Photovoltaic power for Space Station Freedom

The Space Station Freedom is described with special attention to its electric power system. The photovoltaic arrays, the battery energy storage system, and the power management and distribution system are discussed. The current design of Freedom's power system and the system requirements, trade studies, and competing factors which lead to system selections are referenced. This will be the largest power system ever flown in space. This system represents the culmination of many developments that have improved system performance, reduced cost, and improved reliability. Key developments and their evolution into the current space station solar array design are briefly described. The features of the solar cell and the array, including the development, design, test, and fight hardware production status, are given.<<ETX>>


INTRODUCTION
The Space Station Freedom (SSF) Phase 2 configuration Is shown In Fig. I, It will be a manned, multipurpose facility In low earth orbit (LEO) early in the next decade.
It wlll include four pressurized modules, Tile support, data, communicatlons, power, and other subsystems along with provisions for experiments and other equipment both inside and outside the pressurlzed modules. SSF will allow astronauts to live and work in space on a permanent basls.
The planned crew slze ranges from six to elght people.
The operational life of SSF wlll be Indef_nlte since it Is deslgned to be maintainable and can be resupplled on-orblt. SSF will be used for astronomlcal and terrestrial observations and for biological and materials processing experlments In mlcrogravlty.
Satellites and other space hardware can be repaired and _Intalned at SSF. It will also serve as a staging base for men and machines traveling to other orblts or other worlds.
The capabilities of SSF can grow because power and other subsystems are deslgned to increase In capability and to adapt to advanced technology to meet additlonal user and mission requirements.
A more detailed descrlptlon of SSF is given In Ref. I.
The Phase l conflguratlon of SSF (shown in Fig. 2) will be In LEO by 1990. The largest visual feature of SSF will be the eight photovoltaIc (PV) arrays at each end of the main truss.
Other parts of the power system that are less vislble Include the nickel-hydrogen battery energy storage subsystem, the Dower management and distribution subsystem electronics, and _he thermal control subsystem. These subsystems and the structure and other hardware to suoport them are calleO the ohotovoItaic power module (PVPM), which Is also shown in Fig. 2. SSF will nave four PVPM'S and each module has two solar arrays. Two major differences distlnguis_the SSF power _ystem from those of previous satellites: size and llfetlme.
SSF will have an Inltlal Phase I capability of 75 kW to the users with over a 250-kW array output at the beginning of llfe.
It will grow to 125 kW after Phase l and perhaps to as much as a 300-kW user power ultlmately.
The _SF system will dwarf typical satellite power systems which are about 3 kN In size.
Skylab, America's first space station, generated 12 kW of array power.
The Shuttle Orblter generates 22 kN peak from its fuel cells, but only about 7 kW Is usually used. As mentioned previously, the SSF and the power system will have Indefinlte llfe with maintenance.
Typical satellite power systems are not repairable and have Ilfetlmes of about 7 to lO years.
The evolution of the SSF power system from feaslbIllty through the prellmlnary design phase Is descrlbed In Ref. 2.
PO_ER SYSTEM DESIGN The electric power system (EPS) consists of the PVPM, the solar dynamic power module (SDPM), and the power management and distribution (PMAD) system. These hardware will generate, store, distribute, and control electric power.

Solar Dynamic
The SDPM (Fig. 3) will be Incorporated _n Phase Z of the SSF assembly to enable the power system to grow to as much as 300 KW. Solar dynamic (SD) power Is generated by focusing sunlight with an offset parabolic concentrator onto a heat receiver contalnlng a mixture of fluorlde salts. A portlon of the solar energy heats a hellum-xenon gas working fluid and part Is stored In the heat of fusion (solld to llquld) phase change of the Heat stored In the salt Keeps the cycle going during ecllpse.
The SD system advantage Is low 11re-cycle cost, due to the smaller surface area of the collector.
Thls smaller area results In less orbit decaying drag and, therefore, lower reboost fuel and resupply costs.
The area of the SD mirrors Is about one-half to one-thlrd that of the PV arrays. In addltlon, heat storage in the salt Is about 97 percent efflclent compared wlth electrlc storage in the PV batteries which Is about 80 percent efflcient.
CaIculatlons Indlcate that the overall sun-to-user efficiency of the SD system Is about 20 percent compared wlth about 7 percent for the PV system, uslng silicon cells and nickel-hydrogen batteries.
The disadvantages of the SD system are its lack of space operatlng experlence and a sun tracKing/positlonlng requirement more than 20 times more accurate than the PV arrays.
The SD mirror requlres very flne control methods.
Long-term operation of the salt heat storage system, degradatlon of the mirror optlcal quality, and heat engine life are addltlonal technical issues for which solutlons are being worked on.

