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Power/Solar Cells
Brian Shepard
Aerospace and Ocean Engineering
Virginia Tech
[email protected]
http://www.aoe.vt.edu/~hokiesat
Voice: (540) 2961-1483
• Rate of Pressure Change All components must be capable of surviving a pressure change of
± 2068 Pa/sec (0.0204 atm/s).
• Magnetic Fields All components shall be capable of surviving exposure to magnetic
field strengths of up to 34 Gauss.
• 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).
• 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.
• 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-W22759/44. This wire is fluoropolymer-insulated with cross-linked
Ethylene-tetrafluoroethylene copolymer (ETFE). The wire is silver
coated copper.
• 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 outgassing. 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.
Schematic
Panel 1
Panel 5
Panel 3
Panel 2
Panel 6
Panel 4
Anti-Nadir
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
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) Silicone110 grams
3) Kapton25 grams
4) Wiring100 grams
5) Connectors50 grams
This gives a combined mass of 500 grams
Mass & Volume
Specifications
Length
Width
Cell Size
Thickness
Length
Interconnect
Width
Length
Termination
Width
Length
A/R
Width
Coverglass
Thickness
By-pass diode voltage drop
Vopen circuit @ 28C
Vmax power @ 28C
Ishort circuit @ 28C
Imax power @ 28C
Cff @ 28C
Efficiency
Mass (g)
Absorptivity
Emissivity
Value
2.497 in
1.522 in
0.0063 in
0.383 in
0.09 in
2.2 in
0.1 in
2.497 in
1.522 in
0.004 in
0.6 V
2.47 V
2.1 V
370 mA
350 mA
82 %
22.6 %
3.10
0.91
0.88
Subsystem Overview
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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
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.
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) Silicone110 grams
3) Kapton25 grams
4) Wiring100 grams
5) Connectors50 grams
This gives a combined mass of 500 grams.
• 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.
• 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.
• 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.
• 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).
• 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
• 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.
• 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
Power
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Voc = 2.4 V
Isc = .35 A
Diode loss = 6.6 V
Total Voc = 22.2V
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.
On-orbit Operations Summary
• Solar Cells will be taking the sunlight and converting it to usable power
when the satellite is in orbit.
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.
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.
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.
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.
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.
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.
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.