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					         HokieSat
Critical Design Document



     Revision VT ICD C-1
   Issue Date: July 21, 2000




 Virginia Polytechnic Institute and State University
 Department of Aerospace and Ocean Engineering
                215 Randolph Hall
            Blacksburg, Virginia 24061
1.0             Signature Page
    Name                 Signature            Subsystem             Date
   Chris Hall                            Principal Investigator

 Adam Harvey                                    Systems

 Craig Stevens                                 Structures

   Katie Hale                                  Thermal

   Dan Sable                                    Power

 Bryce Bolton                                  Computer

Christian Hearn                            Communications

Kristin Makovec                        Attitude Determination and
                                                 Control
Chris Karlgaard                                Propulsion

                                                Science




VT-ICD-C-1 Issue Date: 21 July, 2000                                       1
2.0          History of Revisions
  Revision       Date                  Comments
                                       Initial input from subsystem leads.
  VT-ICD A-1     25 February 2000
                                       Compiled by Adam C. Harvey.
                                       Includes changes from ICD Review on 4
  VT-ICD A-2     22 March 2000         March 2000.
                                       Compiled by Adam C. Harvey.
                                       Includes changes from ICD Review on 23
  VT-ICD A-3     30 March 2000         March 2000.
                                       Compiled by Adam C. Harvey.
                                       Configuration Freeze
  VT-ICD B-1     19 June 2000
                                       Compiled by Adam C. Harvey.
                                       Updates from Power Subsystem.
  VT-ICD B-2     26 June, 2000
                                       Adam C. Harvey
                                       Includes changes from Andrew Turner.
  VT-ICD B-2b    12 July, 2000
                                       Adam C. Harvey
                                       Includes updates for HokieSat CDR.
  VT-ICD C-1     21 July, 2000         Includes Thermal Subsystem section.
                                       Adam C. Harvey




VT-ICD-C-1 Issue Date: 21 July, 2000                                            2
3.0                Table of Contents

1.0 SIGNATURE PAGE ...................................................................................... 1

2.0 HISTORY OF REVISIONS ............................................................................ 2

3.0 TABLE OF CONTENTS ................................................................................ 3

4.0 LIST OF ACRONYMS AND DEFINITIONS ................................................... 6

5.0 LIST OF TABLES AND FIGURES ................................................................ 7

6.0 DOCUMENT PURPOSE................................................................................ 8

7.0 MISSION AND SYSTEM OVERVIEW ........................................................... 9

8.0 STRUCTURE SUBSYSTEMS ..................................................................... 10

  8.1 Subsystem Overview ........................................................................................... 10
  8.1.1   Description .................................................................................................... 10
  8.1.2   Operational States ......................................................................................... 19
  8.1.3   Interfaces With Other Subsystems ................................................................ 19
  8.1.4   Components .................................................................................................. 20
  8.1.5   Mass Budget.................................................................................................. 20
  8.1.6   Power Budget ................................................................................................ 20
  8.1.7   Subsystem Status .......................................................................................... 20

  8.2 Component Overview ......................................................................................... 20
  8.2.1  Side Panels .................................................................................................... 20
  8.2.2  Top and Bottom Panels ................................................................................. 21
  8.2.3  Top, Side, and Bottom Brackets ................................................................... 21
  8.2.4  Fasteners ....................................................................................................... 22
  8.2.5  Lightband ...................................................................................................... 22

9.0 COMPUTER SUBSYSTEM ......................................................................... 23

  9.1 System Overview ................................................................................................. 23
  9.1.1   Power Requirements ..................................................................................... 23
  9.1.2   Mechanical Interface ..................................................................................... 23



VT-ICD-C-1 Issue Date: 21 July, 2000                                                                                       3
  9.1.3          Thermal Requirements .................................................................................. 23
  9.1.4          I/O Card ........................................................................................................ 24
  9.1.5          Signal Definitions ......................................................................................... 24

10.0     POWER SUBSYSTEM ............................................................................ 27

  10.1 System Overview ................................................................................................. 27
  10.1.1   Description .................................................................................................... 27
  10.1.2   Block Diagram .............................................................................................. 27
  10.1.3   Components .................................................................................................. 27

  10.2 Component Overview ......................................................................................... 28
  10.2.1    Batteries ........................................................................................................ 28
  Solar Cells ..................................................................................................................... 28
  10.2.3    Power Requirements ..................................................................................... 29
  10.2.4    Thermal Requirements .................................................................................. 29

11.0     COMMUNICATIONS SUBSYSTEM ........................................................ 30

  11.1 Subsystem Overview ........................................................................................... 30
  11.1.1   Subsystem Description.................................................................................. 30
  11.1.2   Operations and States .................................................................................... 30
  11.1.3   Interfaces ....................................................................................................... 31
  11.1.4   System Components and Properties ............................................................. 31
  11.1.5   Power Budget ................................................................................................ 31
  11.1.6   Status of Hardware ........................................................................................ 31

  11.2 Component Overview ......................................................................................... 32
  11.2.1  L-3 ST-802-HS S-Band Transmitter ............................................................. 32
  11.2.2  Hamtronics R451 UHF Receiver .................................................................. 32
  11.2.3  APL Crosslink and GPS Hardware ............................................................... 32
  11.2.4  Downlink Patch Antenna .............................................................................. 32
  11.2.5  Uplink Receiving Loop Antenna .................................................................. 32

12.0     ATTITUDE DETERMINATION AND CONTROL SUBSYSTEM .............. 35

  12.1 System Overview ................................................................................................. 35
  12.1.1   Description .................................................................................................... 35
  12.1.2   System Operation .......................................................................................... 37
  12.1.3   Interfaces ....................................................................................................... 38
  12.1.4   Components .................................................................................................. 39
  12.1.5   Mass Budget.................................................................................................. 40
  12.1.6   Power Budget ................................................................................................ 40
  12.1.7   Status ............................................................................................................. 40

  12.2 Component Overview ......................................................................................... 41


VT-ICD-C-1 Issue Date: 21 July, 2000                                                                                                4
  12.2.1        Earth Sensor .................................................................................................. 41
  12.2.2        Magnetometer ............................................................................................... 43
  12.2.3        Torque Coils.................................................................................................. 45
  12.2.4        Rate Gyros .................................................................................................... 49

13.0     PROPULSION SUBSYSTEM .................................................................. 52

  13.1 System Overview ................................................................................................. 52
  13.1.1   Description .................................................................................................... 52
  13.1.2   System Operation .......................................................................................... 54
  13.1.3   Interfaces ....................................................................................................... 54
  13.1.4   Components .................................................................................................. 56
  13.1.5   Mass Budget.................................................................................................. 57
  13.1.6   Power Budget ................................................................................................ 57
  13.1.7   Status of System ............................................................................................ 57

14.0     THERMAL SUBSYSTEM ........................................................................ 58

  14.1 Subsystem Overview ........................................................................................... 58
  14.1.1   Description .................................................................................................... 58
  14.1.2   Operational States ......................................................................................... 58
  14.1.3   Analysis Description and Results ................................................................. 59
  14.1.4   Interfaces ....................................................................................................... 60
  14.1.5   Components .................................................................................................. 61
  14.1.6   Power Budget ................................................................................................ 61
  14.1.7   Status ............................................................................................................. 61

  14.2 Component Overview ......................................................................................... 61
  14.2.1  Thermistors ................................................................................................... 61
  14.2.2  Multi-Layer Insulation (MLI) ....................................................................... 61
  14.2.3  White Paint.................................................................................................... 62




VT-ICD-C-1 Issue Date: 21 July, 2000                                                                                            5
4.0             List of Acronyms and Definitions
3CS                 Three Corner Sat
ADCS                Attitude and Determination Control Subsystem
AFRL                Air Force Research Laboratory
APL                 Applied Physics Laboratory
BIST                Built In Self-Test for computer subsystem
CDMA                Code Division Multiple Access
CPU                 Central Processing Unit
DI                  Discharge Initiation
DOD                 Depth of Discharge
FOV                 Field of View for cameras
FPGA                Floating Point Gate Array
FSK                 Frequency Shift Keying
GPS                 Global Positioning System
GSFC                Goddard Space Flight Center
HH                  Hitchhiker
HFC                 HokieSat Flight Computer
ION-F               Ionospheric Observation Nanosatellite Formation
JHU                 Johns Hopkins University
MLI                 Multiple Layered Insulation
MSDS                Multiple Satellite Deployment System
Nanosatellite       One of three ION-F satellites
OS                  Operating System
Payload             Two satellite stacks onboard the MSDS platform
PCM                 Pulsed Code Modulation
PPT                 Pulsed Plasma Thruster
PPU                 Power Processing Unit
RAM                 Random Access Memory
SHELS               Shuttle Hitchhiker Ejection Launch System
SRAM                Static RAM
Stack               The ION-F stack of three satellites
UHF                 Ultra High Frequency
UNP                 University Nanosat Program
USU                 Utah State University
UW                  University of Washington
VT                  Virginia Tech
VT-ISMM             Virginia Tech Ionospheric Scintillation Measurement Mission




