1. Power/Solar Cells Brian Shepard Aerospace and Ocean Engineering Virginia Tech brshepa2@vt.edu http://www.aoe.vt.edu/~hokiesat Voice: (540) 2961-1483

2. Rate of Pressure Change - All components must be capable of surviving a pressure change of ± 2068 Pa/sec (0.0204 atm/s).

3. Magnetic Fields - All components shall be capable of surviving exposure to magnetic field strengths of up to 34 Gauss.

4. Construction Materials - All materials used in the construction of any spacecraft component must be listed on the NASA approved materials list (http://map1.msfc.nasa.gov/WWW_Root/html/page7.html).http://map1.msfc.nasa.gov/WWW_Root/html/page7.html)

5. Fasteners - All fasteners (nuts, bolts, rivets, inserts, etc.) on the spacecraft must be listed on the NASA approved fasteners list (http://lmd.gsfc.nasa.gov/fasteners/), and obtained through NASA sources. Nuts and bolts must be torqued to specified values and provide back out protection.

6. Wiring, Cables and Connectors - All ground paths in the spacecraft shall be consolidated to a single grounding tree (no ground loops). Cabling must be securely fastened by use of adhesive or approved material (high temperature rated) cable ties. TECSTAR recommended the use of 24-gauge wire for the solar cells. The military specification for this wire is the following: MIL-W- 22759/44. This wire is fluoropolymer-insulated with cross-linked Ethylene-tetrafluoroethylene copolymer (ETFE). The wire is silver coated copper.

7. Out-gassing and Venting - In space, all materials emit gases that reduce the vacuum in closed volumes, coat nearby surfaces with condensable material, and increase the likelihood of an electrical arc. Plastics, polymers, potting compounds, and coatings are particularly common sources of out- gassing. Out-gassing may be diminished by cyclical exposure to heat in a vacuum. Where components are positioned within an enclosure, sufficient vent openings must be present.

8. Schematic

9. System Parameters Detailed list of specific parameters describing the subsystem –This might include pressure, voltage, current, temperature range etc., but not mass, power or volume as these are called out specifically in later slides

10. Mass and Volume Subsystem Mass - A solar array consist of a Solar Cell (CIC), Silicone, Kapton, wiring, and connectors hooked together in a series on a given side of the satellite. The masses are as follows: 1) Solar Cell (CIC)- 250 grams 2) Silicone- 110 grams 3) Kapton- 25 grams 4) Wiring- 100 grams 5) Connectors- 50 grams This gives a combined mass of 500 grams

11. Mass & Volume

12. Subsystem Overview Subsystem requirements Safety requirements Schematic Operations summary Subsystem parameters Mass & volume Power Interfaces On-Orbit Operations Ground Operations Summary Analysis Hardware Subsystem Fabrication and Assembly Integration & test Remaining work Concerns

13. Subsystem Objective Solar arrays are the primary source of power for the satellite. These arrays of solar cells will convert radiation from the sun into usable power through the photovoltaic process. This subsystem is designed to generate power as an entity of the power system, which will regulate and control the distribution of power to the other systems of the satellite.

14. Subsystem Requirements Subsystem Mass - A solar array consist of a Solar Cell (CIC), Silicone, Kapton, wiring, and connectors hooked together in a series on a given side of the satellite. The masses are approximately as follows: 1) Solar Cell (CIC)- 250 grams 2) Silicone- 110 grams 3) Kapton- 25 grams 4) Wiring- 100 grams 5) Connectors- 50 grams This gives a combined mass of 500 grams.

15. Subsystem Power Interface - The solar arrays will need to transfer power from the individual cells of the subsystem to the current and voltage regulation circuitry controlling the distribution of power to the secondary power source (batteries) and the power consuming devices throughout the satellite. Wiring will facilitate this transfer of power. Each string of solar cells will have two wires that will go to the power board (V+, and GND), which will make 26 wires total unless the grounds are tied together at the panel connectors rather than all being tied inside of the electronics box.

