incident
environmental environmental
electron
positive ion
current
current
Secondary
emitted electron
current due to IE
Secondary
emitted electron
current due to II
Back-scattered Photoelectron
electron current current due to
sunlight
due to IE
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Spacecraft Power Systems
Parasitic Structure Current
 A rule of thumb  for every m2 of exposed conductor in LEO at 100 V positive  a
parasitic structure current of  1 mA may be expected.
 Thus, for  100 m2 surface area on a 100 V bus  only  100 mA of structure current may
drain from the power system capacity  it is negligible compared to 100 A that a 10 kW
power system can deliver at 100 V.
 Voltages  than 160 V can be used in LEO with insulated cables, covered in a shielded
enclosure, and be encapsulating all connectors and circuit boards.
 Early in ISS design 270 V DC and 440 V 20 kHz AC
were considered, but, finally  160 V solar array
voltage and 120 V distribution system with stepdown converters for existing 28 V hardware.
Fig. Covered cable tray protecting against voltage breakdown
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Spacecraft Power Systems
System Voltage Options
 Early spacecraft (loads of a few hundred watts)  28 V
 Today’s spacecraft (bus voltages somewhat standardized)  28 V, 50 V, 70 V, 100 V,
120 V, 160V
 160 V limit  primarily comes from the bare conductor interaction with space plasma,
particularly in the low Earth orbit (LEO), potential sparking above  200 V
 If conductors are at high potentials relative to the plasma  snap-over may greatly
increase the electron current and the resulting power drain
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Spacecraft Power Systems
High-Power Systems
 Power levels for commercial, science, and military spacecraft  rising steadily with a
doubling time of  7 years
 Today’s GEO communications satellites  7-15 kW at 70-100 V; the ISS uses 
105 kW at 120 V
 Some strategic Defense Initiative
weapons platforms may require 
steady power in several MW and burst
power in hundreds of MW at voltages
up to 100 kV or more
 Power system up to 100 kW 
designed at voltages up to 200 V
 Distribution systems above 200 V 
considered high-voltage in space due to
various environmental conditions
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Spacecraft Power Systems
Scaling for Power Level
 For new development – characterized by an absence of the heritage design – the
following guideline may be applied:
𝒐𝒑𝒕𝒊𝒎𝒖𝒎 𝒗𝒐𝒍𝒕𝒂𝒈𝒆 = 0.025 𝑥 𝑝𝑜𝑤𝑒𝑟 𝑟𝑒𝑞𝑢𝑖𝑟𝑒𝑚𝑒𝑛𝑡
 This optimum voltage  shifted to the nearest standard voltage
 Example: a 5000 W payload power system would have 0.025 · 5000 = 125 V shifted to the
nearest standard voltage of 120 V.
 For AC systems  the power system mass depends on the power level and frequency:
𝑘𝑊
𝑀𝑎𝑠𝑠 𝑜𝑓 𝐴𝐶 𝑠𝑦𝑠𝑡𝑒𝑚 =
𝑓
𝛼
 = 0.5  for small systems (hundreds of W)
 = 0.75  for large systems (hundreds of kW)
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Spacecraft Power Systems
Large Communications Satellite
Buses
 Standard buses:
 100 V Bus
 70 V Bus
 Under 50 V Buses (42.5 V, 28V)

 100 V DC BSS-702TM Bus
 Standard fully regulated direct
energy transfer (DET) bus
 Dual voltage bus 
100 V DC for high power equip.
30 V DC for low power equip.
 Source  silicon triple junction
solar cells + NiH2 battery
http://spaceflight101.com
 A xenon electric ion-propulsion
system  used for N-S station
keeping  powered from 100 V bus
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Spacecraft Power Systems
Large Communications Satellite
Buses
 70 V A2100TM Bus
 Standard fully regulated DET bus
 Up to 15 kW (20 kW)
 Source  solar array panels with Si,
GaAs or multi-junction PV cells
 The PV cell coverglass  coated with
indium tin oxide (ITO)  to eliminate
arcing due to electrostatic discharge
https://www.pinterest.com
 Batteries  NiH2 cells assembled in
two batteries
 Electro-explosive devices for
deployment  powered directly
from the battery
Illustration of the SBIRS GEO spacecraft. Credit: Lockheed
Martin
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Spacecraft Power Systems
Small Satellite Bus
ASI-150EP/CP
 Typical power system features for small
satellites:
 Solar array covering  wide range of
solar flux and temperatures with wide
swings in I-U characteristics
 PPT architecture
 Body mounted solar array or 3-4 flat
panels, PV cells  silicon or GaAs
 One - two batteries  NiCd or NiH2
http://www.go-asi.com
www.qinetiq.com
 Power control unit  maximizes the
energy delivery to the bus  by driving
solar array to max. power point
 Tracker system does not dissipate
power assoc. with shunt regulation
inside the spacecraft
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Spacecraft Power Systems
Micro-Satellite Bus
 Very small satellites  load power
requirements in wats  simple and lean
architecture
 Solar array, battery, and loads 
permanently connected in parallel
 Battery  feeds to the loads automatically
during an eclipse, recharges during sunlight
 Fully charged battery voltage  relatively
constant, works as buffer
 Excess current absorbed by shunt resistance
connected in parallel with battery
http://librecube.net/
https://directory.eoportal.org
 For loads in tens of W  considered 7-15 V,
however  high costs because requires
components other than standard 28 V class
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Power System Sources
Power sources and systems of satellites and deep space probes
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Solar Array Systems
Intro
 The solar array made of numerous photovoltaic
(PV) cells connected in a series  parallel
combination to obtain required U and I
www.jpl.nasa.gov
 A photovoltaic cell  converts sunlight into direct
current electricity based on the PV effect
 The PV effect  is the electric potential
developed between two dissimilar materials when
their common junction is illuminated with photons
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Solar Array Systems
PV cell
 Energy of absorbed photons  transferred to
electron system of the material  creation of
charge carriers separated at the junction 
potential gradient (voltage)
 Origin of the PV potential  difference in the
chemical potential – Fermi level – of the
electrons in two isolated materials
 Absorption of photon energy in the cell 
knocking out an electron from atom, leaving a
positively charged hole  free electrons return
through external circuit under the potential
difference
http://alihelper.ru
 Voltage and incident energy fraction converted
into electricity  depends on the junction
material characteristics + radiation wavelength
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Solar Array Systems
PV Technologies
The most important measure of the PV cell performance  energy conversion efficiency
 cost per watt capacity

indicate economic competitiveness
of the PV with alternative power
generation technologies
The conversion efficiency of the PV cell 
𝜂=
𝑒𝑙𝑒𝑐𝑡𝑟𝑖𝑐𝑎𝑙 𝑝𝑜𝑤𝑒𝑟 𝑜𝑢𝑡𝑝𝑢𝑡
𝑠𝑜𝑙𝑎𝑟 𝑝ℎ𝑜𝑡𝑜𝑛 𝑝𝑜𝑤𝑒𝑟 𝑖𝑚𝑝𝑖𝑛𝑔𝑖𝑛𝑔 𝑡ℎ𝑒 𝑐𝑒𝑙𝑙
Theoretical maximum  for three common materials 
16 %  Ge
24 %  Si
29 %  GaAs
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Solar Array Systems
PV Technologies
 The energy conversion is different from the energy absorption  not all energy absorbed
by the PV cell is converted into electricity

Very important  sunlight wavelength spectrum
 Certain semiconductors (Si cell)  cut-off wavelength, such as 1.1 m  photons with
longer  are unable to generate hole-electron pairs  they pass through the PN junction
and are absorbed at the cell bottom in the form of heat
 About 2/3 of the solar radiation energy lies
between 0.4-1.1 m
 A 3 eV photon of blue light generates  0.5 eV
electricity, the remaining 2.5 eV is absorbed as
heat  this heat must be dissipated back into
the space; otherwise  
 A high emissivity back surface coating helps to
dissipate the heat effectively  
http://rredc.nrel.gov
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Solar Array Systems
PV Technologies
Major types of PV technologies:
 Single-Crystal Silicon
• Widely available cell material  workhouse of the space industry
• Method of production  Czochralski process  seed crystal placed in liquid Si and
drawn at slow constant rate  solid single-crystal cylindrical ingot
• Manufacturing process  slow, energy intensive, high raw material costs
• To minimize waste  to grow crystals on ribbons and cut by laser
www.nature.com
 Gallium Arsenide (GaAs)
•  conversion efficiency than silicon cells, but expensive
• Relatively insensitive to wide temperature cycling
•  radiation resistant