VT-ICD-C-1 Issue Date: 21 July, 2000                                              6
5.0                 List of Tables and Figures
Table 8.1: External Layout, Side 1 ................................................................................... 11
Table 8.2: External Layout, Side 2 .................................................................................. 11
Table 8.3: External Layout, Side 3 .................................................................................. 11
Table 8.4: External Layout, Side 4 .................................................................................. 12
Table 8.5: External Layout, Side5 ................................................................................... 12
Table 8.6: External Layout, Side 6 .................................................................................. 12
Table 8.7: External Layout, Zenith .................................................................................. 13
Table 8.8: External Layout, Nadir ................................................................................... 13
Table 8.9: Internal Layout, All Sides ............................................................................... 14
Table 8.10: External Layout, Side ................................................................................... 20
Table 9.1: Thermal Requirements of Computer System ................................................. 23
Table 9.2: I/O Board Signal Definitions .......................................................................... 25
Table 9.3: Power Board Signal Definitions ..................................................................... 26
Table 10.1: Power Requirements ..................................................................................... 29
Table 11.1: Subsystem Components ................................................................................ 31
Table 14.1: Subsystem Component Temperature Ranges ............................................... 58
Table 14.2: Material Characteristics ................................................................................ 59
Table 14.3: Orbit Parameters Used in Analysis ............................................................... 59
Table 14.4: Analysis Results............................................................................................ 60

Figure 8.1 Isometric view of the external configuration of HokieSat ............................ 10
Figure 8.2 Front view of the side panel external configuration ..................................... 12
Figure 8.3 Nadir and zenith external configurations of the spacecraft .......................... 13
Figure 8.4 Internal configuration of HokieSat (2 isometric views) .............................. 14
Figure 8.5 Isometric view of the spacecraft bus. ............................................................ 15
Figure 8.6 Isometric views of the isogrid side and end plates ....................................... 16
Figure 8.7 Interior views of the side panels. .................................................................. 17
Figure 8.8 Interior views of the nadir and zenith panels. ............................................... 18
Figure 8.9 Top view of wall scabs ................................................................................. 19
Figure 8.10 Isometric view of Lightband separation system. .......................................... 22
Figure 9.1 Board Positions in Relation to the Backplane............................................... 23
Figure 9.2 I/O Card Overcurrent and Power Circuit Overview ..................................... 24
Figure 10.1 Power Subsystem Block Diagram ................................................................ 27
Figure 10.2 Solar Cell Dimensions .................................................................................. 28
Figure 11.1 Configuration of Loop Antenna .................................................................... 33
Figure 11.2 Dimensions of Loop Antenna and Mounting Brackets (inches) .................. 33
Figure 12.1 Camera Placement ........................................................................................ 35
Figure 12.2 Camera Field of View ................................................................................... 36
Figure 12.3 Location of Torque Coils and Cameras ........................................................ 37
Figure 12.4 Mount on side panel 1 ................................................................................... 41
Figure 12.5 Mount on side panel 3 ................................................................................... 42


VT-ICD-C-1 Issue Date: 21 July, 2000                                                                                     7
Figure 12.6 Mount on side panel 5 ................................................................................... 42
Figure 12.8 Magnetometer Housing ................................................................................. 45
Figure 12.9 Mount of Torque Coils.................................................................................. 46
Figure 12.10 Interface of Torque Coil and Computer .................................................... 47
Figure 12.11 Layout of the A3966SA with the truth table. .............................................. 47
Figure 12.12 Functional Block Diagram of the A3966SA ............................................... 48
Figure 12.13 Interface of Gyros and Computer ............................................................. 50
Figure 13.1 Position of Thrusters ..................................................................................... 52
Figure 13.2 Pulsed Plasma Thruster ................................................................................. 53
Figure 13.3 Thruster Housing ........................................................................................... 54
Figure 13.5 Thruster and Side Structure Interface ............................................................ 56




6.0                 Document Purpose
The purpose of this document is to identify the interfaces of all components onboard
HokieSat. This is done by describing each subsystem, and all components making up
that system, then listing how each component interfaces with other systems.




VT-ICD-C-1 Issue Date: 21 July, 2000                                                                                     8
7.0            Mission and System Overview
The Air Force Research Laboratory’s (AFRL’s) “TechSat-21” program was developed to
investigate the practicality of using small, distributed spacecraft systems, to perform the
missions of larger, single platforms. The University Nanosatellite Program (UNP) is a
subset of this program. The purpose of the UNP is to fund universities to help explore
and implement the technologies of small satellites; 10 schools, including Virginia Tech,
are receiving funding for this program.

At the time that funding was awarded, Virginia Tech had proposed a single satellite
investigation, which is called the Virginia Tech Ionospheric Scintillation Measurement
Mission, or VT-ISMM. More commonly known as HokieSat, the design was quickly
integrated into a team with Utah State University and the University of Washington due
to complimentary scientific interests. HokieSat, USUSat, and Dawgstar thus formed the
Ionospheric Observation Nanosatellite Formation, ION-F. The primary ION-F missions
include several methods of measuring local ionospheric properties, formation flying,
distributed flight and ground control, related technology demonstrations, and high student
involvement. HokieSat’s mission refines these goals even further.

The ION-F formation flying mission is tied closely to NASA-Goddard Space Flight
Center (GSFC). Many formation flying algorithms have been developed at GSFC, but
have not yet been flown. Earth Observer 1 (EO-1) will be the first satellite to implement
these algorithms as it flies with Landsat 7. As ION-F has three satellites each with
differing propulsive capabilities, it will be able to demonstrate more involved formation
flying routines.

On a slightly more global scale, the ION-F team is paired with another UNP funded
project, 3-Corner Sat (3CS). 3CS is composed of satellites from Arizona State
University, New Mexico State, and University of Colorado at Boulder. The two satellite
stacks will be launched together from the Space Shuttle off AFRL’s Multi-Satellite
Deployment System (MSDS). The MSDS is under concurrent development with the
university nanosatellites, and serves to support the TechSat-21 program.

As with the missions described previously, the HokieSat systems must be considered
from several scales. When integrated into the Shuttle, the payload system includes not
only the ION-F stack, but also the MSDS and the 3CS stack. Perhaps the dominant level
in design and testing is the ION-F stack alone; HokieSat is the lowest of the three
satellites in the stack, interfacing between Dawgstar and the MSDS. Finally, the single
satellite design of HokieSat is most central, and is the focus of this document. However,
it should be considered that many HokieSat subsystems are common or complimentary to
the rest of ION-F, and as such the stack level system feeds back into the single satellite
design.



VT-ICD-C-1 Issue Date: 21 July, 2000                                                      9
8.0            Structure Subsystems

   8.1         Subsystem Overview

      8.1.1    Description
The HokieSat structure is designed using three ideas: 1) simple and easily fabricated
design, 2) sized to fit and interface with all components and maximize solar cell area
while staying within the volume and mass constraints, 3) able to withstand the loads
during launch and support the ION-F stack.

The structure is designed in the shape of a hexagonal cylinder (see Figure 8.1).




         Figure 8.1     Isometric view of the external configuration of HokieSat

The external configuration is designed so the side panels overlap the top and bottom
separation systems by 1.0”. This allows a greater surface area to attach components (see
Figure 8.2), while minimizing mass and complying with all stay-out zone requirements.
The solar cells are configured in strings of 12, with a total number of 156 cells. The side
panels interface with components such as solar cells, thruster nozzles, cross-link
antennas, and cameras. The top (zenith) of the spacecraft has 12 solar cells and the GPS
patch antenna (see Figure 8.3). The top solar cell configuration is raised 0.75” to
minimize shading from the Lightband. The bottom (nadir) face of the spacecraft supports
the up-link and downlink antennas.


