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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 Space for Education, Education for Space 2 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 3 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 4 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 Space for Education, Education for Space 5 Power System Architecture Power sources and systems of satellites and deep space probes Space for Education, Education for Space 6 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 7 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 8 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 9 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) Power sources and systems of satellites and deep space probes Space for Education, Education for Space 10 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 11 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 Space for Education, Education for Space 12 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 13 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 14 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 15 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 16 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 17 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 18 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) Power sources and systems of satellites and deep space probes Space for Education, Education for Space 19 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 20 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 21 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 22 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 23 Spacecraft Power Systems Architecture Trades Power sources and systems of satellites and deep space probes Space for Education, Education for Space 24 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 25 The ISS Power System https://spaceflight.nasa.gov Power sources and systems of satellites and deep space probes Space for Education, Education for Space 26 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 27 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. Power sources and systems of satellites and deep space probes Space for Education, Education for Space 28 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 29 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 30 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 31 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) Power sources and systems of satellites and deep space probes Space for Education, Education for Space 32 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 33 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 34 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 35 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 36 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) Power sources and systems of satellites and deep space probes Space for Education, Education for Space 37 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 38 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 39 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 40 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 41 Power System Sources Power sources and systems of satellites and deep space probes Space for Education, Education for Space 42 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 43 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 44 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 45 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 46 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 47 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 49 Solar Array Systems PV Cell Efficiencies Power sources and systems of satellites and deep space probes Space for Education, Education for Space 50 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 51 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 52 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 53 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 55 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 56 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 57 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 58 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 59 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 60 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 61 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% Power sources and systems of satellites and deep space probes Space for Education, Education for Space 62 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] Power sources and systems of satellites and deep space probes Space for Education, Education for Space 63 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 64 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 65 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 66 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 67 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 68 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 69 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 70 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 71 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 72 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 73 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 74 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 75 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 76 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 77 Battery Systems Source: The Electrochemical Society Interface • Fall 1999 Battery types Power sources and systems of satellites and deep space probes Space for Education, Education for Space 78 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 79 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) Power sources and systems of satellites and deep space probes Space for Education, Education for Space 80 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 81 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 82 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) Power sources and systems of satellites and deep space probes Space for Education, Education for Space 83 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 84 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space NASA 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 86 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 87 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 88 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 89 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 90 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 91 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 92 Power Systems Power System Design Trades Flow-Chart Power sources and systems of satellites and deep space probes Space for Education, Education for Space 93 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% Power sources and systems of satellites and deep space probes Space for Education, Education for Space 94 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 95 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 96 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 97 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 98 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 99 Fuel Cell Power Source: Argonne National Laboratory Electrical Performance Power sources and systems of satellites and deep space probes Space for Education, Education for Space 100 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 101 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 102 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 103 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 104 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 105 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 106 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 107 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 108 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. Power sources and systems of satellites and deep space probes Space for Education, Education for Space 109 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 110 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 111 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 112 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 113 Radioisotope Thermoelectric Generator RTG Radioisotopes Source: U.S. Atomic Energy Commission Power sources and systems of satellites and deep space probes Space for Education, Education for Space 114 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 115 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) Power sources and systems of satellites and deep space probes Space for Education, Education for Space 116 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 117 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 118 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 119 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. Power sources and systems of satellites and deep space probes Space for Education, Education for Space 120 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 122 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 123 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 124 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 125 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 126 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 127 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 128 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 129 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 130 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 131 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 132 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 Power sources and systems of satellites and deep space probes Space for Education, Education for Space 133 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. Power sources and systems of satellites and deep space probes Space for Education, Education for Space 134 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. Power sources and systems of satellites and deep space probes Space for Education, Education for Space 135 Thank you for your attention Power sources and systems of satellites and deep space probes Space for Education, Education for Space 136