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LONG SOLAR CELL INFO

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					                                                                              (Document number)**

VT HOKIESAT ASSEMBLY PROCEDURE                                                     DATE:_____________
Part # **____________________

Safety notes, Prep work: Any safety precautions or notes should be listed here. For example, if rubber
gloves, aprons, etc should be worn, specify that here. For clean room assemblies, make a note that all
parts including fasteners should be properly cleaned before entering the clean room area.

Step-by-Step Procedure; __ steps, __ pages

List steps as numbered bullets. Be as thorough as possible.

If procedure requires several major steps/hours/days to complete, include a time log as shown in the
following section.

The last step in your procedure should be to record the new part number in the tracking log, complete
configuration log, and complete assembly log. The numbers for these logs will be assigned when they are
picked up before assembly begins.

Time Log

Date: ________ Start procedure time: _________ a.m. / p.m. Stop procedure time: _________ a.m. / p.m.

Date: ________ Start procedure time: _________ a.m. / p.m. Stop procedure time: _________ a.m. / p.m.

Date: ________ Start procedure time: _________ a.m. / p.m. Stop procedure time: _________ a.m. / p.m.

Etc…


                                           Assembly Verification Record
Assembled by: ________________________               Signature: ___________________________ Date: ________

Witnessed by: ________________________               Signature: ___________________________ Date: ________



Required Tools:

Include a list of the tools used in the procedure.



**Part numbers and document numbers are assigned after submittal.
                                             (Document number)**


               USUSAT CDR / ION-F IDR DOCUMENT:
                         POWER SYSTEM
                    SOLAR CELL SUB-SYSTEM

System Lead:
Chad Fish
Subsystem Lead :
Kenneth Vanhille
                         Document Authors
Kenneth Vanhille
Tyler Loertcher
Darin Thorson




Last Updated:              1/4/01 1:00 PM
Document Number:      AA-22702-DOC02-1.doc
                                                                 (Document number)**



                                SIGNATURE PAGE




Reviewed By:
               Rees Fullmer / Charles Swenson, Co-Investigator                Date


Reviewed By:
               Brian Lewis, USU Systems Engineer, brian.lewis@sdl.usu.edu
      Date


Reviewed By:
               Chad Fish, Power Lead, chad.fish@usu.edu                       Date

Reviewed By:
               Kenneth Vanhille, Solar Array Lead, kvanhille@cc.usu.edu       Date
                                                         (Document number)**


                             REVISION HISTORY
Revision   Description                 Author         Date        Approval
1          Initial Release             Tyler          8/18/00     Pending
                                       Loertscher
2          Changed from a power system Ken Vanhille   1/4/01      Pending
           to a solar array subsystem
           document.
                                                                                                               (Document number)**


                                                TABLE OF CONTENTS
SIGNATURE PAGE..................................................................................................................2
REVISION PAGE .....................................................................................................................3
TABLE OF CONTENTS .......................................................................................................4 - 5
APPLICABLE DOCUMENTS .....................................................................................................6
LIST OF TABLES ....................................................................................................................7
LIST OF FIGURES ...................................................................................................................8
LIST OF ABBREVIATIONS .......................................................................................................9
1         PROBLEM DEFINITION .............................................................................................10
1.1       SYSTEM OBJECTIVE ................................................................................................10
1.2       REQUIREMENTS.......................................................................................................10
          1.2.1      SUBSYSTEM MASS AND VOLUME ...................................................................................... 10
          1.2.2      SUBSYSTEM POWER INTERFACE ........................................................................................ 10
          1.2.3      SUBSYSTEM DATA INTERFACE .......................................................................................... 10
          1.2.4      SUBSYSTEM POWER CONSUMPTION................................................................................... 11
          1.2.5      COMPONENT LIFETIME ...................................................................................................... 11
          1.2.6      SUBSYSTEM SURVIVAL AND OPERATIONAL TEMPERATURE RANGE .................................. 11
          1.2.7      STATIC STRUCTURE LOADS ............................................................................................... 11
          1.2.8      DYNAMIC STRUCTURE LOADS ........................................................................................... 11
          1.2.9      RATE OF PRESSURE CHANGE ............................................................................................. 11
          1.2.10     MAGNETIC FIELDS ............................................................................................................. 11
          1.2.11     CONSTRUCTION MATERIALS.............................................................................................. 11
          1.2.12     FASTENERS ........................................................................................................................ 12
          1.2.13     WIRING, CABLING AND CONNECTORS .............................................................................. 12
          1.2.14     OUTGASSING AND VENTING .............................................................................................. 12
2         CONCEPTUAL DESIGN .............................................................................................13
2.1       TRADE STUDIES ......................................................................................................13
          2.1.1      REGULATED BUS ............................................................................................................... 13
          2.1.2      UNREGULATED BUS........................................................................................................... 13
2.2       PRELIMINARY TESTING OR SIMULATIONS ...............................................................13
2.3       CONCLUSIONS .........................................................................................................14
3         DETAIL DESIGN ......................................................................................................14
3.1       SUB-SYSTEM OVERVIEW ........................................................................................14
          3.1.1      BATTERY SYSTEM ............................................................................................................. 14
                                                                                                             (Document number)**