• Production  growing GaAs film on Ge substrate
GaAs PV cell
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Solar Array Systems
 The major types of PV technologies (cont.):
PV Technologies
 Semicrystalline and Polycrystalline
• Production  molten silicon poured into rectangular crucible  under
controlled cooling rate form partial and/or multiple crystals
• Eliminates  squaring-up process and waste materials
• Manufacturing  relatively fast and inexpensive
•  conversion efficiency, but much  cost  net reduction in cost/watt
• At present  uneconomical for space apps
 Thin Film
• Thin film materials  copper-indium-gallium-diselenide (CIGS)
and cadmium-telluride (CdTe)  a few m or  in thickness
directly deposited on substrate material
• Much  material per square FC area   expensive per watt
of generated power
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Credit: NASA
48
Solar Array Systems
 The major types of PV technologies (cont.):
PV Technologies
 Amorphous
• Si vapor deposited in  m thick amorphous film on glass or stainless steel rolls
• Poor conversion efficiency   10% at present
• Manufacturing  inexpensive, cost/watt generated  significantly lower
• Amorphous silicon (a-Si) efficiency   half of crystalline silicon (c-Si)  not
serious candidate for space power now
 Multi-Junction
•  demands on spacecraft power systems  double and triple junction cells with
improved performance, radiation resistance, and  cost
• Recently  tandem multi-junction GaAs cell
• Efficiencies  up to 35%  double junction and 40%  triple junction spacequalified cells (theoretical limit of triple junction  50%)
• Today  GaInP/GaAs on Ge substrate multi-junction cells  increasingly used
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Solar Array Systems
PV Cell Efficiencies
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Solar Array Systems
 Electrical characteristic of PV cell  I-U curve
I-U and P-U characteristics
 Important parameters  open circuit voltage UOC,
and short circuit current ISC
 Short circuit current  measured at full illumination
 ISC (photocurrent IS)  max current the cell can deliver
 I-U characteristics  from test data at various
illuminations, T, and ionized radiation doses
 Maximal power production  at knee point
 Panels designed  to operate closed to knee point
 Left-hand side  cell like constant current source
generating voltage to match with load resistance
 Right-hand side  cell like a constant voltage source
with internal resistance
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Solar Array Systems
I-U and P-U characteristics
Silicon PV cell
performance data
8x8 cm, 8 mm thick,
14% efficiency in AM0
GaAs/Ge PV cell
performance data
4x4 cm, 3.6 mm thick,
19% efficiency in AM0
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Solar Array Systems
Power from Albedo
 Albedo 
 Ratio of sunlight reflected by a planet to received by it
 Earth’s albedo  can contribute to power generation  can be used if the array is
designed accordingly
 Potential increase in array current by  10-15 % and  5% in voltage
 Usually albedo effects ignored  for simplicity, mainly in GEO satellites
Planet’s Albedo
http://www.nature.com
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Solar Array Systems
Back-Surface Illumination
 PV cells sometimes mounted on  flexible cloth
rather than rigid panels
Credit: ESA
Hubble with its second set of ESA
designed solar blankets
ISS solar arrays
Credit: NASA
 Flexible cloth array under back side illumination 
transmit significant energy to front  conversion
into useful power by front cells
 ISS back-surface illumination tests:
Back-illuminated cells  produced 41% of the
front-illuminated power at essentially the same
voltage
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Solar Array Systems
 Traditional construction method  cells mounted on
rigid substrate from aluminum honeycomb with skin
Rigid Panels
 Skin material  aluminum, but fiber composite  graphite, NomexTM, KevlarTM also
successfully used for mass reduction
 Array protection from space environment  primarily by coverglass
 Antireflective coating  to minimize light reflection + enhance light absorption by cell 
more power
Mars Atmosphere and Volatile Evolution, or
MAVEN spacecraft
Photovoltaic module with rigidizing backplane
Patent: US 20070074755 A1
https://spaceflightnow.com
FV cells
Flexible top sheet
Credit: ESA
EVA (Ethylenvinylacetat)
Flexible back sheet
Skin
material
ESA
Honeycomb
material
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Solar Array Systems
 Applications  small satellites for science missions
Body Mounted Array
 Cells mounted directly on spacecraft body without using honeycomb substrate
 Saves substrate mass and gimbals
 Disadvantage  needs of 57% more cells for the same power
 Rotation of body mounted array  illuminated strings at rated voltage, dark strings at
zero voltage
 Remedy  use of bleed resistance across isolation diode  equalizes potential of sunny
side and dark side through bleed current
Credit: NASA
Credit: ESA
Rosetta mission
FASTSAT
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Solar Array Systems
Three or More Wings
 Wings  can be canted at different angle as required for powering
 Application  small science mission satellites with peak power tracking design
Artist's rendering of the Juno spacecraft
www.engadget.com
www.thetechjournal.com
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Solar Array Systems
Flexible Array
 Installing PV cells on Kevlar cloth – open weave or close knit – can reduce array mass
 Array folded when stowed and deployed like accordion panel or rolled out
 Used on  the first Hubble Space Telescope, EOS-AM, Olympus, and ISS
 Design  withstand thermal snap on exiting the eclipse  large T gradient builds up
across cloth thickness  bowing in array can occur  electrical power reduced, since
array is not completely flat, comes back in 30 min
MegaFlex™ solar array system for all space environments
The UltraFlex solar arrays powered 2008’s
Mars Phoenix Lander. Source: NASA
www.orbitalatk.com
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Solar Array Systems
Inflatable Array
 Specific type of flexible array
 In balloon shape  flexible thin-film PV sheet on flexible composite laminate structure
densely packed for launch, and deployed in space using inflation gas
 Significantly reduces mass + packed in small volume  reduces costs
3U CubeSat Concept
Lightweight Inflatable Solar Array (LISA)
137 W min power
Earth Observation Nano-Satellite
Array radius 0.65 m
Credit: NASA
80 W to Payload, 1-2 year life
Array weight 0.75 kg
Credit: NASA
Credit: NASA
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Solar Array Systems
Array Performance
 Major factors influencing SA electrical performance 
 Sun intensity
 Sun angle
 Operating temperature
 Sun Intensity
 I-U characteristic  shifts  at  sun intensity with
small reduction in voltage
 However  photoconversion efficiency 
insensitive to solar radiation in working range,
falling off rapidly only bellow ¼ sun intensity
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Solar Array Systems
 Cell output current 
Sun Angle
𝐼𝑆 = 𝐼0 ∙ 𝑐𝑜𝑠𝜃
Photocurrent with normal sun ( = 0)
Sun angle  angle between the SA plane
 Cosine law holds well for    0, 50
and the sun-pointing vector
 Beyond 50  electrical output deviates significantly from cosine value
 No power generation beyond 85, although math prediction gives 7.5% power
 Actual power-angle curve of PV cell  called Kelly cosine

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Solar Array Systems
Temperature Effect
 With  temperature  short circuit current , open circuit voltage 
  Increase in current is much less than  drop in voltage  net effect  P
𝑷 = 𝑼𝟎 𝑰𝒔𝒄 − (𝜷𝑰𝑺𝑪 − 𝜶𝑼𝟎 )∆𝑻
𝑈0
𝐼𝑆𝐶
𝛼, 𝛽
∆𝑇
- open circuit voltage at reference T
- short circuit current at reference T
- temperature coefficients
- temperature gradient
 Typical 2 x 4 cm single crystal silicon cell 
𝛼 = 250 A/C, 𝛽 = 2.25 mV/C 
𝑷 = 𝑼𝟎 𝑰𝒔𝒄 𝟏 − 𝟎. 𝟎𝟎𝟓∆𝑻
  for every C  in operating temperature  cell power output  by 0.5%
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Solar Array Systems
Temperature Effect
 Power vs voltage characteristic 
max. power  at  temperature  cold T better for PV cell  more power generation
 Pmax points  at different voltages  to extract Pmax at all T  PV system must be
designed to  array output voltage (U2) for capturing Pmax2 at  T and vice versa
 Peak Power Tracking architecture  needed, otherwise excess power must be wasted in
shunt circuits
 Array undergoes  wide temperature cycle in each orbit  during sunlight  50-60 on