VT-ICD-C-1 Issue Date: 21 July, 2000                                                     10
Table 8.1: External Layout, Side 1
      Side          Component
  1              Solar Cell String 1A
  1              Solar Cell String 1B
  1                    PPT A1
  1                   Camera 1




Table 8.2: External Layout, Side 2
      Side          Component
  2              Solar Cell String 2A
  2              Solar Cell String 2B
  2                    PPT A2
  2                  C/L Patch 1




Table 8.3: External Layout, Side 3
          Side      Component
      3          Solar Cell String 3A
      3          Solar Cell String 3B
      3               Camera 2
      3                PPT B3




VT-ICD-C-1 Issue Date: 21 July, 2000    11
Table 8.4: External Layout, Side 4
      Side       Component
    4         Solar Cell String 4A
    4         Solar Cell String 4B
    4               PPT B4
    4             C/L Patch 2




Table 8.5: External Layout, Side5
      Side        Component
    5         Solar Cell String 5A
    5          Solar Cell String 5B
    5               Camera 3




Table 8.6: External Layout, Side 6
         Side      Component
     6          Solar Cell String 6A
     6          Solar Cell String 6B
     6              C/L Patch 3




            Figure 8.2     Front view of the side panel external configuration




VT-ICD-C-1 Issue Date: 21 July, 2000                                             12
Table 8.7: External Layout, Zenith
    Side         Component
  Zenith       Solar Cell String Z
  Zenith         GPS Antenna
  Zenith         Magnetometer
  Zenith           Lightband




Table 8.8: External Layout, Nadir

  Side         Component
 Nadir         U/L Antenna
 Nadir         D/L Antenna
 Nadir           Starsys




        Figure 8.3    Nadir and zenith external configurations of the spacecraft




VT-ICD-C-1 Issue Date: 21 July, 2000                                               13
The internal configuration is not volume constrained at this time (see Figure 8.4). The
components are arranged according to specifications and in an attempt to optimize the
structural and thermal properties. If there are any problems with the current
configuration layout, please specify the details in the mass properties list.




          Figure 8.4     Internal configuration of HokieSat (2 isometric views)



Table 8.9: Internal Layout, All Sides

       Location                      Component
 Nadir                   Electronics Enclosure
 Nadir                   Battery Enclosure (Not Pictured)
 Zenith                  Hexagonal Torque Coil
 Side 1                  Earth Sensor
 Side 1                  PPT 1A (Not Pictured)
 Side 2                  PPT 2A (Not Pictured)
 Side 3                  Earth Sensor
 Side 3                  PPT B1 (Not Pictured)


VT-ICD-C-1 Issue Date: 21 July, 2000                                                      14
 Side 3                 Rectangular Torque Coil 1
 Side 4                 Rectangular Torque Coil 1
 Side 4                 PPT B2 (Not Pictured)
 Side 4                 Rectangular Torque Coil 2
 Side 5                 Earth Sensor
 Side 6                 Rectangular Torque Coil 2

The bus is fabricated out of aluminum 6061-T4, which is readily available through the
Virginia Tech AOE shop (see Figure 8.5 and 8.6). The bus is assembled out of eight
isogrid plates. There are six identical side panels measuring 13.725” in height(see
Figures 8.6a and 8.6b). The top and bottom overhangs measure 1.25” in height and
protrude up and around the top and bottom plates and separation systems. The “actual”
stack height measures 11.725” between the separation system interfaces. The top and
bottom panels are identical 18.00” hexagonal plates with a thickness of 0.25”. The
isogrid is designed with 0.025” skin and nodes that have been spaced 2” apart. Each
node will have a hole drilled that measures 0.25” in diameter. A smaller hole may be
drilled if any component requires a different sized hole for mounting. ANY CHANGES
SHOULD BE MADE AVAILABLE ASAP IN THE MASS PROPERTIES TABLE.




                  Figure 8.5    Isometric view of the spacecraft bus.




VT-ICD-C-1 Issue Date: 21 July, 2000                                                15
            Figure 8.6      Isometric views of the isogrid side and end plates




           Side          Component
       1                   PPT A1
       1                  Camera 1



                                                   13.725”                            10.425”




                                                                         9.0”




VT-ICD-C-1 Issue Date: 21 July, 2000                                             16
                         Side         Component
                     2                  PPT A2




                                    Side         Component
                                3                 Camera 2
                                3                  PPT B3



                         Side         Component
                     4                  PPT B4




                                    Side         Component
                                5                 Camera 3




                         Side          Component
                     6                 Data Test Port




                  Figure 8.7        Interior views of the side panels.




VT-ICD-C-1 Issue Date: 21 July, 2000                                     17
   Side       Component
 Zenith       Magnetometer




                                                    18.00”




  Side        Component
 Nadir          None




             Figure 8.8   Interior views of the nadir and zenith panels.




VT-ICD-C-1 Issue Date: 21 July, 2000                                       18
The side plates are designed to connect using wall brackets (see Figure 8.7), which are
placed at all internal corners of the spacecraft. ANY components that require mounting
in the “stay out” areas should list the requirements specifically in the mass properties list
ASAP.




                           Figure 8.9     Top view of wall scabs


      8.1.2    Operational States
Launch state - the most stress filled state during the spacecraft lifetime.
On-orbit state - where large temperature variations will be experienced.

The structures subsystem is designed to support the payload according to the NASA
GSFC load and stress requirements. The same material (aluminum 7075-T6) is used
throughout the spacecraft to alleviate any expansion/contraction differences that will
stress the connections due to on-orbit temperature changes.

      8.1.3    Interfaces With Other Subsystems
All components are mounted to the bus using NASA GSFC approved fasteners. All
subsystems should place fasteners every two inches to allow for easy mounting to the bus
(see Figure 2b). The components should use a relatively simple mounting scheme that
uses at least four fasteners (GSFC requirement) to connect to the bus. Any components
that require special mounting configurations should list the requirements in the mass
properties table.




VT-ICD-C-1 Issue Date: 21 July, 2000                                                       19
      8.1.4     Components

Table 8.10: External Layout, Side

 Component             QTY
 Side Panel             6
 Top Panel              1
 Bottom Panel           1
 Wall Scabs             6
 Fasteners             TBD

      8.1.5     Mass Budget
The total mass of the structure subsystem is presently 15.83 lbm, or 7.194 kg.

      8.1.6     Power Budget
This subsystem requires no power.

      8.1.7     Subsystem Status
The structures subsystem is currently undergoing finite element analysis. This will
determine if any modifications need to be made in order to satisfy the requirements set
forth by GSFC. The main concern of the subsystem is to build a structure that will satisfy
the lowest mode natural frequency requirements (AFRL/GSFC requirements).
Optimization is ongoing and updates will be made periodically.

   8.2          Component Overview

      8.2.1     Side Panels
Panels are made of 7075-T6 Aluminum.

 Component QTY              W                 H
                         cm, (in)         cm, (in)
 Side Panel        6   23.1, (9.085)   34.9, (13.725)

         8.2.1.1       Mass Budget of Side Panels
 Component QTY              Mass                Mass
                        Each g, (lbm)       Total g, (lbm)
 Side Panel        6   521.5, (1.1497)     3129.78, (6.900)




VT-ICD-C-1 Issue Date: 21 July, 2000                                                   20
       8.2.1.2           Status of Side Panels
The aluminum is presently in the Virginia Tech AOE shop and some plates may need
ordering. The isogrid must be milled out either on-campus or in Roanoke. The labor cost
involved with these components will be the milling out process cost (approximately
$20.00 per hour with one side panel taking approximately 30 minutes to mill) and the
time for students to build the structure.

      8.2.2    Top and Bottom Panels
  Component        QTY       DIA
                              (in)
  Top Panel          1       18.00
 Bottom Panel        1       18.00

       8.2.2.1           Mass Budget of Top and Bottom Panels
  Component        QTY          Mass                Mass
                             Each g, (lbm)      Total g, (lbm)
  Top Panel          1      1133.5, (2.499)    1133.5, (2.499)
 Bottom Panel        1      1133.5, (2.499)    1133.5, (2.499)

       8.2.2.2           Status of Top and Bottom Panels
The aluminum is presently in the Virginia Tech AOE shop and some plates may need
ordering. The isogrid is milled out on-campus in Whittemore Hall. The labor cost
involved with these components is the milling out process cost (approximately $20.00 per
hour with one side panel taking approximately 30 minutes to mill) and the time for
students to build the structure.