Power Management and Distribution
The PMAD system dlstrlbutlon frequency and voltage are key features that determine the EPS system architecture.
Thls archltecture has changed several times durlng the SSF program. Initially, trade studies Indlcated that a 20-kHz dlstrlbution frequency was optlmum.
Several single-phase transitions were made for prlmary dlstributlon and secondary dlstrlbutlon.
Currently the system prlmary distribution Is 160 V do. The secondary dlstributJon Is 120 V do. Factors involved In these changes include cost, wlre slze, weight, power loss, electromagnetic Interference and compatlbIllty to power user's system, and experiment requirements.
The PMAD system is deslgned so that any combination of two failures wl]l not cause a loss of all electrIca] power.
Thls is accomplished by incorporatlng redundant swltching and controlling units and muItlpIe independent cables to each critical user IncIudlng the pressurlzed modu]es and the var-1bUS experimental pa]let load centers on the station truss.
Control of the EPS will be by semiautonomous local controllers linked to a central controller. Thls system will monitor and detect Faults, isolate malfunctionlng c1rcults, and reconfigure and recover system performance.
It wlll manage and, to an extent, schedule power use to help prevent overloads and to Insure full battery charge at the start of each ecllpse perlod.

PHOTOVOLTAIC POWERMODULE
The PVPM shown In Fig.  2  This cable carries the 0ower to the base of the array.
The folded blankets are contained In boxes (Fig. 6) for protectlon during launch.
The panels are compressed by latch and pre]oad mechanlsms to prevent solar cell damage due to launch vlbratlon and acce]eratlon forces.
Motor drive assemblles release the latches and allow the boxes to be _elatched _hen required.
The two-blanket contain-•ent boxes are attached to the top of a mast canis-:er which contains a foldable articulated square :russ <FAST) mast. Motor drives deploy the FAST •ast, and the mast pulls the folded panels out of the b;anket boxes to deploy and tenslon the array cn orolt.
Durlng launch the mast Is stored in its :a_isTer and the contalnment boxes are arranged _Ide oy side For mlnlmum packaging slze In the cargo bay of the Orblter.
Further details on the array design are glven In Ref. 3.
It Is 10-ohm-cm, p-type base resistivity wlth a shallow phosphorous diffused n-type junctlon and wlth a boron back surface Field.
SlgnlfIcant features of the cell are Its wrapthrough front contacts and grldded back contacts.
The Front collector grids converge on four round Insulated holes In the ceil.
The front metalllzatlon wraps through these holes to allow n contact _nterconnectlon on the back of the cell. The back contact Is grldded to allow transmisston of infrared (IR) llght that Is not absorbed by the silicon.
In conventional ceils with full coverage back contacts, the IR Is absorbed in thls metaI11zatlon and causes heatlng which reduces cell voltage and power output.
The IR energy comlng out the back of the cell Is also transmitted through the transparent panel substrate.
Thus the array operates at a lower temperature with hlgher voltage and power output.
Another benefit of wrapthrough contacts Is that both the n and p connections of the cell to the copper clrcult can be done In one operation from the back side of the ceil.
This slgnIflcantly reduces array panel assembly tlme and cost. The Interconnectlon of the cell sl]ver contacts to the copper traces In the panel substrate is accomplished by semiautomated paralle] gap welding. These interconnects have demonstrated a 15-year l|fetlme In a slmulated LEO thermal cycling environment.
At the time of this wrltlng, the array blanket components are In various stages of deslgn, The solar cells have undergone prellmlnary and cr_tlcal design reviews.
Quallflcatlon testing is in progress.
A total of over 265 000 fllght-type cells will be requlred.
The bypass diode has had Its preliminary and crltlcal design reviews.
Production of about 33 000 diodes is planned after quallfIcatlon testing at Advanced Optoelectronics.
The panel preliminary design revlew Is complete, and detailed design Is In progress.
Over 1344 panels w111 be produced by Lockheed.
Other array components are in the prellm|nary deslgn stage Includlng the containment box, the FAST mast, and other array mechanisms.
Development tests and thermal, structural, and fatlgue analysis are In progress. Extensive quallflcation and acceptance level testing at the component, coupon, panel. ORU, and assembly levels is planned.