the front face, 40-50 on the back face  with 10 gradient
 Solar array thermal equation  [Solar flux + Earth’s albedo + Earth’s thermal radiation +
heat from adjacent components of spacecraft] = [electrical power output + heat radiated
back into space]
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Solar Array Systems
Temperature Effect
 Incidence + radiated heat from spacecraft body  varies with location in the orbit
 During eclipse in GEO  temperature drops exponentially to  -175 C
 Front to back face temperature gradient for rigid array  5-10 C under steady sun,
up to 20 C on sunlight snap after eclipse
LEO
GEO
Heat input and output of satellite in
Earth orbit
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Solar Array Systems
Sun Acquisition
 Array wings after deployment  oriented to acquire the sun normal to array surface
 Sun sensor  two PV cells mounted on two 45 wedges + connected differentially in
series opposition through actuator motor
 The sun is perfectly normal  currents from both cells equal to  I0 ∙ cos45
 Net motor current is zero  array stays put
 Array not normal to the sun  error angle   the sun angles on two cells are different
 different currents  motor turns toward the sun until current is zero
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Solar Array Systems
Sun Tracking

Required for solar array  to face the sun continuously as spacecraft orbits the Earth

Done by actuator  follows the sun like sunflower

Two types of sun trackers:
 One-axis gimbals  follows the sun from east to west during day
 Two-axis gimbals  tracks the sun from east to west during day + from north to
south during seasons of the year
Credit: NASA
Beta Gimbal Assembly
(BGA or Beta Gimbal)
Solar Alpha Rotary Joint
(SARJ or Alpha Gimbal)
Close view of the ISS gimbals
ISS  and  gimbals
https://geektimes.ru
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Solar Array Systems
Peak Power Extraction
 Gimbal motor  drives module to face the sun for collecting maximum solar flux
 However  module must electrically operate at voltage corresponding the peak power
point
 Operating point  intersection of the source line and the load line
 Necessary condition for electrically stable operation of solar array 
Load matching of PV source with constant
resistance load
𝑑𝑃
𝑑𝑈
>
𝑙𝑜𝑎𝑑
𝑑𝑃
𝑑𝑈
𝑠𝑜𝑢𝑟𝑐𝑒
Load matching of PV source with constant
power load
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Solar Array Systems
Peak Power Extraction
 Operating at U and I on the I-U curve  power generation P = UI
 If operation moves from above point to  I + I, and U + U  new power:
𝑷 + ∆𝑷 = (𝑼 + ∆𝑼)(𝑰 + ∆𝑰)
∆𝑷 = ∆𝑼𝑰 + ∆𝑰𝑼
Dynamic impedance of the source Static impedance
 P should be zero at peak power point 
𝒅𝑼
𝑼
=−
𝒅𝑰
𝑰
 PP extraction methods 
 Impedance method
 Power gradient method
 Voltage ratio method
 Current regression method
 Power regression method
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Solar Array Systems
 Array  many parallel strings of series connected cells
Shadow Effect
 Large array  may be partially shadowed (due to structure interfering with sun line)
 Completely shadowed cell in long string  loses photovoltage but still must carry current
(series connection)  without internally generated voltage no power production  acts as
load producing local I2R loss + heat  power loss
 No problem  mild shadow on small area
 However  more cells shadowed 
I-U curve reduced below operating voltage 
string current fall to zero  losing all power
www.solartec.eu
Shadow effect on long SA PV string
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Solar Array Systems
Shadow Effect
 Common method of eliminating the loss  to subdivide string length into segments with
bypass diodes
 Diode across shadowed segment  bypasses only that string segment
 Causes  proportionate loss of string U and I without losing whole string P
Bypass diodes in PV strings to minimize
power loss under heavy shadow
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Battery Systems
Intro
 Main purpose  the energy storage  to meet spacecraft load demand during eclipse +
when demand exceed power generation at any time
 Battery  stores energy in electrochemical form
 Basic types:
 Primary battery (PB)  irreversible electrochemical reaction (no reuse after full
discharge)  very short duration missions as primary power source.
Specific energy  an order of magnitude  compared to secondary battery
 Secondary battery (SB)  reversible electrochemical reaction  can be recharged 
long duration missions (independent source required for periodical recharging), most
satellites use SB
Space Vector Corporation:
Lithium-ion Battery
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Battery Systems
Primary Battery
 Long storage capability
 Dry  without electrolyte until needed
 Activated  by introducing electrolyte into dry battery
 Electrolyte may be solid at room temperature  activated heater to melt electrolyte
(Thermal battery)
 Typically  fairly large energy density, and long shelf life
 Lithium primary batteries 
Li-SO2, Li-SOCl2, Li-BCX, Li-(CF)x, Li-MnO2  used by planetary probes, rovers, and
astronaut equipment
Li-(CF)x  has been used in launch vehicles
Li-SO2  used in long duration exposure facility
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Battery Systems
Primary Battery
 Examples 
 ESA’s Huygens probe  relied on non-rechargeable Li-SO2
primary batteries  could be left inactive during 7-year
trip to Saturn  but still retain sufficient capacity for
landing on Titan
 Foton M3 missions  Li-SOCl2 primary batteries 
allowed performing experiments for more then 10 days
before reentering the atmosphere
 Ariane 5  power its autonomous operations during brief
but crucial flight to orbit  inertial navigation + guidance,
engine firings and booster separation
 Used also  for early major mission events, short duration
missions
Credit: NASA
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Battery Systems
Electrochemical Cell
 Battery  numerous electrochemical cells in series-parallel combination to obtain 
required U and I
 Cell voltage  depends solely on electrochemistry, not physical size
 Common electrochemistries  produces 1.5-3.5 V when fully charged
 Cell capacity (C)  depends on physical size  Ah charge the cell deliver at room T until
reaches cut-off voltage  2/3rd of fully charged voltage
 Battery voltage rating  average voltage during discharge
 Product of voltage and Ah rating  energy rating in Wh the battery can deliver from fully
charged state
 Battery charge and discharge rates (C-rate)  stated in fraction of the capacity
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Battery Systems
Electrochemical Cell
 State of charge SOC 
Ah capacity remaining in the battery
SOC =
rated Ah capacity
 SOC affects  cell voltage, specific gravity, and freezing point of electrolyte
 Fully charged battery   specific gravity and  freezing point of electrolyte
 Fully discharged battery   freezing point of electrolyte  important to keep
battery fully charged when exposed to low temperatures
 Depth of discharge DOD 
DOD =
Ah capacity drained from fully charged battery
rated Ah capacity
DOD = 1 − SOC
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Battery Systems
Electrochemical Cell
 Major secondary batteries used at present  nickel-cadmium (NiCd), nickel-hydrogen
(NiH2), and lithium-ion (Li-ion)
 New electrochemistry  continuously researched for space apps  for example: lithiumpolymer (Li-poly) and nickel-metal-hydride (NiMH)
 Selection of electrochemistry  matter of performance + cost optimization
 Figure of merit used in comparing relative performance of various electrochemistries:
 Specific energy  gravimetric energy density  energy stored per unit mass, Wh/kg
 Energy density  volumetric energy density  energy stored per unit volume, Wh/l
 Specific power and power density  power the battery can deliver per kg of mass,
resp. liter of volume
 Cycle life  number of charge/discharge (C/D) cycles the battery can deliver while
maintaining the cut-off voltage
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Battery Systems
Battery types
 NiCd  workhouse of spacecraft industry since the earliest missions, still in use in some
missions
 NiH2  battery for general use since the mid-1980s  provides deeper DOD for
comparable cycle life  requiring lower Ah capacity  lighter spacecraft weight
 At present  moving towards lithium based batteries for potentially 2-5 x the specific
energy compared to NiH2 chemistry
Ni-Cd battery pack used
on SPOT-4 satellite.