      8.2.3    Top, Side, and Bottom Brackets
The brackets are 0.25” aluminum 7075-T6, which is also available from the Virginia
Tech AOE shop.

These scabs are connected to the plates with bolts that are approved by NASA Goddard
Space Flight Center (GSFC).

The brackets will line all internal corners of the spacecraft such that the spacecraft
components cannot be mounted in these areas of the spacecraft.

       8.2.3.1           Mass Budget of Brackets
  Component       QTY          Mass                Mass
                           Each g, (lbm)       Total g, (lbm)
 Wall Brackets      6      51.71, (0.114)     310.71, (0.685)



VT-ICD-C-1 Issue Date: 21 July, 2000                                                     21
       8.2.3.2        Status of Brackets
The aluminum is readily available through the Virginia Tech AOE shop.

      8.2.4    Fasteners
The structural fasteners are ordered and awaiting arrival. The fasteners are available
from the GSFC web page: http://lmd.gsfc.nasa.gov/fasteners/. Once any other needed
fasteners have been selected, there is a short lead-time to receive the components after
order. The cost of these items depends on the type of fastener that is needed (ranging
from $0.27 to $3.23 per item).

Number 10 fasteners must be used if located in the load path. Otherwise, smaller
fasteners may be used.

      8.2.5    Lightband
Lightband is the separation system that is used in the ION-F constellation. HokieSat
requires the bottom half of Lightband on the zenith face (see Figure). The system is
designed by PSC and purchase is correlated by the University Nanosat program. The
mass of the lower section of Lightband is 0.799 kg, or 1.7578 lbm. More detailed
specifications of the Lightband separation system may be found at the following URL:
http://www.aa.washington.edu/research/nanosat/docs/docs.htm




              Figure 8.10    Isometric view of Lightband separation system.




VT-ICD-C-1 Issue Date: 21 July, 2000                                                       22
                                                    Data Bus Backplane                                                  SPI Bus Backplane
                                                                                         Common Backplane

                                                                                         USUSAT




                                                       Telemetry Board




                                                                                                                                                             Power Board 1

                                                                                                                                                                             Power Board 2
                                                                          Camera Board
                          Gyro Board
           Rate Gyros




                                        CPU Board




                                                                                                            PDP Board
                                                                                               IO Board




                                                                                                                             Spare 2


                                                                                                                                                  Spare 1
9.0                     Computer Subsystem

  9.1                   System Overview
                         Slot 1 Slot 2 Slot 3                            Slot 4               Slot 5      Slot 6           Slot 7                Slot 8     Slot 9
                                                    Data Bus Backplane                                                  SPI Bus Backplane

                                        boards. These are the CPU board, the telemetry
The computer system has three commonCommon Backplane
board, and the I/O board. Figure 9.1 shows various board positions in relation to the
                                        Dawgstar
Backplane.




                                                                                                                             Torque Coil Board
                                                       Telemetry Board




                                                                                                                                                             Power Board 1

                                                                                                                                                                             Power Board 2
                                                                          Camera Board
                          Gyro Board
           Rate Gyros




                                        CPU Board




                                                                                                            PDP Board
                                                                                               IO Board




                                                                                                                                                  Spare 1
                         Slot 1        Slot 2        Slot 3              Slot 4               Slot 5      Slot 6           Slot 7                Slot 8     Slot 9
                                                    Data Bus Backplane                                                  SPI Bus Backplane
                                                                                         Common Backplane

                                                                                         HokieSat
                        Figure 9.1                    Board Positions in Relation to the Backplane

      9.1.1             Power Requirements
Peak power consumption: Less than 3W
Power Interface: 3.3  0.3V, 5  0.25V, AGND, GND

      9.1.2             Mechanical Interface

The electronics enclosure will measure 9” x 6” x 4”.


      9.1.3             Thermal Requirements

Table 9.1: Thermal Requirements of Computer System
Operating Temperature     -40 to +85C
Survival Temperature      -55 to +125 C
Cooling Mechanism         Conduction Cooled




VT-ICD-C-1 Issue Date: 21 July, 2000                                                                                                                                                         23
      9.1.4     I/O Card




              Figure 9.2   I/O Card Overcurrent and Power Circuit Overview

This circuit in Figure 9.2 allows for passive monitoring of the I/O board current level.
Power switching capability is controlled via a single 3.3V backplane signal called
IO_CARD_ON/OFF#. Asserting this signal high applies power to the board via a SEGR-
resistant P-FET switch circuit. The board current is monitored using a single-supply
instrumentation amplifier. A low-value (1-10 Ohm) sensing resistor (Rs) is used to view
the circuit current as a voltage. Rg sets the instrumentation amplifier gain. We can scale
Rg so that the typical current through the IO card is viewed in the 0-4V range at the
instrumentation amplifier output. The next stage, a comparator, is used to sense when the
current is above its nominal range. A rise above 4.5VDC at the positive input of the
comparator will cause a rising clock edge on a D Flip-Flop. The overcurrent condition
will be signaled to the microprocessor via a shared backplane interrupt
OVERCURRENT_INT#. The microcontroller resolves which card caused
OVERCURRENT_INT# by reading Vcurrent monitor, or by reading the value of 'Q' at
the overcurrent D Flip-Flop output. Alternatively, the design could be altered to allow for
multiple interrupt, lines for faster overcurrent resolution.

      9.1.5     Signal Definitions
Table 9.2 shows signal definitions of the I/O board and Table 9.3 shows signal definitions
of the power board.




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Table 9.2: I/O Board Signal Definitions
                     Directi
                     on
                                          Back-   Top Rang
Sub       Signal      wrt IO Signal       plane   pin e     Description
System    Type        board Name          Pin #   Pin #
          3.3V                                          [0- Current Monitor for on
          ANALOG      OUT                 ?       N/A 3.3V] I/O board
          3.3V                                          [0- Overcurrent Interrupt for
          ANALOG      OUT                 ?       N/A 3.3V] on I/O board
          3.3V
          DISCRET                                     [0,3. A/D conversion complete
          E           OUT                 ?       N/A 3V] (I/O board)
          Discrete        DO_2_XLI
GPS/Xlink I/O         OUT NKPWR    N/A            ?    0,5V Power on/off# XLINK
Uplink/Do Discrete        DO_3_TXP
wnlink    I/O         OUT WR       N/A            ?    0,5V Power on/off# transmitter
          Discrete        DO_4_RXP
          I/O         OUT WR       N/A            ?    0,5V Power on/off# receiver
          5V
          DISCRET         DO_DISC_                     [0,5V
IO Board E            OUT 0         N/A           ?    ]     Power on/off PPT board
          Discrete        DO_6_FIRE
PPT       I/O         OUT 1         N/A           ?    5V? Fire 1
          Discrete        DO_7_FIRE
PPT       I/O         OUT 2         N/A           ?    5V? Fire 2
                          DI_PPT_IN
PPT       A/D         IN  1         N/A           ?    0,5? Voltage 1
                          DI_PPT_IN
PPT        A/D        IN  2         N/A           ?    0,5? Voltage 2
           Discrete
Rate Gyro I/O         OUT DO_#_XON N/A            ?    0,5 Power on/off x
           Discrete
Rate Gyro I/O         OUT DO_#_YON N/A            ?    0,5 Power on/off y
           Discrete
Rate Gyro I/O         OUT DO_#_ZON N/A            ?    0,5 Power on/off z
Rate Gyro A/D         IN     DI_           N/A    ?    ?    Measure x
Rate Gyro A/D         IN     DI_           N/A    ?    ?    Measure y
Rate Gyro A/D         IN     DI_           N/A    ?    ?    Measure z
Need to define which discrete IO are thru from CPU and which are generated or read via
logic on IO Board.