Array Deslqn Heritaqe
The design heritage for the solar array traces Its roots to research and development programs which began In the 1970's.
These programs Included the solar eIectrlc propulsion stage (SEPS) for Interplanetary missions.
SEPS Incorporated a llghtwelght, hlgh-performance solar array uslng 2 by 4 cm wraparound contact cells to supply power (4). Thls array concept is the foundatlon on which the current space station array deslgn Is based. The SEPS program built a demonstration singleblanket wing whlch was later upgraded to the solar array fllght experiment (SAFE).
SAFE was successfully flown on the space shuttle In 1984 and demonstrated the array packaglng concept.
It a]so prov|ded data on array structural dynamic behavlor In zero gravity (5). In the early 1980's, the SEPS array design was proposed as a primary power source for the shuttle orbiter In the power extension package (PEP) program.
PEP studied the adaptlon of the SEPS Interplanetary array to the LEO environment, incIudlng the effects of thermal cycling on solar cell interconnects.
PEP developed and demonstrated pilot production of a 6 by 6 cm wraparound contact cell as a means to lower :e11 and array assembly costs through reduced plece parts handling (6).
After the First shuttle f11g_ts, the damaging effects of atomic oxygen (AO) In LEO were dlscovered.
Atomlc oxygen attacks certain spacecraft and solar array materlals and poses a threat to hardware survivability for long-term mlsslons.
In the mld-lgBO's, the space station program inltlated a development contractcalled photovoltalc array environmental protectlon (PAEP) to protect or replace array materials subject to damage by AO. The PAEP contract has developed and demonstrated coatings that successfully withstand the effects of AO (7). These protectlve materlals can be applled in large-scale processes and are compatlble with all array manufacturing and handllng steps.
Small materials samples and 11mtted scale flight tests are planned.
Preliminary results from the recently retrieved Long Duration Exposure FacIllty are encouraging and glve confidence in the AO protection methods.
The PAEP contract has also provlded protected array panel coupon segments and Full-scale panels. The coupons have been thermally cycled to slmulate LEO temperatures.
They have successfully survived IS years (87 000 cycles), a key requirement For the welded interconnects and the thln-panel substrate (8) The PAEP panels were used in a test that slmuiated the LEO space plasma envlronment to determine the Interactlon of the electrically actlve panel with the plasma. No signiFIcant parasitic currents or arcing events were detected under normal operatlng condltlons (g).

CONCLUSIONS
She electrlc power system is vitai to the success of Space Statlon Freedom.
Few of Freedom's research and exploratlon missions can be accomplished without rellable electrlc i)ower. The current EPS deslgn meets system requirements In a safe, cost effective manner.
It incorporates a b]end oF Flight-proven PV modules and promising hlgh-performance SD modules linked by the PMAD system.
The PV arrays and solar cells used on Space _ta:ion _reedom represent the product of numerous advances in array and cell design as well as advance_ in understanding the operatlonal environ-• ents.
These advances resulted From many research and development activltles throughout the 1970's and 1980's -only a portlon oF which have been mentioned _n this paper.
The photovoltalc community can be proud of a Job well done In providing Space Station Freedom, the largest'space PV system ever, wlth ]Ightwelght, fllght-proven designs, and manu-Facturlng capabillty.
More advanced cells and array designs being developed today extend the state oF the art beyond that of silicon cells and arrays.
HIth continued development, these new designs w_ll provide the space station wlth opportunltles For PV power system growth and technology evolution Into the 21st century. I , 2. 3.