Previous Ni-H2 ISS battery ORU
(Orbital Replacement Unit)
©Saft
©2007 Space Systems Loral
Saft Li-ion 28V space
grade battery
©Saft
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Battery Systems
Source: The Electrochemical Society Interface • Fall 1999
Battery types
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Battery Systems
Battery Types
Battery Type
Energy Density (Wh)/kg
Silver–zinc (Ag-Zn)……………………………………………………………………………………………………………………………….... 120 –130
Silver-cadmium (Ag-Cd)…………………………………………………………………………………………………………….. 60 –70
Nickel-cadmium (Ni-Cd)…………………………………………………………………………………………………………….. 20 –30
Nickel-hydrogen (Ni-H2)……………………………………………………………………………………………………………… 60 –70
Nickel-metal hydride (Ni-MH)………………………………………………………………………………………. 120 –130
Lithium Thionyl Chloride (Li-SOCl2)…………………………………………………………………… 650
Lithium Vanadium Pentoxide (Li-V2O5)………………………………………………….... 250
Lithium Sulfur Dioxide (Li-SO2)…………………………………………………………………………………... 50 –80
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Battery Systems
 Relatively  specific energy, T sensitive,
shorter life cycle + cadmium placed under
environmental regulatory scrutiny
Nickel Cadmium
 Major disadvantage  memory effect 
remembers its DOD, does not work well
beyond that  partial loss of unused
capacity for subsequent use
 Reconditioning  remedy to counter memory
Memory effect in NiCd cell
effect for restoring the battery to full capacity
 Battery  fully discharged to almost zero voltage, then fully charged to  1.55 V/cell
 In spacecraft power system  done twice/year at convenient time in the orbit
 In GEO  done outside the two eclipse seasons
 In LEO  spacecraft must carry an extra battery just for
reconditioning purposes (adding significant mass + cost)
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Battery Systems
Advanced Nickel Cadmium
 Advanced NiCd  improvements in terms of life + specific energy
 Competes with NiH2 for some low power spacecraft, particularly in defense apps
 Negligible memory effect
 Several-fold  cycle life than conventional NiCd and a few-hold  than NiH2
 Used typically in small LEO satellites  capacities ranging from 5-50 Ah
 Better performance than the old NiCd only bellow 50 Ah,
matches with NiH2 at  capacities
 Primary limitation for capacity restriction  heat removal difficulty across the plates
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Battery Systems
www.nasa.gov
Nickel Hydrogen Battery
 Nickel Hydrogen:
 The most widely used  during the last 20
years in LEO + GEO satellites
Nickel-Hydrogen batteries for Hubble
 Takes the best from two electrochemistry 
nickel oxide electrode from NiCd + hydrogen
catalyst electrode from Fuel Cell
www.nasa.gov
 Withstanding some abuse in over-charging
and over-discharging
 Disadvantages  low energy density;
pressure vessel rupture, handling, and safety
consideration; high self-discharge rate
 0.5%; high loss of capacity on storage
Hubble's new batteries integrated onto the Super Light-weight Interchangeable Carrier
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Battery Systems
Nickel Hydrogen Battery
 Design  each elementary cell developing 1.25 V
packed in individual pressure vessel
 Electrodes  (+) sintered nickel, (-) platinum
 Separator  on each side of the (+)
 Electrolyte  26 or 31% KOH
NiH2 CPV
positioned
Vertically
 Pressure vessels  individual (IPV), common (CPV),
single (SPV), dependent (DPV)
 Cell capacity  function of loading level of the
positive active material, KOH concentration, and T
http://en.academic.ru
 Compared to NiCd, the NiH2 chemistry  superior charge/discharge cycle life + low
internal resistance
 No noticeable memory effect  reconditioning not needed, however advised  in all
batteries with (+) nickel electrode
 Capacities of commonly available NiH2  2-400 Ah (IPV), 60-200 Ah (CPV), 15-50 Ah (SPV)
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Battery Systems
Nickel Metal Hydride Battery
Nickel Metal Hydride 
 Anode  metal hydride eliminating environmental concerns of cadmium
 Improvement in  specific energy over that in NiCd
 Negligible memory effect
 Disadvantages  less capable of delivering high peak power,  self-discharge rate,
sensitive to  T risk of damage due to over-charging, expensive compared to NiCd
 Electrochemically  similar to NiH2 except  hydrogen stored as solid metallic hydride
rather than gas  no internal pressure
 Significantly improved  energy density and safety,  volume to pack,  support structure
 Cells  prismatic + the same operating voltage as NiCd  easily interchangeable
 Produces heat during charging  opposed to endothermic charging of NiCd
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Battery Systems
Lithium-Ion Battery
 New development  significantly  specific energy + energy density over NiH2
 Prismatic shape  higher energy density
 Design  electrodes: (+) LiNiCoO2, (-) mixture of two graphites, electrolyte: LiPF6 salt,
stainless steel prismatic case
 End-of-discharge cut-off voltage of 2.7 V, average discharge voltages of 3.5 V,
end-of-charge voltage of 4.2 V
 Average discharge voltages of 3.5 V vs 1.25 V for NiCd and NiH2  require  1/3rd the
number of cells in series for given battery voltage  reducing mass + costs
Photo: Saft's qualified high-energy space
Li-ion batteries for GEO telecom satellites
Installation of new
Li-ion batteries on the ISS
SAFT Li-ion
batteries used
in satellites
Saft
Saft
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85
Battery Systems
The ISS Lithium-Ion Batteries
2 Power Channels/Integrated Equipment Assemblies (IEAs)
6 Ni-H2 ORUs/channel  48 total
1 Li-ion + 1 Adapter Plate to replace 2 Ni-H2  24 total Li-ion
GS Yuasa 134 Ah cells
Credit: NASA
ORU Layout – three Cell
“10-Packs”
ESA AOES Media lab.