16. Subsystem Data Interface - The flight computer will monitor the current coming off of each of the arrays of solar cells. This current reading will be generated by a set of 7 current sensors that will read the current from each of the arrays. This is the only piece of data that needs to be transferred to the rest of the satellite. There is no data that will be read into the subsystem.

17. Subsystem Power Consumption - This system provides power to the rest of the satellite. The solar cells that make up the arrays are designed to convert 23% of the sun’s radiation to usable power. The power exiting the current sensors going into the rest of the power system should be around 94% of the power that will be generated by the solar cells. This loss of power is due to the voltage drop across the reverse current protection diode that is located on each line coming to the power regulation system from the solar array.

18. Component Lifetime - The operational lifetime of all components shall be no less than four months. The non-operational storage lifetime of all components shall be no less than 24 months (two years).

19. Subsystem Survival and Operational Temperature Range - Component Survival Temperature(  C) Operating Temperature(  C) Kapton -250 to 400 -250 to 400 Solar Cells -100 to 150 -40 to 150 Silicone -115 to 260 -115 to 260 Connectors -40 to 80 -40 to 80

20. Static Structure Loads - No individual component or its structural interface shall experience yielding or buckling at 13.5 g’s in all axes simultaneously. This corresponds to a safety factor of FS > 1.2.

21. Dynamic Structure Loads - All components and it’s structural interface must be capable of withstanding 40 g shock loads in all axes, and 9.1 g RMS random vibration at the spectrum shown

22. Power Voc = 2.4 V Isc =.35 A Diode loss = 6.6 V Total Voc = 22.2V Voc = 2.4 V Isc =.35 A Diode loss = 6.6 V Total Voc = 22.2V

23. Interfaces Structural –Cells are bonded to substrate using the epoxy CV 2568. Power –24 gauge wire will be used in string to string connection. An Sn-62 (62.5% tin, 36.1% lead, 1.4% silver) will be soldered to end terminations of strings.

24. On-orbit Operations Summary Solar Cells will be taking the sunlight and converting it to usable power when the satellite is in orbit.

25. Ground Operations Summary Procedures for ground operations: –When solar cells are hooked to the satellite testing is recommended outside in natural sunlight between the hours of 11 a.m. and 1 p.m.

26. Analyses Detailed list of the analyses and tools that have been used to support the design –Solar cells to be set up at USU. –Facilities should provide accurate measuring devices and assembly tools.

27. Hardware Item 1, etc Description: GaInP2/GaAs/Ge high efficiency (22.6%) cascade solar cell. Included is an antireflective coverglass, interconnects for connecting cell to cell, and terminations. Heritage: Widely used in the Aerospace field. Manufacturer: Techstar Part number: Properties: Mass: ~3.07 grams Status: Available at USU for panel setup and layout.

28. Subsystem Fabrication and Assembly Describe how the subsystem is put together, –VT, USU, and UW will be assembling solar cell strings at USU at a later date. –Cells are placed on a specially designed aluminum block. This block keeps any stresses off of the fragile interconnects and allows for precise soldering of cells, interconnects, and terminations. –Groups of two or three are working together in order to assure soldering is satisfactory. –The stings of cells are attached to the surface skin of the panels using an epoxy that has been carefully created while soldering and cleaning the cells.

29. Integration & Test Initial test can be performed at USU to see if the cells are producing the desired current. The solar cells will be powering the satellite and therefor will need the other components of the satellite to test upon.

30. Remaining Work Sheets of Mylar with printed cell locations and 70% area for the epoxy will need to be made. These are a necessity when assembling the solar arrays. 5 days will need to be set aside so that two representatives from Virginia Tech can visit USU and assemble the individual strings of cells. Once panels have solar cells placed in there respective locations they will need to be transported back to Virginia Tech.

31. Unresolved Issues The solar cells are extremely delicate and expensive. Laying them out onto the panels should not be a problem with the expertise supervision at USU, but transporting them back here may cause cracking in the cells. Once the cells are on the panels if there is any more work that needs to be done in or around the panels it could potentially harm the cells. If a cell should crack, does VT have the facilities to fix or replace a solar cell? Time is running short and there is a lot of work to do.