VT-ICD-C-1 Issue Date: 21 July, 2000                                                 25
Table 9.3: Power Board Signal Definitions
        Sign wrt              Back- Top Rang
Sub     al   Pwr Signal       plane    pin    e     Description
             boar
System Type d     Name        Pin #    Pin #
Power             V_BATT_
Card A/D IN TEMP              USER_0 N/A 0-5v Battery Temperature
                  V_BUS_V
        A/D IN OLT            USER_1 N/A 0-5v Total Bus Voltage
                  V_BUS_C
        A/D IN URR            USER_2 N/A 0-5v Total Bus Current
                  V_CELL_
        A/D IN SIDE1          USER_3 N/A 0-5v Voltage side 1
                  V_CELL_
        A/D IN SIDE2          USER_4 N/A 0-5v Voltage side 2
                  V_CELL_
        A/D IN SIDE3          USER_5 N/A 0-5v Voltage side 3
                  V_CELL_
        A/D IN SIDE4          USER_6 N/A 0-5v Voltage side 4
                  V_CELL_
        A/D IN SIDE5          USER_7 N/A 0-5v Voltage side 5
                  V_CELL_
        A/D IN SIDE6          USER_8 N/A 0-5v Voltage side 6
                  V_CELL_
        A/D IN SIDE7          USER_9 N/A 0-5v Voltage side 7 (top)
Route signals
from IO board via
user defined IO.
Power outputs pin #s are not defined here. They are defined in the ION-F computer backplane definition, but
should be added here.




VT-ICD-C-1 Issue Date: 21 July, 2000                                                 26
10.0          Power Subsystem

  10.1        System Overview

     10.1.1 Description
   The power system needs to distribute power from the solar arrays to all systems in the
   satellite. The voltage bus will be determined from the solar array configuration,
   voltage, and battery cell voltage. Our voltage bus will range from 16.5V to 22.5 V.
   Note that the battery cells are directly connected to the solar array. A DC to DC
   converter will distribute this voltage among the sub-systems. Voltage limits need to
   be on the realm of standard values in order for the DC to DC converter to be efficient.
   Batteries will be used to provide power during eclipse conditions. These batteries
   will be charged during sun exposure.

     10.1.2 Block Diagram




                    Figure 10.1    Power Subsystem Block Diagram




     10.1.3 Components

    Component                       Purpose
    Voltage regulator circuit       To regulate voltage from solar cells.
    Current regulator circuit       To regulate and remove unwanted current
                                    (Shunt regulator to be placed outside of satellite


VT-ICD-C-1 Issue Date: 21 July, 2000                                                     27
                                     to dissipate heat) Prevents battery from being
                                     over charged.
    Chargeable Batteries             To provide power to satellite during eclipse
    DC to DC converter               To distribute voltages to all sub-systems
                                     (To be built by VPT, Dan Sable)
     UController/Chrg Controller      Determines charge rate of battery


  10.2       Component Overview

     10.2.1 Batteries
Sanyo Cadnica model KR-1400AE
15 cells
1.4 A-hr
Nominal Voltage = 16.8V
45% DOD
Total mass: 465g, (1.0251lb)

     10.2.2 Solar Cells




                           Figure 10.2   Solar Cell Dimensions


Desired:
    Each string should have 12 cells.
    13 strings
    Total number of solar cells = 156

Solar cell characteristics:
    Maximum average weight per 100 cells is 2.20 grams.
    Isc = .35 A
    Nominal current = 2.2 A


VT-ICD-C-1 Issue Date: 21 July, 2000                                                  28
       Avg. S/C load = 1.43 A
       Voc = 2.4 V ≥ 2.4 V x 12 = 28.8 V
       Vsa ~≤ 1.85V
       Vbatt ≤ 22.2V – account for diode drop

By these characteristics, the battery is left with about .77A for charging during 54-minute
periods. The battery can only replenish 0.69 A-hr, so the load should be less then 22 W
(minus converter losses).

       10.2.3 Power Requirements
Table 10.1: Power Requirements
 Sub-System     Voltage            Power             Comment
 Computer       5V and 3V          >3.0W (max)
 A/D                                                 Can it take +15V
 S Band Xmit                       14W               Can it take 15.4-21V
 UHF Rcvr       11-15V             5.4W – 15W        Can it take 15.4-21V
 Torque Coils 3.3 V                0.25W             Can it take 5V
 Cameras        +/- 5V             0.85W  3
 Magnet         +/- 15V            0+.3W             (4A/1 usec pulse)
 Rate Gyro      +/- 5V             1.2W  3
 PPT            28 4V             13W
 Relays         28V                                  Can it take 15.4-21 V
 GPS            5V                 1.4W

Table 10.2: Voltage Distribution from DC-DC Converters
 T. Coil          DVSA283R3S + 3.3V
 Cameras          DVSA2805D  5V
 Magnetom         DVSA2815S        +15V
 Rate Gyro        DVSA2805D  5V
 GPS              DVSA2805S        +5V
 Computer         DVSA283R3S + 3.3V
 HV PS            Direct Bus        16.5 V– 22.5V
 Comm.            Direct Bus        16.5 V– 22.5V
 UHF receiver
 Relays           Direct Bus        16.5 V– 22.5V

       10.2.4 Thermal Requirements
The DC-DC converters have an operating temperature range of –55C to +125C.




VT-ICD-C-1 Issue Date: 21 July, 2000                                                    29
11.0           Communications Subsystem

   11.1        Subsystem Overview

       11.1.1 Subsystem Description
The communication system in HokieSat is comprised of three major links: the crosslink,
the uplink, and the downlink.

The crosslink will operate in the S-Band, and will be implemented using hardware that
will be provided by JHU APL. This hardware includes the GPS receiver, crosslink
transceiver and the antenna, and it will use CDMA.

The GPS hardware will use a patch antenna operating in the L-Band.

The crosslink transceiver will use a patch operating in the S-Band. The required
bandwidth on the crosslink is approximately 100KHz.

The downlink operates in the S-Band. The center frequency lies somewhere in 2200-
2290MHz, and will be determined as soon as frequencies are assigned, or as soon as a
firm decision is made regarding them. The required bandwidth is approximately
200KHz, and the link operates using FSK. This link will be used to transmit mostly
science data. It will use a patch antenna on the nadir face of the spacecraft.

The uplink will operate at approximately 450 MHz, and will use a loop antenna on the
nadir face of the spacecraft. It will require about 100KHz of bandwidth. It will also use
FSK.

       11.1.2 Operations and States
The downlink transmitter will operate when the spacecraft is visible to the earth station.
This will occur for about several minutes in every 90-minute orbit. The command to
downlink the data will come from the computer. Once this command is received the
transmitter receives the baseband data from the buffer, modulates and amplifies, and then
sends it out to the antenna.

       There only two states of operation of this hardware are either on or off.

       The uplink receiver is turned on when data from the GPS and orbit propagator
        show that a ground station pass is about to occur. Its senses and acquires the
        carrier, and then demodulates it to baseband, at which point the data is passed to
        the computer for analysis.



VT-ICD-C-1 Issue Date: 21 July, 2000                                                    30
       The crosslink and GPS hardware operates using and RS-232 interface with the
        computer. At this time, it is assumed that the computer will have direct control
        over the operation of this piece of hardware.

       11.1.3 Interfaces
The crosslink hardware box will have an RS-232 port on it and will interface with the
computer. The computer will then operate it when necessary.

       11.1.4 System Components and Properties
Table 11.1 Subsystem Components
 Component            Description               Power       Voltage    Mass g, (lb)
 L3 ST 802 S          S-Band Transmitter        14W         28V        130 (0.2866)
 Hamtronics R451* UHF Receiver                  1.5W        11-15V     105 (0.23125)
 X-Link/GPS           Hardware from APL         1.4W        5V         750 (1.65)
 Patch Antenna (3)    S-Band Transmitting       NA          NA         30, (0.0661)
                      Antenna
 Loop                 UHF Receiving             NA          NA         10, (0.0220)
                      Antenna
 GPS Patch Antenna GPS L-Band Patch             NA          NA         10, (0.0220)
                      Antenna
 Total (approx.)       Assume 5 W for X-Link    ?                      378, (0.8333)
   *- (11V-15V at 36-100mA)

       11.1.5 Power Budget
Power requirements stated above are for the on state. The receiver stays on at all times,
whereas the transmitter operates when needed. The above total is based upon the
assumption that the crosslink takes 5 W during operation. Based upon that, the Comm
system power requirements vary from about 4.48 W to 24 W.