ORS: Orbital Replacement Unit, BCDU: Battery Charge/Discharge Unit
Saft DC Switching Unit
DCSU:
Heat Barrier (12/ORU)
• High margin against
thermal runaway
propagation
• Reflects 787
reach-back safety
addictions
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Battery Systems
 Electrochemistry  high charge efficiency
Lithium-Ion Battery
 High rate capability + potentially long cycle life at
high DOD
 Low temperature sensitivity  operation over wide
temperature range
 Low internal impedance  high charge and
discharge rates
 Delivering  high short time peak power without
adverse effect on life
 Potential of achieving  150 Wh/kg and 400 Wh/l
with abuse tolerance
Credit: US Department of Energy
 At present   cost, but cost competitive with NiH2 in the same volume production
 Vulnerable to damage  under over-charging or other shortcomings 
requires more elaborate charging circuity with protection against over-charging
 Does not tolerate  over-charge or over-discharge
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Battery Systems
Lithium Polymer and Silver Zinc
Battery
 Lithium Polymer 
 Electrochemistry  solid polymer electrolyte (Li-poly) that encapsulates electrodes
 Solid polymer  enhances specific energy by acting as both  electrolyte + separator
 Improved safety  reduced flammability compared to Li-ion
 Silver Zinc (AgO-Zn) 
 High specific energy but shorter cycle life
 Suitable in missions requiring  low number of C/D cycles
 Specific energy ranges  125-250 Wh/kg, energy density  200-600 Wh/l
 Prismatic cells available in wide range of Ah in low-rate (LR) and high-rate (HR) types
 Resistance of AgO electrodes vary little with state of charge  stable operating
voltage provided until nearly all capacity is dawn
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Battery Systems
Properties and Performance
 Electrical performance  depends on electrochemistry + many other parameters in highly
nonlinear manner
 Battery design  one of the most difficult tasks for power system engineers
 Charge/discharge characteristics  depends on charge rate, discharge rate, temperature,
age of the cell
 Typical average NiH2 and NiCd battery
cell voltages:
 1.55 V  when fully charged
 1.45 V  average during charge
 1.25 V  average during discharge
Internal electrochemical voltage
Internal battery resistance
 1.10 V  at 80 % DOD
 1.00 V  when fully discharged
Battery line vs load line intersecting at operating point
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Battery Systems
Properties and Performance
 Charge/discharge ratio  defined as Ahs input over
Ahs output for full state of charge  depends on
charge and discharge rates + temperature
 Charging in LEO  usually at much faster rate than
in GEO
 LEO C/D = 1.05 -1.10, GEO C/D = 1.1-1.2
 High C/D ratio  cause excessive corrosion of Ni
electrodes + large amount of oxygen leading to
heating and popping during recombination on
hydrogen electrode
NiH2 C/D ratio vs temperature
 Low C/D ratio  causes reduced usable capacity,
low discharge voltage + capacity degradation
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Battery Systems
Properties and Performance
Cell Ah capacity vs
temperature
 New cell stores  over 20 %  capacity when
charged at -20 C compared to +20 C 
 But only  10 %  after  100 cycles of C/D cycles
(1 year in GEO)
 Max charge voltage  varies with temperature
Full charge voltage
at C/10 for 16h
 Better battery performance  under slow charge
and discharge rates
 Fast discharge rate  fast voltage degradation +
 Ah capacity  high charge and discharge rate
apps require different design considerations
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Battery Systems
Properties and Performance
 Fully charged battery  any additional charge converted into heat
 Any excessive over-charging  excessive gassing scrubbing the electrode plates
 Continuous scrubbing  excessive heat, wears out electrodes and shortens the life 
need for regulator to cut back to trickle charge rate once battery is fully charged
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Power Systems
Power System Design Trades
Flow-Chart
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Fuel Cell Power
Intro
 Fuel cell (FC) or “gas battery”
intermediate-term source for space applications
 Principle  direct conversion of
fuel’s chemical energy into  DC electricity
 Operation  producing electricity as long
as the fuel is supplied  has to be recharged
 Typical fuel  hydrogen or hydrogen-rich
mixture + oxidant
 Working of FC  reversed process to water
electrolysis  hydrogen + oxygen combined to
produce  electricity + water
 Isothermal process  conversion efficiency not
limited by Carnot efficiency  converts  % of
fuel’s chemical energy directly to electrical energy
The Apollo 13
Fuel cell located
immediately
above the H2 and
O2 cryogenic
tanks
Image Source: NASA
 Fuel cell efficiency   twice of thermodynamic
converter  65-80%
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Fuel Cell Power
Intro
 Previous use  moon buggy (the first), Gemini, Apollo, Space Shuttle (STS Orbiter)
 At present  used routinely + new developments for planetary rovers
 Applicability 
 space missions  few days - few weeks when battery is not practical
 auxiliary power source for orbit transfer vehicles
 as regenerative fuel cell  attractive mass saving for LEO satellites
 (it was a serious candidate in place the battery for the ISS)
Apollo FC, 27-31 V
563-1420 W
max. 2300 W
Gemini 1 kW
Fuel Cell
Credit: NASA
Space Shuttle Fuel Cell
12-15 kW
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Fuel Cell Power
Electrical Performance
 Fuel cell  works as voltage source with internal resistance
 Theoretical potential difference in hydrogen-oxygen fuel cell  1.25 V
 Potential different  reaction dependent
 Multiple fuel cell  series-parallel combinations using heavy graphite pallets
 Voltage drops significantly  due to various losses
Ideal cell voltage
 Primary loss mechanism  ohmic loss in electrodes
𝑈𝑑𝑟𝑜𝑝 =∝ +𝛽 ∙ 𝑙𝑛𝐽
Total Loss
Constants dependent on
T and electrode surface
Current density at
electrode surface
 Electrical performance  given by electrode voltage
vs surface current density  known as polarization
curve or U-I curve
Pe = UI
Fuel cell polarization curve –
voltage vs current
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Fuel Cell Power
 Ideally  H2-O2 fuel cell  produce 1.25 V DC
Electrical Performance
 Undesirable ions + products of intermediate irreversible reactions   cell potential
 Voltage drop under load 
 Resistance polarization  electrical resistance of the electrolyte + electrodes
 Concentration polarization  accumulation of ions + reaction products + depletion of
ions and reactants in the electrolyte near electrode surface  energy losses associated
with mass transport
 Activation polarization  reluctance of fuel and oxidant to undergo reaction at each
electrode  losses associated with reactions
 All ohmic losses within cell  electrodes, current
collectors, contacts, ionic impedance of the electrolyte
 Polarization result  cell produces 0.5–1.0 V DC at 100–400 mA/cm2 of cell area
 Increasing of performance  by  cell temperature + reactant pressure
 Practical operating range  controlled by ohmic losses
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Fuel Cell Power
Electrical Performance
 Terminal U-I relationship:
(with  t  k  and U0 )
𝑈 = 𝑈0 − 𝑘𝐼
Open circuit voltage
 Power at any operating point:
 The maximum power  when
Constant
𝑃 = 𝑈𝐼 = (𝑈0 − 𝑘𝐼) ∙
𝑑𝑃
𝑑𝐼
(𝑈0 − 𝑈)
𝑘
= 0  at 𝑉 = 1/2 ∙ 𝑈0 
𝑃𝑚𝑎𝑥
1 𝑈02
=
4 𝑘
 FC uses on-board fuel  not operated at Pmax  operated at maximum fuel
efficiency until the EOL
 Open circuit voltage as a function of time:
𝑈0 (𝑡) = 𝑈0 (0) − 𝐾0 ℎ
Number of hours since FC
is put in operation
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Fuel Cell Power
Electrical Performance
PEMFC
FC operating characteristic degradation with time
© 2014 by ASME
Pmax/Prated ratio vs time - determines the life
 Fuel cell life  time needed for voltage
to fall below required input voltage, resp.