      3.1.2     SOLAR ARRAYS ................................................................................................................. 14
      3.1.3     SHUNT ............................................................................................................................... 16
      3.1.4     POWER BOARD .................................................................................................................. 17
      3.1.5     MASS AND VOLUME .......................................................................................................... 21
      3.1.6     POWER REQUIREMENTS ..................................................................................................... 21
      3.1.7     DATA REQUIREMENTS........................................................................................................ 21
      3.1.8     POWER CONSUMPTION ...................................................................................................... 21
      3.1.9     PARTS LIST ........................................................................................................................ 21
4     SUB-SYSTEM VERIFICATION ...................................................................................22
4.1   QUALITY CONFORMANCE METHODS ......................................................................22
      4.1.1     INSPECTION ....................................................................................................................... 22
      4.1.2     DEMONSTRATION .............................................................................................................. 22
      4.1.3     ANALYSIS .......................................................................................................................... 22
      4.1.4     TEST .................................................................................................................................. 23
4.2   VERIFICATION MATRICES .......................................................................................24
5     SUB-SYSTEM OPERATION .......................................................................................25
5.1   OPERATIONS BLOCK DIAGRAMS .............................................................................25
5.2   OPERATING SOFTWARE ...........................................................................................25
6     POST CDR / IDR SCHEDULE ...................................................................................27
7     SUMMARY ...............................................................................................................27
                                         (Document number)**


                  APPLICABLE DOCUMENTS


Document Number   Title
NASA-STD-7002
                                                                                   (Document number)**


                                         TABLE OF TABLES


TABLE 1.1 SUBSYSTEM MASS AND VOLUME ...................................................................10
TABLE 1.2 SUBSYSTEM SURVIVAL AND OPERATIONAL TEMPERATURE RANGE ................11
TABLE 1.3 SUBSYSTEM SURVIVAL AND OPERATIONAL TEMPERATURE RANGE ................21
TABLE 1.4 VERIFICATION MATRICES ................................................................................24
TABLE 2.1 TRIPLE JUNCTION CELL MODELING ASSUMPTIONS...........................................17
TABLE 2.2 POWER DIFFERENCES FROM SATELLITE ORIENTATION.....................................18
TABLE 1.5 POST CDR/IDR SCHEDULE..............................................................................27
                                                                                      (Document number)**


                                          TABLE OF FIGURES
FIGURE 1.1 RANDOM VIBRATION SPECTRUM ......................................................................12
FIGURE 3.1 SOLAR CELL LAYOUT ......................................................................................15
FIGURE 3.2 SOLAR CELL DIMENSIONS ................................................................................16
FIGURE 3.3 SOLAR ARRAY BLOCK DIAGRAM .....................................................................16
FIGURE 3.3 USUSAT DET POWER SYSTEM ........................................................................17
FIGURE 3.4 WIRING/CONNECTOR DIAGRAM .......................................................................18
FIGURE 3.5 BATTERY MONITOR..........................................................................................18
FIGURE 3.6 CHARGE CONTROL CURRENT REGULATOR.......................................................19
FIGURE 3.7 SHUNT ACTIVATION CIRCUIT ...........................................................................20
FIGURE 4.1 SOLAR CELL CURVES .......................................................................................23
                                                                                              (Document number)**


                                         LIST OF ABBREVIATIONS
FSK ................................................................................................. Frequency Shift Keying
STK .......................................................................................................... Satellite Tools Kit
FPGA…………………………………………..…………Field Programmable Gate Array
USUSat ................................................................................. Utah State University Satellite
SDL .......................................................................................... Space Dynamics Laboratory
CIC ........................................................................ Coverglass-Interconnect-Cell Assembly
ETFE .....................................................................Ethylene-tetrafluoroethylene Copolymer
IPA .......................................................................................................... Iso-propyl Alcohol
PCB ..................................................................................................... Printed Circuit Board
                                                                     (Document number)**

Sub-System 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 Mass And Volume
                     Table 1.1
      Solar Array           Volume         Mass (g)
      Subsystem
      Solar Cells (CIC)     Unknow         250
                            n
      Silicone              Unknow         110
                            n
      Kapton                Unknow         25
                            n
      Wiring                Unknow         100
                            n
      Connectors            Unknow         50
                            n
      Total                 Unkno          500
                            wn
1.1.1 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 20 wires
total unless the grounds are tied together at the panel connectors rather than all being tied
inside of the electronics box.