       11.1.6 Status of Hardware
 Hardware             Current Status
 Transmitter          Has been decided upon
 Receiver             Has been decided upon, and is being modified
                      by USU
 X-Link Hardware      APL is working on that
 Antennas             Currently being designed




VT-ICD-C-1 Issue Date: 21 July, 2000                                                    31
   11.2          Component Overview

      11.2.1 L-3 ST-802-HS S-Band Transmitter
This piece of hardware will interface with power and computer. Connectors are shown
below.

J-1                       MDM-15S
Pin Connectors            1 through 8 Freq. Selection
                          9 – PWR RTn
                          10 – Spare
                          11 – Spare
                          12 – Spare
                          13 - +28V
                          14 – PWR RTN
                          15 – PWR RTN
J-2 RF Output             SMA Female
J-3 Modulation Input      SMA Female

      11.2.2 Hamtronics R451 UHF Receiver
This piece of hardware will interface with power and computer. Pin details are still
unknown.

      11.2.3 APL Crosslink and GPS Hardware
At this time, the only information available about this piece of hardware is that it has an
RS-232 output port to connect with the computer. It is believed at this time that there are
a total of three crosslink antennas. They will be patches measuring approximately 1.7” 
1.7”, and will need to be located on every other side panel.

      11.2.4 Downlink Patch Antenna
This antenna is operated in the S-Band, and will be mounted flush with the nadir face of
the spacecraft. It requires no DC power. It will connect to the S-Band transmitter
through an SMA or SSMA connector.

The antenna will be a patch, approximately 3in square, and will be located on the Nadir
face. This can be placed inside the loop antenna.

      11.2.5 Uplink Receiving Loop Antenna
This antenna operates at 450 MHz, and it must be mounted onto a continuous metallic
surface. It will connect to the UHF receiver through a SMA connector, and requires no
DC power.




VT-ICD-C-1 Issue Date: 21 July, 2000                                                      32
The antenna will be a loop located on the Nadir face, and will need to extend between 1in
and 1.3in from the backplane, as shown in Figure 11.1 and 11.2.




                      Figure 11.1 Configuration of Loop Antenna




       Figure 11.2 Dimensions of Loop Antenna and Mounting Brackets (inches)




VT-ICD-C-1 Issue Date: 21 July, 2000                                                  33
VT-ICD-C-1 Issue Date: 21 July, 2000   34
12.0           Attitude Determination and Control Subsystem

   12.1        System Overview

      12.1.1 Description
The purpose of the Attitude Determination and Control System (ADCS) is to determine
the attitude of HokieSat at a particular point in time, compare the measured attitude with
the desired orientation, and make necessary corrections.

The desired attitude of HokieSat is a 3-axis stabilized orientation, with a certain
hexagonal face designated as “down” in the nadir direction. This orientation must be
obtained so that antennas and propulsive devices will not have to be redundant. With this
position, their orientation of the antennas in the roll and pitch directions relative to the
Earth and the orbit will be constant.

The attitude of the HokieSat is determined through the use of earth sensors mounted on
the sides of the satellite, and by a magnetometer measuring the magnetic field of the
Earth. The rate of change of the attitude is measured with a rate gyro. Running a current
through three copper magnetic torque coils makes attitude corrections.

The earth sensor is composed of three cameras that are used to take pictures of the
Earth’s horizon in order to determine the attitude of HokieSat. The three cameras are
located alternating sides of the satellite (Figure 12.1), and are flush with the outside
surface.




                             Cameras



                             Figure 12.1    Camera Placement




VT-ICD-C-1 Issue Date: 21 July, 2000                                                       35
The field of view of each camera is approximately 67, leaving gaps of 53 in coverage
around the circumference of the satellite. This is depicted below in Figure 12.2.




                          Figure 12.2     Camera Field of View


Analysis shows that during orbit, the Earth’s horizon (neglecting shadowing) appears in
the FOV of all three cameras for almost all conditions. This includes if the satellite is
rotated up to 15 from nadir pointing when the altitude is 380 km, or up to a 20 rotation
at 250 km. This suggests that even when shadowing is considered, the horizon should be
in view of at least one camera at all times. A problem could occur if the position of
HokieSat became such that one camera was pointing directly at the Earth and the others
were pointing out into space. In this instance, no horizon would be detected, but a coarse
nadir vector could be estimated since one camera would be pointing directly at the Earth
and would show brightness, while the others would depict darkness.

The torque coils are composed of copper wire wound in three loops. The three coils are
located on the inside of the satellite near one of the surfaces. One hexagonal loop is
oriented parallel to the top surface of the satellite with an inner radius of 5.686 inches,
and two rectangular shaped coils with inner dimensions of 8.5 in  10 in are placed such
that they are mutually orthogonal, with one parallel to a side face. Figure 12.3 shows the
location of the torque coils and cameras.




VT-ICD-C-1 Issue Date: 21 July, 2000                                                     36
                                Torque Coils




     Cameras
     Mounted
     on Sides




                  Figure 12.3    Location of Torque Coils and Cameras


      12.1.2 System Operation
The ADCS of HokieSat is only turned on when measurements are being taken or attitude
corrections are being made. One reason for this is to save power. Constant attitude
measurements and adjustments should not be necessary. In addition, the magnetometer
and magnetic torque coils cannot be on at the same time because they affect each other’s
readings and output. The magnetometer must also be turned off when the propulsion
system is in use.

Before the magnetometer can be used, a set/reset pulse must initialize it. The
characteristics of the magnetometer device are upset when exposed to a large magnetic
field. HokieSat will be exposed to such a field resulting from the close proximity during
launch of the permanent magnets being used on Utah State University’s satellite. This re-
initialization of the magnetometer is accomplished by running a current pulse through the
system for a short period of time.

The orbital location of HokieSat is known from the GPS system. From this position, it is
possible to determine the percentage of the Earth HokieSat can expect to see, as
illuminated by the sun. The earth horizon sensors are turned on, and determine the
horizon boundaries. The magnetometer then measures the magnetic field and compares
with the known field at that orbital location. The rate gyro measures the rate of change of


VT-ICD-C-1 Issue Date: 21 July, 2000                                                    37
attitude. This data is sent to the computer system. The attitude determination system
then turns off.

A computer program performs analytical calculations and compares the measured attitude
with the desired nadir-facing attitude. If corrections need to be made, the magnetic
torque coils turn on, and run a calculated current through the copper coils. After the
attitude correction has been made, the torque coils turn off, and the system is ready to
begin the process of taking measurements again.

As time progresses, it is necessary to reset the rate gyros occasionally. A drifting occurs
in the measurements, which over time becomes substantial enough to cause errors in
readings. At a determined interval, the magnetometer and earth sensors are used to
calculate the rate of change of attitude, and this is compared with the measurements from
the rate gyros. If a discrepancy occurs, the rate gyros are reset according to these
measurements.

       12.1.3 Interfaces
The main systems with which HokieSat interfaces are thermal, computer, power,
structures, GPS, and propulsion.

        12.1.3.1       Thermal Requirements
All of the components of the ADCS subsystem have a temperature range of -40 to
+85C, with the exception of the CMOS cameras for earth sensing which has a maximum
operating temperature of 60C. This implies that none of the systems located internally
require any heating or cooling in excess of what is done to the satellite in general.

       The earth sensor is located on the outside of the satellite, and therefore the
        cameras are exposed to direct sunlight in addition to shadowing.

        12.1.3.2       Computer Interface
The connections between the earth sensors, magnetometer, and rate gyro have similar
requirements. For these systems, the computer needs to tell the device when it is time to
turn on, based on calculations made in specifically written software.

       The computer also needs to accept data back from these systems to perform
        subsequent calculations.
       For the torque coils, the computer is required to calculate the amount of current
        needed to perform a maneuver, and pass this to the coils. In addition, a reading
        will be taken of the applied current to make sure that it is within tolerance limits.

        12.1.3.3       Power Interface
The power system needs to provide power to all of the ADC systems. In order for the
devices to work, the earth sensor, magnetometer, and rate gyros must receive the proper


VT-ICD-C-1 Issue Date: 21 July, 2000                                                        38
amount of power when required. The torque coils must receive a varying amount of
current to perform different maneuvers.

        12.1.3.4        Structural Interface
The interface with the structures system mainly includes the placement of each device, as
well as the method of mounting.