maximum power below required output
power
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Fuel Cell Power
Source: Argonne National Laboratory
Electrical Performance
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Fuel Cell Power
Fuel Cell Types
 Fuel cell classified by  types of fuel and electrode materials
 Solid polymer electrolyte (SPE)  in the early manned missions lasting < 1 week
 Proton exchange membrane (PEM)  Gemini
 Alkaline (AFC)  missions to the moon
 Improved aqueous alkaline technology NASA’s space shuttle fleet
 Lightweight hydrogen-oxygen alkaline fuel cell stack  uses propellant H2 + O2 to
generate peak power of 3.5 kW for average load of 2.2 kW
Alkaline Fuel Cell
Power for Space Shuttle
Lightweight Fuel Cell
Power for Aerial Drones
The Lynntech
Flightweight PEM
Fuel Cell Stack
https://spinoff.nasa.gov
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Fuel Cell Power
Fuel Cell Types
Space shuttle fleet:
 No batteries
 High power FC stack  operating on cryogenic
reactants H2 and O2  peak power of 12 kW,
brief overloads of 15 kW, average load to 4.5 kW
(increased on 7 kW)
 Before launch  starting power from the ground 
then fuel cells
The Space Shuttle Discovery Fuel Cell
 Each shuttle  3 FC generating 30-36 V DC,
also conversion to 400 Hz AC
 Output power capability  12-15 kW 
specific power 100–120 W/kg
 Electrical startup time 10-20 min, shutdown is
instantaneous, the whole system  requires  2 h
warm-up time before loading, design life = 5000 h
The Space Shuttle Endeavour Fuel Cell
Source: NASA
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Fuel Cell Power
Alkaline Fuel Cell
 Alkaline fuel cell (used in the space shuttle) 
combines H2 and O2 to produce  electricity + water
as a byproduct
 Principle 
 O2 enters the cell through sintered-nickel cathode
 catalyst produces OH- anions delivering them
to KOH (potassium hydroxide ) electrolyte
Image: www.eere.energy.gov
𝐀𝐧𝐨𝐝𝐞 𝐑𝐞𝐚𝐜𝐭𝐢𝐨𝐧 𝐨𝐱𝐢𝐝𝐚𝐭𝐢𝐨𝐧
2𝐻2 + 4𝑂𝐻 − → 4𝐻2 O + 4𝑒 −
𝐂𝐚𝐭𝐡𝐨𝐝𝐞 𝐑𝐞𝐚𝐜𝐭𝐢𝐨𝐧 𝐫𝐞𝐝𝐮𝐜𝐭𝐢𝐨𝐧
𝑂2 + 2𝐻2 O + 4𝑒 − → 4𝑂𝐻 −
𝐍𝐞𝐭 𝐑𝐞𝐚𝐜𝐭𝐢𝐨𝐧 𝐫𝐞𝐝𝐨𝐱
2𝐻2 + 𝑂2 → 2𝐻2 O
 Anions drift through the alkaline electrolyte to
anode  react with delivered H2 to form  water
molecules + release electrons  electrical power
delivered to external load
 Energy released by 1 kg of H2 + 8 kg of O2 
 34 kWh, conversion efficiency of  60-70 %
 Byproduct water  used for the crew
 Operational temperature  70 – 100 C
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Fuel Cell Power
Proton Exchange Membrane
Fuel Cell
 Proton exchange membrane (PEM) technology 
more suited for spacecraft applications at present
 PEM offers  enhanced safety, longer life,  weight,
 reliability,  peak power capability, compatibility
with propulsion fuel,  cost
 Compared to PV array   specific power +
flexibility  no need for the sun pointing
Image: www.eere.energy.gov
 Disadvantage  needs to carry fuel on board
 Operational Temp.  60-80 C, Efficiency  50-60%
𝐀𝐧𝐨𝐝𝐞 𝐑𝐞𝐚𝐜𝐭𝐢𝐨𝐧 𝐜𝐚𝐭𝐚𝐥𝐲𝐭𝐢𝐜𝐚𝐥 𝐬𝐩𝐥𝐢𝐭𝐢𝐧𝐠
2𝐻2 → 4𝐻 + + 4𝑒 −
𝐂𝐚𝐭𝐡𝐨𝐝𝐞 𝐑𝐞𝐚𝐜𝐭𝐢𝐨𝐧 𝐨𝐱𝐲𝐠𝐞𝐧 𝐫𝐞𝐝𝐮𝐜𝐭𝐢𝐨𝐧
𝑂2 + 4𝐻 + + 4𝑒 − → 2𝐻2 O
PEM Fuel Cell
Source: NASA
𝐎𝐯𝐞𝐫𝐚𝐥𝐥 𝐑𝐞𝐚𝐜𝐭𝐢𝐨𝐧 𝐫𝐞𝐯𝐞𝐫𝐳𝐢𝐛𝐥𝐞 𝐫𝐞𝐚𝐜𝐭𝐢𝐨𝐧
2𝐻2 + 𝑂2 → 2𝐻2 O
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Fuel Cell Power
 Presented non-regenerative FCs  convert energy one way:
fuel  electricity
Regenerative Fuel Cells
 Recharging FCs  requires electrolyzer to decompose water
back into hydrogen + oxygen
 Electrolyzer  generally a separate unit 
the two cannot operate simultaneously
 Hydrogen-oxygen RFCs  developed for energy storage in
short-term space missions
The Unitized
Regenerative
Fuel Cell
 RFC  considered for the ISS but dropped in favor of NiH2
battery  due to poor round trip energy efficiency
https://str.llnl.gov
 However  high-power long-term manned missions to the
moon or Mars  find RFC practical
 RFC  based on hydrogen-oxygen PEM FC + electrolyzer
technology
  peak power capability   10 x the base power 
useful for power pulses
http://www.sciencedirect.com
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Fuel Cell Power
 Disadvantage  peripheral pumps and
plumbing for handling fluids reliably over a
long mission life in GEO
Regenerative Fuel Cells
 Providing redundancy in such system 
may incur significant mass penalty
 Possible RFC system for LEO  mission of
5 years, specific power of 5-8 Wh/kg,
round trip energy efficiency of 60-70 %
RFC System at NASA Glenn Research Centre
 High power GEO apps  mass-optimized 
20-35 Wh/kg  large due to the need for
small electrolyzer unit  relatively long
recharge time available in GEO
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Fuel Cell Power
Regenerative Fuel Cells for
Satellites
 RFC  may potentially replace batteries in satellites (not in near future)  because of
 round trip energy efficiency than battery + radiators and heat pipes in RFC  in terms
of mass and integration into spacecraft design
 However in future designs  RFC may find apps in high-power spacecraft  for example
in 10 MW pulse power directed energy weapon platforms in LEO
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Interplanetary and Deep Space Missions
 Power systems for interplanetary and deep space
missions  significantly different environments
Intro
 Closer to the sun or farther away  extreme
temperature  high or low  large impact on the
performance
 Long flight times  need for power systems with
 10 years life
http://voyager.jpl.nasa.gov
 Mass is at a absolute premium  need for power
systems with  specific power and scalability
 Some missions require ion propulsion to reduce
flight time  need for  power capabilities with
 specific power and  cost
http://www.businessinsider.com
 Over 3 order of magnitude decrease in solar flux
from Earth to Pluto
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Interplanetary and Deep Space Missions
Solar Irradiance in Deep Space
40
1
 10
15
30
5
50
2
2
610
20
Ref
1358
4
<1
2200
Distance (AU)
Solar flux (W/m2)
 Solar flux at any distance
in deep space 
𝐼𝐸𝑎𝑟𝑡ℎ
𝐼=
𝑅2
solar flux in the Earth orbit (1358 W/m2)
distance from the sun in AU
Patel, M. R.: Spacecraft Power Systems.
Boca Raton: CRC Press, 2005.