1.1.2 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 5 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.

1.1.3 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 24% 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%
                                                                    (Document number)**

of the power that will be generated by the solar cells. This lose 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.

1.1.4 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).

1.1.5 Subsystem Survival and Operational Temperature Range

                                      Table 1.2
            Component              Survival           Operating
                                   Temperature(      Temperature(
                                   C)                 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



1.1.6 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).


1.1.7 Wiring, Cabling 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.

1.1.8 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.
                                                                    (Document number)**




1.1.9 Solar Array Assembly Location
ION-F has taken the approach of allowing USUSat to lead the constellation for the
research of solar cell assembly methods. As a result of the need to spearhead the solar
array development, USUSat sent a representative during September of 2000 to
TECSTAR and TRW (both located in L.A.) to learn about the assembly process from the
solar cell manufacturers. A summary of this trip is included in the Conceptual Design
section of this document. Because of the expertise that developed at USU, UW and VT
wish to collaborate with USU in the assembly of their respective solar arrays. This study
was conducted to identify a satisfactory location and procedure for this collaboration to
take place.

2.1.1.3 ION-F Comes to USU to Assemble the Panels
Each of the universities sends a representative to USU to learn to assemble the solar
panels and then they will stay at USU to assemble their own arrays, which they will then
take back to their respective universities. The advantage to this plan is that each school
will have their own personnel assemble their own arrays. The advantage also exists in
the fact that they will have the availability of the facilities of SDL, which are probably
superior to other facilities. There is also the advantage in that UW and VT will have the
resource of Ken to speak with as they assemble their panels. Disadvantages are that each
university will have to send someone to USU for an extended period of time (probably
around a week). There is also the disadvantage that the assembled product would have to
be shipped from SDL to the respective universities. The final assessment of this option is
that it would be the best choice.

2.1.1.4 Final Decision
The third option is recommended option. UW and VT could send someone to USU with
their panels and the solar arrays would be assembled here and then the finished panels
would be sent back. This sounds like the best idea because the solar cells will be at SDL
to start out with. The ability to watch USU assemble their own solar arrays and then
assemble in the same facilities is advantageous. It is possible for the other universities to
solder together the solar cells and then mount them on something temporary to be
shipped back to the other universities depending on whether the flight panels are ready to
have the solar cells mounted to them at the time of mounting. Sometime mid to late
February is the projected time for USU to be ready to host the other universities for solar
cell assembly. Each university would probably need about a week to assemble their
arrays.

1.1.10 Silicone Selection
Quite a few different silicone substances were looked at to use in the bonding process
between the satellite and the solar cells. This study involved looking at the differences
between the silicones used to bond the solar cells to the layer of Kapton.
                                                                   (Document number)**


1.1.10.1       CV10-2568
This silicone was found a little later in the design. Nusil, who also manufactures the
other silicones as well, manufactures this variety. The drawback of this silicone as it is
related to the CV 2568 is that it is a rather complicated silicone to apply because of its
two-part nature. The advantage of this silicone is that it has the desirable properties of
CV 2568, but it is also heat cured. This silicone is not manufactured by Dow Corning as
is the other two silicones, for this reason it was not found until later.

1.1.10.2       Silicone Summary
CV10-2568 was chosen because of the advantage that it has over the CV 2568 since it is
heat cured. The disadvantages and the problems posed by CV 1149 were to imposing to
ignore, so it was not chosen despite its simplicity to apply. The CV10-2568 will produce
the quality of bond that is required for our application with the solar cell arrays. Despite
the increased complexity of its application, the advantages outweigh the drawbacks for
this silicone.


Tyler Loertcher (USU) performed MATLAB simulations of the Emcore and TECSTAR
solar cells and combined those simulations with that of the entire power system to
determine the compatibility of the solar cell design with that of the rest of the power
system. The results of these simulations indicate that the efficiency of the power system
as a whole should be anywhere from 80%-90% if it is mated properly to the rest of the
satellite. Results of those simulations with the coding for the simulation can be found in
their entirety in his documentation for his senior project, however some of the pertinent
graphs and information can be found in the following figures and tables. Two major
simulations were conducted by Tyler, which will be explained in the following sections
of this document. He has labeled these simulations “rough” and “advanced” simulations
and the nomenclature is repeated here.

Due to mass limitations and the desire to keep the mechanical design of the spacecraft as
simple as possible, it was decided at this early design stage to use only body mounted
cells in the power system design. Every available surface was thus assumed to be
covered with solar cells. Throughout the design process the exact size, number, and type
of solar cells was refined in the simulation files. However, the concept of the rough
simulation remained the same.