       The earth sensors are currently located on the odd numbered side faces of the
        satellite.
       The magnetometer is placed outside the satellite. It is fastened to HokieSat in a
        location that is as far away as possible from anything magnetic.
       Two of the three torque coils are located parallel to a specific face of the satellite.
        One hexagonal coil is mounted near the top hexagonal face, one rectangular coil
        is located near a side panel, and the other rectangular coil is placed orthogonal to
        those.
       The three rate gyros are placed such that they are mutually orthogonal.

        12.1.3.5        GPS Interface
Data from GPS is required in order for the system to operate correctly. The received
GPS data allows the satellite to know its orbital location. With this location known, the
attitude determination system knows what percentage of the Earth is in shadow as seen
by HokieSat. In addition, this data allows the magnetometer to compare the calculated
magnetic field with the actual.

        12.1.3.6        Propulsion
The ADCS interfaces with the propulsion system such that the two cannot be turned on at
the same time. The thrusters used for orbital maneuvers produces a magnetic field that
interferes both with the magnetometer measurements and with the magnetic field
produced by the torque coils.

       12.1.4 Components
The entire ADCS system is composed of a earth sensors, a magnetometer, three torque
coils, and rate gyros.

The earth sensor includes:
    3 CMOS cameras (Fuga 15d)
    3 Infinite Conjugate MicroVideo Imaging Lenses – focal length 4.8mm (Edmund
       Industrial Optics K53-221)
    3 aluminum mounts
    3 Optical Windows

The magnetometer includes:
    1 Three-Axis Magnetic Sensor Hybrid (Honeywell HMC2003)


VT-ICD-C-1 Issue Date: 21 July, 2000                                                         39
       Set/Reset Circuit & Signal Conditioning Circuit
       Aluminum Mount

   The Torque Coils include:
    Magnet Wire (Dearborn 30SP)
    Fasteners
    Kapton tape

The rate gyros include:
    3 Systron Donner Model QRS11-00050-100
    aluminum mount

       12.1.5 Mass Budget
  Component        Quantity length (in) width (in) height (in) total mass (lb)
  Earth Sensor       3          3           3          1.5           0.88
 Magnetometer        1          1         0.75         0.5         0.0661
Torque Coil (hex)    1        11.372     9.848        0.25         0.3188
Torque Coil (rect)   2         10.5         9         0.25         0.7296
   Rate Gyro         3        1.635      1.635        0.64         0.5423
      Total                                                        2.5368



       12.1.6 Power Budget
 Component Current (mA) Power (W) Voltage (V) # of Components
Earth Sensor   170         0.85      +/- 5           3
Magnetometer    20         0.3       6-15            1
 Torque Coil   0-75      0 - 0.25     3.3            3
  Rate Gyro     80         1.2       +/- 5           3


In addition, the magnetometer needs an initial 3 to 4 Amp pulse for approximately 1 sec
at the beginning of the mission in order to initialize the device. This requirement to reset
is due to the strong magnetic field due to the permanent magnets HokieSat is subjected to
from a nearby satellite.

       12.1.7 Status
At the current time, 1 CMOS Fuga 15d camera has been obtained, and others can be
ordered as needed. The placement of the cameras on the satellite have been determined,
as well as the method of horizon sensing. A rough outline of the algorithm to be used to
calculate the horizon from the rough sensor measurements has been explored. The
method of mounting the cameras to the satellite structure has been designed and currently
is being prepared for manufacturing.

The magnetometer device has been chosen, and has been purchased. Currently, it is
being tested. The connections to the computer and power systems are not known, and the
governing software has not been written. The structural location is not known.


VT-ICD-C-1 Issue Date: 21 July, 2000                                                     40
The final sizing for the magnetic torque coils has been finalized, and a prototype of the
system is being built. The method of current generation is under investigation.

One rate gyro is on-hand, and others will be purchased as needed.

In addition, the schedule for how often to take attitude readings must be decided, and
software uniting the system must be written.

   12.2        Component Overview

      12.2.1 Earth Sensor

       12.2.1.1        Operation of Earth Sensor
The GPS system determines the orbital location of the satellite. The percentage of Earth
not shadowed, or the expected view from HokieSat, is determined by specially written
software. The earth-horizon sensor cameras turn on and take readings. Data is sent to the
computer for calculations. The earth sensor turns off, and a computer program
determines the attitude of HokieSat based on the location of the horizon.

       12.2.1.2        Interfaces of Earth Sensor
The earth sensor interfaces with many systems on HokieSat. The computer tells the
sensor when to turn on and off, and the earth sensor sends data to the computer upon
completion of measurements. The power system sends power to the earth sensor when it
is turned on.

Structural interfaces with the earth sensor include the placement of the cameras, and
mounting. The cameras are mounted on three side faces of the satellite on odd-numbered
panels. Three different mount designs are required in order to fit onto each particular
side panel of the satellite. The cameras located on sides 1 and 3 require that a bar of the
isogrid panel be removed, while the camera on side 5 is located in the middle of an
isogrid triangle.

The method of mounting the camera, lens, camera board, and connectors to the isogrid is
shown below.




                            Figure 12.4 Mount on side panel 1




VT-ICD-C-1 Issue Date: 21 July, 2000                                                        41
                           Figure 12.5 Mount on side panel 3




                           Figure 12.6 Mount on side panel 5

The earth sensors are exposed to the outside of the satellite, and therefore have an
interface with the thermal system. The cameras are exposed to direct sunlight in addition
to shadowing. The operational thermal range of the cameras is between -20 and 60C.

The earth sensor system has an interface to the GPS system. Using the GPS system, the
orbital location of HokieSat is known, and therefore the portion of the Earth in eclipse
can be determined. This is necessary for checking the earth sensor data.

       12.2.1.3       Subcomponents of Earth Sensor
The earth sensor includes:
    3 CMOS cameras (Fuga 15d)
    3 Infinite Conjugate MicroVideo Imaging Lenses – focal length 4.8mm (Edmund
       Industrial Optics K53-221)
    3 aluminum mounts
    3 Connector sets (AirBorn, Inc. WTB30PR9SY, WTB30SAD11SY)
    1 Connector set (AirBorn, Inc. WGA122PR9SY, WGA122SACSY)
    Flat Ribbon Cable
    12 Fasteners 4-40 (NAS1352C04-12)
    9 Fasteners 10-32 (MS51958-67)
    6 Fasteners 4-40 (MS51957-17)
    6 Nuts 4-40 (MS21043-04)
    21 Standoffs
    3 Optical Windows




VT-ICD-C-1 Issue Date: 21 July, 2000                                                   42
       12.2.1.4       Mass Budget of Earth Sensor
    Component Mass (lb) Mass (gram)     # Total mass (lb) Total Mass (gram)
      Lens     0.0276      12.5         3     0.0828             37.5
      Board    0.0165       7.5         3     0.0495             22.5
      Cable     0.073       33          3      0.219             100
     Mounts    0.0882       40          3     0.2646             120
    Connectors                                 0.154              70
      Misc                                      0.11              50
      Total                                   0.8799             400



The mass of the mounting brackets is undetermined.

       12.2.1.5       Power Budget of Earth Sensor
      Component      Current (mA) Power (W) Voltage (V)      #
     CMOS Camera         170        0.85         5           3


       12.2.1.6       Status of Earth Sensor
At the current time, 1 CMOS Fuga 15d camera has been obtained, and others can be
ordered as needed. The placement of the cameras on the satellite has been determined, as
well as the method of horizon sensing. A rough outline of the algorithm to be used to
calculate the horizon from the rough sensor measurements has been explored. The
method of mounting the cameras to the satellite structure has been designed.

      12.2.2 Magnetometer

       12.2.2.1       Operation of Magnetometer
At the beginning of the mission, it is necessary to reset the magnetometer. This is
because the exposure to the permanent magnets from the nearby Utah State University
satellite causes a residual magnetic field. In order to reset the magnetometer, a set/reset
switch is implemented by running a 3 to 4 Amp current through the device for a very
short period of time. Furthermore, this reset will occur after all torque coil operations
and transmissions. This will prevent the magnetometer from measuring a saturation error.

During normal attitude acquisition, the magnetometer turns on, measures the magnetic
field, and then turns off. The measured data is sent to the computer, where it is compared
to the actual magnetic field at that location. The computer uses the magnetometer data,
along with the earth sensor calculations, to determine the attitude of the satellite. The
magnetometer must be turned off before the torque coils are turned on.

       12.2.2.2       Interfaces of Magnetometer
The interfaces of the magnetometer include the computer, the power system, structures,
GPS, and propulsion.