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Interplanetary and Deep Space Missions
Temperature in Deep Space
 Surface temperature is determined by:
𝑇=
𝐼
𝜎
𝛼
𝜀
1
4
𝜎 Stefan-Boltzmann constant
  absorptivity
  emissivity of the surface
http://www.azimuthproject.org
 Example  spacecraft in deep space at 2 AU distance from the sun receives solar flux ¼
compared to Earth. Solar array with / = 0.8 would give rise to a temperature of:
𝑇=
1375/4
5.67∙10−8
0.8
2
1
4
= 221K  -52 ℃
 Missions beyond Mars  cannot effectively generate PV power  very  solar intensity
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Interplanetary and Deep Space Missions
Power System Options
 Interplanetary + deep space missions  not effective PV power generation due to
insufficient solar flux
 Spacecraft must carry on board a primary energy source  radioactive isotope or nuclear
reactor
 Radioisotope  heats a thermoelectric (TE) material such as lead telluride  generates
electrical potential based on the Seebeck effect
 Radioisotope thermoelectric generator (RTG) or “nuclear battery”   several hundred W
 Nuclear reactor  considered for high power in 30 – 300 kW range
The Russian TOPAZ Nuclear Power System
150 kWth/5 kWe, used in US-A reconnaissance satellites
Source: Ponomarev-Stepnoi et al. [2000]
Radioisotope Thermoelectric Generator
 100 We. Credit: NASA
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Interplanetary and Deep Space Missions
Power System Options
 Advantage 
 Provides power for a long period of time  eliminating the need for a battery
 Independent of the spacecraft orientation and distance from the sun
 Suitable for missions  far away from the sun, too close to the sun, lunar missions
with long eclipse periods
 Disadvantage  heavy radiation shielding required around electronic components;
expensive nuclear fuels (238Pu, 244Cu),
inexpensive and easily available 90Sr is unsafe
 Radiated thermal power 
𝑃 𝑡 = 𝑃0 𝑒
0.7𝑡
−𝑇
1
2
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Radioisotope Thermoelectric Generator
Radioisotope 238Pu
 Half-life  87.74 a, power density  0.56 W/g,
kinetic energy of alpha particle  5.544 MeV
INL: Advanced Test Reactor
 Exceptionally  gamma and neutron radiation levels
 The lowest shielding requirements  needs less than 2.5 mm Pb
shielding, in many cases  no shielding is needed, as the casing
itself is adequate
 The most widely used isotope for RTGs, but  relatively 
stockpile (USA restarted production at  1.5 kg/year in 2013)
INL
1
0𝑛
237
1
+ 238
92𝑈 → 2 0𝑛 + 92𝑈
237
92𝑈
Production of 238Pu
→
0
−1𝑒
1
0𝑛
+ 237
93𝑁𝑝
+ 237
93𝑁𝑝 →
Alpha decay of 238Pu
238
94𝑃𝑢
→
234
92𝑈
A glowing 238Pu RTG element
+ 42𝛼(5.544 𝑀𝑒𝑉)
238
93𝑁𝑝
+ 00𝛾
238
93𝑁𝑝
→
0
−1𝑒
+ 238
94𝑃𝑢
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Radioisotope Thermoelectric Generator
RTG Radioisotopes
Source: U.S. Atomic Energy Commission
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Radioisotope Thermoelectric Generator
RTG Conversion Efficiency
 RTG consists of numerous thermoelectric converter (TEC) cells  to obtain required U, I
electrical power output
 Power conversion efficiency defined as: 𝜂 =
thermal power depletion
 Efficiency depends on  material properties + hot and cold junction temperatures (Th, Tc):
Material 1
𝜂=
𝑇ℎ − 𝑇𝑐
𝑇ℎ
Tc
Material 2
Th
Heating
U
 The most used 238Pu + SiGe TE cells gives   7% conversion efficiency
 Removing remaining 93 % of system energy as waste heat  significant design challenge
 Specific electrical power output  typically low  it was 2 W/kg in SNAP-19, and 5 W/kg
in Galileo (SNAP: Systems Nuclear Auxiliary Power – experimental RTG program)

238Pu
costs more than a couple of million $/kg
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Radioisotope Thermoelectric Generator
Principle
 Working principle  based on Seebeck effect 
generates electrical potential when two dissimilar
materials are maintained at different temperatures
Single-stage unicouple TEC
 Involves electron or hole transfer between two
dissimilar materials under thermal energy
 Two materials  conductors or semiconductors
 TEC for space power apps  uses semiconducting
materials  p-type and n-type
 At temperature difference T12  electrical potential
difference U12 is produced at junction:
𝑼𝟏𝟐 = 𝜶𝟏𝟐 ∙ ∆𝑻𝟏𝟐
Source: Idaho National Laboratory, USA
Differential Seebeck Coefficient of the couple (V/C)
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Radioisotope Thermoelectric Generator
Principle
 Each arm of two materials  has electrical resistivity  + thermal conductivity k
 Parameters: , k, 𝛼  vary with operating temperature
𝛼2
 Figure of merit for each material in the couple  𝑍 =
𝜌𝑘
 Figure of merit of both materials together in the couple  𝑍12 =
𝛼 212
𝜌1 𝑘1 + 𝜌2 𝑘2
2
 Good TE junction has:
  Seebeck coefficient   generates higher voltage
  electrical resistivity 𝝆  results in low ohmic loss
  thermal conductivity k  gives high T
 Electrical power 
𝑇𝑐
𝑃𝐸 = 1 −
𝑇ℎ
1 + 𝑍𝑇 − 1
𝑇
1 + 𝑍𝑇 + 𝑇𝑐
ℎ
𝑇=
𝑇ℎ + 𝑇𝑐
 Some semiconductors   Seebeck coefficient in the range: 300 – 500 V/C
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Radioisotope Thermoelectric Generator
Principle
 Typical space-qualified RTG uses 
 High-grade SiGe thermoelectric converters + 238Pu heath source
 SiGe  doped with phosphorus for n-leg and boron for p-leg
 Z12  0.001  30 % depending on material grade and quality of manufacture
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Radioisotope Thermoelectric Generator
Single-Stage Unicouple
 Single-Stage Unicouple
 The most basic elementary construction of TEC
 One end of: p/n arms  kept in common
reservoir at Th
 The other ends  kept at Tc
 DC voltage  proportional to T  generated
between p and n terminals at the cold end
 Conversion efficiency   7%
 Cold T reservoir  liquid metal bath  p/n
terminals must be kept at two different
reservoirs electrically isolated to withstand the
generated voltage
MHW/GPHS SiGe Unicouple
(used in Galileo and Cassini)
Source: http://thermoelectrics.matsci.northwestern.edu
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Radioisotope Thermoelectric Generator
Single-Stage Multicouple TEC
 Single-Stage Multicouple Thermoelectric Converter:
 Two or more p/n couples  stacked together
 Couples  thermally in parallel  the same T
electrically in series  higher output U
Source: Patel, M. R.: Spacecraft Power
Systems. Boca Raton: CRC Press, 2005.