The rough simulation took as input a sun vector ephemeris, generated by the satellite
propagation software STK (Satellite Tool Kit). This input was then processed using
information such as satellite orientation, solar cell configuration, and cell efficiencies to
produce an output of expected output power from the solar arrays based on an optimum
load. These data were then further processed to give estimations of orbital average
power.
                                                                    (Document number)**

The functionality of the solar arrays is simplistic in nature, but the task of assembling
these arrays is indeed daunting. The following sections will examine the specific parts
involved in the assembly process and an outline of the actual process. The outline of the
process is taken from a document comprised of notes and information gathered by Ken
Vanhille after his trip to TECSTAR and TRW September 26-27, 2000.


3.1.2 Solar Cells
Due to the great expense of high efficiency solar cells the only option open to the
designers of HokieSat has been the TECSTAR, triple junction cells. These were chosen
because the Air Force has provided them to the program. These cells have a typical
efficiency of 24% and are being provided with bypass circuitry (diode), cover glass, and
mechanical interconnects for each cell (CIC, pronounced “kick”). This leaves the
construction of the arrays as the major remaining hurdle to overcome. Being that the
construction of the solar arrays is mainly a mechanical problem, not much on this design
had been considered beyond the number of cells on each surface of the spacecraft until
the beginning of August 2000.

There are a total of 156 solar cells in the design. These are strung in strings of twelve to
provide the desired bus voltage. Each of these thirteen strings is protected by a diode to
prevent current from flowing in the reverse direction through the array. This is not only
necessary to prevent the batteries from discharging through the array, but also to prevent
illuminated strings from discharging through shaded strings.
3.1.3 Wiring
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) and is silver coated copper. It is stranded wire containing roughly 36 strands
within the wire. This spec was taken from the document MIL-W-22759/44A, which
describes in great detail the proper wire for this application. This wiring will run from
the end terminations through holes located upon the corresponding panel of the satellite.
After passing inside of the satellite, these wires will run to a connector located on each
panel that will serve to join the wire from the cells to the wires running into the power
board. At the connector the grounds will be tied together as well as the positive pins for
each individual panel. A table showing the necessary connectors for the flight is included
below. A diagram of this wiring pattern for each of the panels is shown below.
                                                                                               (Document number)**


1.1.11 Parts List
                                                  Table 3.3 Parts List
USUSat Parts List
Subsystem or Component : Solar Arrays
Responsible Party : Kenneth Vanhille


Part Description       ID Number         # of Parts       Supplier                    Lead Time   Contact Information
Flight Solar Cells     Triple Junction                 87 Tecstar                     Next Week   Bob Langdon (626) 934-6541
Mechanical CICs        Triple Junction                100 Tecstar                     At SDL      Bob Langdon (626) 934-6541
Protection Diodes      15MQ040N                        10 International Rectifier     ??          Jeff Grutter (310) 563-1476
Current Sensors        MAX471                           5 Maxim Integrated Products   ??          Maxim
Solar Cell Silicone    CV 10-2568                100 g NuSil                          At SDL      Brian W. Burkitt (805) 566-4166
Substrate Silicone     CV 2289                 100 mL NuSil                           At SDL      Brian W. Burkitt (805) 566-4166
Primer                 CF2-135                    4 oz. Dow Corning                   At SDL      Brian W. Burkitt (805) 566-4166
Kapton                 200HPP-ST            25” x 400” Fralock                        At SDL      Oscar Perdomo (800) 372-5625
End Terminations       205342-12                      382 American Etching and Man.   At SDL      Gary Kipka (818) 896-1187
Solder Wire            Sn62Pb36Ag2                ¼ lb Technical Devices              At SDL      Pam (801) 972-5939
Wire                   Mil -W-22759/44       Unknown Raychem                          Unkown      None



The parts list can also be found in its entirety on the Thidwick server at
:(ftp://thidwick.ece.usu.edu/pub/nanosat/ususat/powersystem/AA-22702-DOC02-1.doc).

1.2         Assembly Procedures
September 26, 2000 Ken Vanhille (Utah State University) represented ION-F in a visit to
TECSTAR and TRW in the Los Angles area. These visits were necessary to enable ION-
F a better understanding of the procedure needed to assemble the solar strings and arrays
once TECSTAR delivers the solar cells that have been ordered. Robert Langdon has
been the ION-F contact at TECSTAR for the solar cell order. He arranged for a meeting
with Jim Hanley, Director of Solar Panels, at the aforementioned company. Marshall
Cannedy, Director of Solar Panels at TRW, welcomed Ken Vanhille and Dr. Charles
Swenson for a discussion and tour of their facilities.

TECSTAR agreed to outline the entire process necessary for the construction of solar
panels. This process starts with receiving the CICed cell and continues all the way
through testing the assembled panels. The acronym CIC stands for
Coverglass/Interconnect/Cell, which refers to the condition of the solar cells that we will
be receiving. The following information will be a synopsis of the process outlined by
TECSTAR with additional insights from Marshall Cannedy as well. Prices on all of
these items have not been included, but they will be included in subsequent revisions.
                                                                    (Document number)**


1.2.1 Assembly of Strings
Assembly of the strings is the first task that ION-F will encounter once the cells are
received. String assembly in industry is done by welding, a process that has been taking
over in the last 5 years. ION-F will use soldering, which is still used for manual touch-up
jobs. Welding requires facilities that are much more elaborate than anything that is
conceivably possible for the resources of ION-F. Soldering should be done using Sn 62
wire. This wire contains 62.5% tin, 36.1% lead and 1.4% silver. This alloy is more
desirable than a pure mixture of lead and tin because of the silver plating on the back of
the cells and the interconnects as well. TECSTAR uses Kester wire that is .020” in
diameter. The diameter is not too thin, so this shouldn’t be too difficult to find.
TECSTAR does not normally use any additional flux. They use wire with an R-flux
core, which seems to be enough for them to use in their applications. Standard Weller
soldering workstations can be used for the soldering work with a temperature of near 300
degrees Celsius or 750 degrees Fahrenheit. USUSat acquired a quarter pound of solder
wire from Technical Devices in Salt Lake City on sample.

The solar cells should be set on a specially designed block to be soldered. The figure is
on the following page. This block can be made out of aluminum, and the blocks that
were used at TECSTAR were made of aluminum. This assembly block is necessary
because the interconnects are stress relieved. This stress relief creates the necessity of a
notch in the assembly block to keep the face of the cells completely level that must rest
upon the assembly block. Without the notch in the block the cells would be much more
likely to break because of the interconnect as pressure is applied during the soldering
process. The assembly block should be long enough to allow for the assembly of an
entire string of solar cells with end terminations on the outside sides of the end cells.
With the block as long as the string this will allow the assembler to assembly any
combination of solar cells smaller than this size as well. For instance, if a string were to
be placed in two rows such that one part contained 3 cells and the other part contained 5
cells, this would be possible to do on this assembly block.

                                                     End
                                Interconnect         Terminatio
                                Slot                 n




                         Figure 3.4 Rough Assembly                    Cell
                         Block                                        Space
                                                                 (Document number)**

If soldering touch-up work needs to be done once the cells are placed on the substrate,
make sure that the substrate doesn’t get to much heat such that the heat damages the
Kapton or substrate. When soldering wires to the end terminations, be careful not to
reheat the connection between the end termination and the interconnect from the cells
such that previously done connection goes bad. These ideas should cover the assembly
of the solar cell strings. One item of note is that these solder joints should be well
cleaned with a Cotton swab. Use two solutions to clean the cells. These two solutions
should include first Acetone and then Iso-Propyl Alcohol (IPA).




                              Figure 3.5- Assembly Block
                                                                     (Document number)**


1.2.2 Array Layout
Once the strings have been connected together they will need to be laid out on the
substrate of the Solar Panel. This preparation includes examining the cells to make sure
that the cracks that have developed are not critical. TECSTAR was confident in saying
that cracks will occur at this point in time. It is quite definite that they will be present.
The hope is that all of the cracks run parallel to the flow of current rather than
perpendicular to the current flow. If the cracks run parallel to the flow of current, there
shouldn’t be a problem, but if the cracks are more than a quarter across the cell it should
be replaced. A dinged corner of a cell shouldn’t be a problem if there are not any large
cracks that have been produced by the ding. See the following figure for details.

Before the front of the solar cells can be examined the solar cells must be taken from the
off of the assembly block and flipped so that the bottom side is down. Getting a sturdy
material that can be used as a strong back and placing it over the assembled sting can do
this. The entire combination including the sting, the strong back and the assembly block
can now be flipped over together so that the string assembly is now side up. From here
the coverglass of the solar cells can be examined so that the cracks that were referenced
to in the proceeding paragraph may be looked for.
                             Figure 3.6 Cell Crack
                             Diagram




                                                             Cracks are
                            Cracks are
                                                             unacceptable.
                            OKokaaceptableac
                            ceptable.
The separate strings for each panel will now be joined together so that the entire panel
                                                                  (Document number)**

                                                 Kapton
                                                 Tape




                                Figure 3.7 Taping of the
                                Solar Array



can be manufactured. In order to accomplish this the first thing that must be done is to
plot two 1:1 copies of the panel on Mylar. Vellum can be used rather than Mylar, but
Mylar is the best material that can be used. What needs to now be done is that the strings
should be placed on one copy of the panel in the exact location that they are to occupy.
The top of the cell should be face up, so that the strings lay flat. These should then be
taped together with Kapton tape so that each string is connected to the string next to it.
Take the other sheet of Mylar and punch two holes in each of the spaces for the solar
cells so that the holes line up horizontally. These holes will be important for connecting
the cells together.