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Radioisotope Thermoelectric Generator
Multistage Multicouple TEC
 Multistage Multicouple Thermoelectric Converter:
Electrical insulation
 Heat extracted in two stages
Cu strap
 Temperature drops from Th  Tm
in the 1st stage having two couples
Cu strap
 From Tm  Tc in the 2nd stage
having two couples
Electrical insulation
Cu strap
 Electrical isolation  high thermal conductivity
to minimize temperature drop across it
 Copper strap  high thermal and electrical
conductivities  conducts electrical current +
distributes the heat uniformly across the device
 all couple in the same T
El. insulation
El. insulation
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Cu strap
121
Radioisotope Thermoelectric Generator
RTG Assembly:
 Seebeck voltage per junction   300 – 500 V/C
RTG Assembly
 numerous junctions connected in series  desired voltage
 numerous strings of series junctions connected in parallel  desired current
 series-parallel combination  maximum power transfer to the load
Possible causes of damage:
 Refractory metal insulation shield + graphite isotope capsule  can be damaged under
atmospheric oxygen operating at high temperature
 Sublimation damage may occur before the
mission start  RTG filled with pressurized
xenon gas until after launch
 Very  thermal conductivity of xenon 
possible to store and launch RTG hot and
generate needed pre-launch and launch
power
 Once in space  xenon is vented
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Radioisotope Thermoelectric Generator
GPHS-RTG Design
GPHS-RTG: General Purpose Heat Source – Radioisotope Thermoelectric Generator
 Specific RTG design used on US space missions 
 Ulysses  robotic space probe to orbit + study the sun (1990)
 Galileo  atmospheric-entry probe to the Jupiter (1989)
 Cassini-Huygens  space probe to the Saturn (1997)
 New Horizons  interplanetary space probe to perform flyby study of the Pluto (2006)
New Horizons GPHS-RTG, 300 W
Ulysses GPHS-RTG, 300 W
Cassini GPHS-RTG, 3x300 W
Credit: NASA
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Radioisotope Thermoelectric Generator
GPHS-RTG Design
 Design parameters 
 Overall diameter of 42.2 cm, length of 114 cm, mass of  57 kg
 Electrical power of 300 We at the start of mission, specific power of 5.2 We/kg, U = 30 V
 Radioisotope mass of  7.8 kg 238Pu, chemical form PuO2, number of GPHSs =18
Galileo:
Electricity (2xRTG)
and Heat (RHU)
http://nuclear.mst.edu
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Radioisotope Thermoelectric Generator
GPHS-RTG Design
238Pu
RTG pellet glowing red-hot from
its internal heat generation
1.4 kg
1.5 kg
1.6 kg
GPHS Module Illustration provided by DOE
Evolution of the GPHS
Module
GPHS Fuel Pellets. Foto: LANL
Source: Atomic Power in Space II, INL/EXT-15-34409
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Radioisotope Thermoelectric Generator
Heat Rejection Fin Generator Outer Housing Thermocouples
GPHS-RTG Design
Idaho National Laboratory
Electrical Cable
Connector
Graphite and
Iridium Protective
Layers
General
Purpose Heat
Source
Modules
Encapsulated
238PuO Fuel
2
Pellets
Image: INL RPS Program
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Radioisotope Thermoelectric Generator
MMRTG Design
MMRTG: Multi-Mission Radioisotope Thermoelectric Generator
 New generation of RTG  developed by the NASA and the DOE
 Designed to operate on Mars and in the vacuum of space
 Flexible modular design  to meet the needs of a wide variety of missions
 Optimized power level   110 W at launch, minimum lifetime of  14 years, ensuring
high degree of safety
www.nasa.gov
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Radioisotope Thermoelectric Generator
MMRTG Design
MMRTG
Thermoelectric
Couple
MMRTG: Multi-Mission Radioisotope
Thermoelectric Generator
 Composed of  8 standard GHPS modules
Thermoelectric
Modules
Housing
Thermal
Insulation
www.nasa.gov
Mounting
Interface
 Fuel  total of 4.8 kg PuO2 (including 238Pu)
  2000 Wth and 110 We
 Thermoelectric materials  PbSnTe, TAGS,
and PbTe  demonstrated extended
lifetime + performance capabilities
Cooling Tubes
Fin
8 GPHS Module Stack
Source: http://thermoelectrics.matsci.northwestern.edu
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Radioisotope Thermoelectric Generator
MMRTG Design
MMRTG: Multi-Mission Radioisotope Thermoelectric Generator
 The first NASA mission with MMRTG  Curiosity Mars rover (launched in Nov 2011)
The Curiosity rover with its
MMRTG visible at the rear
MMRTG 110 W
Installing MMRTG into Curiosity rover
www.nasa.gov
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Radioisotope Thermoelectric Generator
Electrical Model of RTG
 Total voltage generated by the Seebeck effect 
internal voltage:
𝑈 = 𝑈𝑂𝐶 −𝐼𝑅𝑖
𝛾 = 𝐼𝑆𝐶 /𝑈𝑂𝐶
𝐼 = 𝐼𝑆𝐶 −𝛾𝑈
Electrical equivalent circuit model
𝑃 = 𝑈𝑂𝑃 ∙ 𝐼𝑂𝑃
I-U characteristics of the RTG
𝑈𝑂𝐶
𝐼𝑆𝐶
𝑅𝑖
𝛾
𝑈𝑂𝑃
𝐼𝑂𝑃
𝐼𝑆𝐶 = 𝑈𝑂𝐶 /𝑅𝑖
- open circuit voltage
- short circuit current
- internal resistance
- characteristic admittance of the RTG
- voltage at operating point
- current at operating point
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Radioisotope Thermoelectric Generator
Maximum Power Transfer
 Power transferred from TEC to  the load:
𝑃 = 𝑈𝐼 = 𝑈(𝐼𝑆𝐶 − 𝛾𝑈) = 𝑈𝐼𝑆𝐶 −𝛾𝑈 2
 To extract the maximum power  the load must operate at given operating voltage:
𝑑𝑃/𝑑𝑈 = 𝐼𝑆𝐶 − 2𝛾𝑈 = 0 → corresponding operating voltage → 1 2 𝑈𝑂𝐶 and 1 2 𝐼𝑆𝐶
 Maximum possible power transfer:
𝑃𝑚𝑎𝑥 = 1 2 𝑈𝑂𝐶 ∙ 1 2 𝐼𝑆𝐶 = 1 4 𝑈𝑂𝐶 𝐼𝑆𝐶
 RTG conversion efficiency 
maximum at the max. power
transfer point
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Radioisotope Thermoelectric Generator
Effect of Temperature and Aging
 The I-U line shifts  for  T and  for  T
 The amount of shift  characteristics of the couple material
TEC current, power, and efficiency vs
operating voltage at given T
TEC current vs voltage at different T
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Radioisotope Thermoelectric Generator
 Aging  little effect on RTG output  heat source
has half-life in decades
Effect of Temperature and Aging
 Power generation degrades  little
 Most power degradation  due to slow precipitation
of phosphorus doping in the n-type leg of the
thermocouple
 I-U and P-U  shifts uniformly with time and
temperature such the max. P remains at the same U
I-U-P characteristics of RTG with age
TEC conversion efficiency vs temperature
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Radioisotope Thermoelectric Generator
Segmented TECs
 Unicouple TEC  the same heat seen by each segment in each leg
 High conversion efficiency  obtained by operating TEC material over wide temperature
range + using different materials in each T range where posses optimum performance
 Achieved by  Segmented TEC:
 p/n legs made of multiple segments of different materials in series 
conversion efficiency  15% vs 7%
 using state of-the art BiTe, FeSi, PbTe, SiGe alloys
 overall operating temperature drops from 1275 C to 300 C in typical segmented TEC design
a)
b)
Schematic principles of segmented (a)
and cascaded (b) TEC
Source: Pham, H. N., et al.: Design and Optimization of Segmented TE
Generator for Waste Heat Recovery. Roskilde: TU of Denmark, 2015.
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Radioisotope Thermoelectric Generator
Advanced RTG
 To support  defense and deep space missions (USA)
 Goal  triple conversion efficiency of the present TECs based on  advance SiGe cells
with improved thermal conductivity
 Power output  100 W
 Technologies:
 Segmented alkali metal thermal to electric converter (AMTEC) with 15% efficiency,
7-9 W/kg specific power
 Stirling engine with 25% efficiency and 6-7.5 W/kg
 Thermo-photovoltaics (TPV)
Goals of advance RTG R&D programs
 In each case  heat produced from
238Pu pellets, clad + protective layers
of graphite + iridium
Source: Patel, M. R.: Spacecraft Power Systems. Boca Raton: CRC Press, 2005.
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Thank you for your attention
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