                     Figure 3.8 Mylar                            Holes punched
                     Layout                                      out




                                          Mylar sheet with 1:1 solar array
                                                                    (Document number)**

It may be necessary to make cutouts in the Mylar sheet for the interconnects as well.
This was not something that was discussed at TECSTAR, but in thinking further about it,
it may be necessary to get the Mylar to lay flat on the tops of the solar cells. Strips of
Kapton tape are placed on the sheet over the holes. This would mean that the holes must
not be in line with the interconnects or else the tape would have to arch over the
interconnects rather than laying down flat. A Cotton swab is then used to affix the
Kapton tape through the holes of the Mylar to the tops of the solar cells on the other side.
Apply enough pressure with the Cotton swab to affix the coverglass of the solar cell to
the Kapton strips of tape. This will allow the assembler to pick up the solar cells with the
bottoms of the cells exposed so that it can be eventually placed upon the substrate.

Either at this time, or else when the strings of solar cells are still on the assembly block
the backs of the solar cells should be cleaned. This cleaning process will include first a
clean with Acetone and then a clean with IPA, which should be done three times before
proceeding. This is something that is necessary to insure that there is a good surface for
the primer to be applied to.

                         Kapton tape stripes      Figure 3.9 Taped
                                                  Mylar




                                           Directly over holes, touching
                                           coverglass.

1.2.3 Preparing the Substrate
As the preparations are made for the backs of the solar cells to be affixed to the substrate,
the substrate must also be prepared for priming as well. All of these processes should be
done in as clean of an environment as is possible. TECSTAR recommended that work be
accomplished in a class 10,000 environment, although they said that a class 100,000
would probably be acceptable.

There needs to be a dielectric check performed on the substrate. The dielectric check will
insure that there is not any leakage of charge through the Kapton, so that the backs of the
solar cells are certain to be electrically isolated. This can be done with a megohmmeter
                                                                   (Document number)**

and alcohol or water. Do not use water because it takes longer to dry than alcohol; wait
24 hours after applying water to insure that all of it has dried. The Kapton layer on top of
the substrate should be 2 mils thick. Apply alcohol evenly to the substrate and then
slowly pass the probe of the megohmmeter over the entire surface to know if charge is
leaking through the Kapton. If there is a problem with charge leaking there must be a
new layer of Kapton applied to the substrate. Once this has been done a very thorough
cleaning of the substrate must be conducted. To clean use a cotton swab and wipe in a
pattern where all of the wiping is in on direction and then all of the wiping is done in a
perpendicular direction to that. This pattern should be employed rather than a circular
pattern for all cleaning applications. The circular pattern will not clean; it will only
spread the dust and other particles around the surface to be cleaned.

Kapton HPP-ST is the variety of Kapton that is being used for the design of the satellites.
Figure 3.10 is an example layout of the Kapton sheet
needed for all of ION-F. This is actually 25X400”,
but it is shown here as being 50X200” to save on
space. What it entails is enough Kapton to completely
cover each satellite twice. This is enough Kapton to
allow for errors in design and for prototyping as well.
It was decided to choose this amount of Kapton so
that there will be enough for the flight panels and so
as to be able to have enough Kapton for ample testing
and an error range as well. Since all of this will be
done at USU, we will not need nearly as much of a
margin for testing Kapton. The panels for UW and
VT were arbitrarily chosen to be approximately
14X10” and the USU side panels are 6X10”. The top
and bottom panels are the same size for each of the
satellites, major diameter at 20” and the minor
diameter at 17.1”. Visio was used to draw this
drawing.

The cells and the substrate should be primed at the
same time. Use SP 120 or SP 121, either of which can
be purchased from Nusil. Wait at least a half hour
before applying the silicone, but no more than 18
hours; this is the window in which the primer will
produce the strongest bond. This should all take place
in the clean room that was specified earlier.

Another test that should be done on the substrate is to
put it through four cycles with a thermal vacuum. Use
the temperature limits that the satellite will see as the
limits for the thermal cycle. If the Kapton layer stays
well sealed to the substrate, there should not be a
problem with the bond strength between the Kapton



                                                               Figure 3.10 Kapton
                                                               Usage
                                                                    (Document number)**

and the aluminum substrate. Make certain to use space-grade Kapton for the dielectric
layer. The adhesive used to bond the Kapton to the substrate should be CV 2289.

1.2.4 Bonding the Assembly
Once the cells and the substrate have been primed the next step is to bond the cells to the
substrate. TECSTAR, as well as most other solar cell manufacturers, use a silicone
called CV 2568 or some other variation of that silicone. Before the trip to TECSTAR it
was the intention of ION-F to use a version of the CV 2568 silicone for the satellites,
however TECSTAR felt that a CV 1142 would meet the needs of ION-F. CV 1142 is not
as complicated to use to bond as the CV 2568 because it is a one-part silicone and the CV
2568 is a two-part silicone. CV 1142 has a longer working time, about four hours,
compared to an hour with CV 2568. Marshall Cannedy pointed out the problem with CV
1142 because its bond line doesn’t hold as well as the bond line of the CV 2568. As
such, the CV 2568 will be utilized as was originally planned for the bonding of the solar
arrays to the substrate.

To prepare the CV 2568 for bonding the two-part mixture needs to be mixed together in
the proper ratio. Once this has been done it is necessary to place the mixture under a
vacuum to remove all of the bubbles from the bond material that were created when the
two parts were combined together. After this has been finished it is necessary to begin
bonding, as there will not be a lot of time to waste.

In order to do the bonding the substrate needs to be prepared before this has happened.
The name of the method used to apply the bond material is called skiving. This step of
the process has been changed from what was outlined before. It was determined that it
would be too difficult and time consuming to use Kapton tape to form the 70% area grid
for the silicone bonding. Because of the complexity of outlining a Kapton gird, the idea
was envisioned to do something like what is done with solder paste and a stencil. A 6-
mil stainless steel sheet is used for the grid. This stainless steel sheet has holes punched
out in it where the silicone will be placed on the substrate. The underlying idea is to have
stencils with the proper sized and position holes to lay the silicone on the stencil and then
squeegee the silicone into the holes. The stencil is then removed from the substrate and
the solar cells are aligned and placed on the silicone to be bonded.

The following information has been outdated, however to show the evolution of the
design this information concerning the Kapton grid is included. Note that a stainless steel
stencil is being used instead. To set up the substrate for skiving there must be a grid of
Kapton tape affixed to the substrate in a pattern such that the proper amount of adhesive
will be on the
                                                                    (Document number)**



                              Figure 3.11 Bond                            Size of Solar
                              Guide                                       Cell




                                        Kapton Silicone Guide, 70% of Cell
                                        Area.

substrate in the proper positions. Normally the size of the patch of silicone per solar cell
should be 70% of the area of the solar cell that would be put onto it. A bond line of 6
mils is desired; so three stripes of 2-mil Kapton can be placed on top of each other to
allow for 6 mils of clearance above the surface of the substrate. This grid is depicted in
the figure below. A full-sized drawing of this grid pattern should be plotted out for
referencing the necessary layout of the Kapton tape. As had been stated earlier this
method of using Kapton tape for a guide has been replaced by the use of a stencil.

Place enough of the silicone within each square of the grid necessary to evenly fill the
squares with silicone. Here a squeegee will be used rather than a skiving tool to get the
silicone to the proper level. This process with a squeegee and a stencil will be much like
what is done with solder paste for surface mount PCB assembly. A skiving tool will be
used to evenly distribute the silicone so that it is exactly level with the top of the Kapton
tape grids. A picture of a possible skiving tool is included below. This tool does not
have to be done in exactly this manner; it merely needs to be something with which it is
possible to pull the adhesive across the square of the grid. It might be helpful to have a
tool that will bridge the gap of the grid in the small way, but not in the long way. This
tool can be made out of aluminum or even some sort of plastic.
                                                                      (Document number)**




                                     Figure 3.12 Skiving
                                     tool
Once the adhesive has been laid in the proper fashion the next step is to remove the
Kapton grid from the surface of the substrate. The solar cells should now be placed on the
substrate in its final position. The solar cells are still affixed to the Mylar plot of the solar
cell layout by the Kapton tape that makes contact with the tops of the solar cells through
the holes that were punched out of the Mylar. It will be possible to maneuver the solar
cells to their proper position because the bottoms of the cells will be exposed. The edges
of the substrate should be bordered with acrylic (or some other plastic) at this point in
time so that under vacuum conditions the substrate will not receive any undue stresses.
The following figure depicts the set-up for the vacuum bag that should be used to
uniformly bond the cells. The desired pressure for the bond is 12 inches of water, which
should be held for 18 hours. A bit of the adhesive should be cured in a tin cup alongside
the panel assembly, indicating the relative characteristics of the assembly bond. This will
complete the process necessary to assemble the solar panels.




                                                                                      Vacuum Bag


                                                                                      Hard Back

                                                                                      Mylar

                                                                                      Solar Cells

                                                                                      CV 2568
                                                                                      Kapton

                                                                                      CV 2289

                                                                                      Aluminum Panel
                                                                   (Document number)**


1.2.5 Wiring Specifications
The wiring should be either 24-gauge or 22-gauge wire. 22-gauge is probably generous,
because the 24 gauge wires should be able to work for our purposes. It is possible to get
away with 26-gauge wire for an individual string going from four cells to the rest of the
string. TECSTAR used Raychem wire, the part number is M22-759-44A/22 . This
normally comes in a white jacket, but it can come in any color, such as black and red for
positive and negative voltage wiring. The wiring can be lain along the front of the panel,
if that is necessary. The easiest way to run the wiring through the panel is to drill holes.
These holes should have a telfon ring to have the wire rub against the ring rather than
against the aluminum of the panel. The wires should be crimped every four inches and
then affixed to the panel between the crimps by a dot of adhesive. You can use a Cotton
swab for the model to crimp the wire around it. The following figure contains an
example of the necessary wiring configurations. The 4-inch spaced crimp is necessary
because of expansion issues due to changes of temperature that the wire will experience.




1.2.6 Inspection
Inspection is a method of verification consisting of investigation, without the use of
special laboratory appliances or procedures, to determine compliance with requirements.
Inspection is generally non-destructive and includes (but is not limited to) visual
examination, manipulation, gauging, and measurement. There isn’t a lot of inspection
that can be done with electrical systems because of the necessity of lab equipment to
measure electrical systems.

1.2.7 Demonstration
Demonstration is a method of verification that is limited to readily observable functional
operation to determine compliance with requirements. This method shall not require the
use of special equipment or sophisticated instrumentation. Documentation is the major
                                                                     (Document number)**

means of fulfilling this section of verification. Data sheets for all parts are being
acquired.


1.2.8 Test
Test is a method of verification that employs technical means, including (but not limited
to) the evaluation of functional characteristics by use of special equipment or
instrumentation, simulation techniques, and the application of established principles and
procedures to determine compliance with requirements. LabVIEW has been and will
continue to be used to control testing. There will be LabVIEW testing done on a solar
array with the power system and the battery to demonstrate that the entire power system
functions as required as a system.

For part of the functional testing it must be determined whether the Kapton has
sufficiently bonded to the substrate. To complete this test NASA recommends placing
the material to be tested in a thermal vacuum chamber for 10 cycles (NASA-STD-7002).
The temperature extremes for the cycles should be ten degrees lower than the lowest
temperature that the material will see and 10 degrees higher than the highest temperature
that the material will see. From thermal analysis calculations done at USU the
temperature extremes for the solar arrays should be between –40 degrees Celsius and +50
degrees Celsius.


There are several tests that should be run for the solar cell assembly and then to test the
capability of the solar cells. The tests that should be run for the solar cell assembly
process should include the following. There is going to be a couple of test runs with CV
2568 and the bond process to ensure that the proper way of bonding has been established.
These tests could be done initially with aluminum plates rather than solar cells to be
certain that the bond process is comfortable. ION-F is in the process of talking with
TECSTAR about the possibilities of getting mechanical cells for test purposes. Soldering
processes need to be outlined and extensively tested. It would be conceivable useful to
do the largest array with plain aluminum plates rather than solar cells to start with.


To test that the solar cells are functioning well there needs to be some sort of light source.
TECSTAR felt that it would be easier and more accurate to go outdoors to use the sun as
a light source rather than using special lighting. TECSTAR and TRW both use a Xenon
light source with filters to simulate the light of the sun. A normal halogen lamp will not
work well because the frequencies of light are not the same as sunlight. The best time to
test the solar cells is between 11:00 a.m. and 1:00 p.m. Because of the angle of the earth
during wintertime the solar cells will not see one sun, but it is possible to measure the
amount of sunlight and it should be at an intensity of .7 suns.


Testing will be an important step in solar cell manufacturing to ensure that a viable
method has been chosen to mount the solar cells. This is the most delicate part of the
design and development left to accomplish. The system will be tested in vacuum
                                                                 (Document number)**

conditions with lights to determine how well the solar cells supply the necessary energy
to the satellite. There is other testing that is yet to be done, this is still TBD.
                                                          (Document number)**




2      Post CDR / IDR Schedule
                                           Table 1.5
Task                                                   Manpower
                                 Estimated Time
Design Finalization              2 weeks               Two people
Board Layout                     2 weeks               Two people
Prototype Construction           2 weeks               Two People
Prototype Testing                4 weeks               Two People
Design Modifications             3 weeks               Two People
Construction of Flight Unit      6 weeks               Two People

				
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