VT-ICD-C-1 Issue Date: 21 July, 2000                                                    43
The computer turns the magnetometer on and off as necessary for attitude measurements.
In addition, the computer measures the magnetometer voltage for measuring the Earth’s
magnetic field components. The computer interface is defined in Figure 12.4.



                  +12 V
                  Pwr Gnd
                  mag_sr

                  mag_x
                  mag_y
                  mag_z

                  mag_x_gnd
                  mag_y_gnd
                  mag_z_gnd
                                                     I/O board


                     Figure 12. 7 – Interface of Magnetometer and I/O
                                           Board
The power system sends power to the magnetometer when it is turned on.

The structural interfaces include the placement of the magnetometer, as well as the
method of mounting. The magnetometer will be located on the bottom facing out as far
away from devices that can act as magnets. Over time, the torque coils act to magnetize
everything and this causes the magnetometer measurements to be affected. In order to
minimize this, the greatest distance is put between the magnetometer and any magnetic
device.

The magnetometer relies on GPS. This is because the actual magnetic field must be
known in order to be compared to that measured by the magnetometer. If the orbital
location is known from GPS, the magnetic field at that location can be easily calculated.

The magnetometer cannot be operating at the same time as propulsion system. The
thrusters produce a magnetic field that interferes with the magnetometer readings.

       12.2.2.3       Subcomponents of Magnetometer
The magnetometer consists of a Three-Axis Magnetic Sensor Hybrid (Honeywell
HMC2003). There is also a set/reset circuit that will pass the required 3A current through
the device to reset it, removing error due to magnetic saturation from external devices. A
signal conditioning circuit will amplify the output voltages so that there is little or no
noise in the signal as it is sent to the IO board.



VT-ICD-C-1 Issue Date: 21 July, 2000                                                    44
The magnetometer will be mounted on the bottom of the satellite using the aluminum
mount shown below.




                           Figure 12.8 Magnetometer Housing

       12.2.2.4        Mass Budget of Magnetometer

 Component     Mass (lb)
Magnetometer     0.01
Mounts / Wires 0.0561
    Total       0.0661


       12.2.2.5        Power Budget of Magnetometer

  Component       Current (mA) Power (W) Voltage (V)
 Magnetometer          20         0.3       6-15


       12.2.2.6        Status of Magnetometer
At the current time, the magnetometer device has been chosen, and the interfaces
determined. The governing software has not been written. A magnetometer has been
purchased, and is in the process of being tested.

      12.2.3 Torque Coils

       12.2.3.1        Operation of Torque Coils
The torque coils operate after the computer software concludes that an attitude correction
is required. The computer calculates the amount of current needed to make the



VT-ICD-C-1 Issue Date: 21 July, 2000                                                    45
correction, and this current is then passed through the wire. After the correction has been
made, the coils are turned off. It is important that the magnetometer and the torque coils
are not operating at the same time because the two systems will affect each other
unfavorably.

       12.2.3.2        Interfaces of Torque Coils
The Torque Coils interface with the computer, power system, and structure. The
computer calculates the amount of current that needs to be run through the coils for a
specific attitude adjustment. The power system provides the current.

The torque coils interface to the structure in the way that they are mounted. The one
hexagonal coil is mounted on the inside of the satellite to the top hexagonal face, and the
two square coils are mounted such that they are mutually orthogonal, with one located
parallel to a side panel.

The method of mounting consists of using a bent piece of aluminum and standard
fasteners to hold the coil in place, as shown in Figure 12.5.




                                                             4-40 UNC
                                                               bolts
                           Copper
                                                 Aluminum
                            wire
                                                  support

                          Figure 12.9    Mount of Torque Coils


The bolts will be placed such that they fit the 2-inch spacing on the isogrid and not
require any further drilling. The area on the coils that is in contact with the aluminum
mount will be covered with silicon gel to prevent rubbing of the two surfaces.

The computer interface is defined as shown in Figure 12.10:




VT-ICD-C-1 Issue Date: 21 July, 2000                                                       46
                    coil_x_pos                 coil_x
                    coil_x_neg

                    coil_y_pos                 coil_y
                    coil_y_neg

                    coil_z_pos                 coil_z
                    coil_z_neg

                    coil_x_volt_pos
                    coil_x_volt_neg

                    coil_y_volt_pos
                    coil_y_volt_neg                      Torque
                                                         board
                   coil_z_volt_pos
                   coil_z_volt_neg
                 Figure 12.10 Interface of Torque Coil and Computer

Allegro Microsystems’ A3966SA Dual Full Bridge PWM Motor Driver will drive the
torque coils.

An example of the schematic of the chip is as follows:




                Figure 12.11 Layout of the A3966SA with the truth table.




VT-ICD-C-1 Issue Date: 21 July, 2000                                             47
                Figure 12.12 Functional Block Diagram of the A3966SA


Each chip is capable of driving two coils. Therefore, two of the surface mounted chips
will be used in the process.

Depending on the amount of voltages received in the SENSE 1 and SENSE 2 as
compared with the reference voltage, either Output 1 or Output 2 will be activated. Then
a program will be written to produce the desired amounts of current that will be sent
through the coils. The voltage that is produced in the output is proportional to that
received through the PHASE node.

It is important to note that the torque coils affect many of the surrounding systems.
Anything that has the ability to become magnetized is at risk of obtaining a residual
magnetic field. This gradually builds up over the duration of the mission.

       12.2.3.3       Subcomponents of Torque Coils
The torque coils are made of magnet wire. Each coil will be covered Kapton Tape for
stiffness. Aluminum fasteners are used.

The hexagonal coil consists of 316 loop of wire. The two rectangular coils are each made
of 341 loops. The cross sectional dimension of the wire bundles is approximately 0.25 in
 0.25 in.




VT-ICD-C-1 Issue Date: 21 July, 2000                                                     48
       12.2.3.4        Mass Budget of Torque Coils


   Component         Mass (lb)
Hexagonal Coil (1)    0.2818
Rectangular Coil (2) 0.6556
      Misc             0.11
      Total           1.0474


       12.2.3.5        Power Budget of Torque Coils


Component Current (mA) Power (W) Voltage (V) #
Torque Coils  0-75       0-2.5       0.3     3


       12.2.3.6        Status of Torque Coils
The final sizing for the magnetic torque coils is complete, and the coils are in the process
of being manufactured. The chip used for current generation has been tested in a simple
test environment, and has produced the expected results.


      12.2.4 Rate Gyros

       12.2.4.1        Operation of Rate Gyros
The rate gyro provides an output signal that is proportional to its angular rate about its
input axis. For three-axis applications, three rate gyros are necessary.

       12.2.4.2        Interfaces of Rate Gyros
The rate gyros are an integral component of the attitude determination and control
system, and the interface is an analog rate output signal. They interface with the
spacecraft structure through small mounting brackets. They interface with the power
system through the power supply and ground lines.

The computer interface is defined as shown in Figure 12.13:




VT-ICD-C-1 Issue Date: 21 July, 2000                                                         49
                     +5V
                     Gnd
                     -5V
                 rate_x
                 rate_y
                 rate_z

                 rate_x_pos
                 rate_x_neg

                 rate_y_pos
                 rate_y_neg

                 rate_z_pos
                 rate_z_neg

                 rate_x_bit
                 rate_y_bit                    gyro board
                 rate_z_bit


                    Figure 12.13 Interface of Gyros and Computer

      12.2.4.3       Subcomponents of Rate Gyros
The control moment gyro used is a Systron Donner Model QRS11-00050-100. Three of
these are used.

      12.2.4.4       Mass Budget of Rate Gyros

  Component      Mass (lb) # Total Mass (lb)
  Rate Gyro       0.1323 3       0.3768
 Mounts, wires                   0.1455
    Total                        0.5423


      12.2.4.5       Power Budget of Rate Gyros
 Component   Current (mA) Power (W) Voltage (V)       #
 Rate Gyro        80         1.2       +/- 5          3




VT-ICD-C-1 Issue Date: 21 July, 2000                                           50
       12.2.4.6      Status of Rate Gyros

A single rate gyro has been bench tested, and produced nominal results. One QRS11 is
on-hand, and others will be purchased as needed. The gyro mount and gyro board are
being developed and tested by the ADCS team at the University of Washington.




VT-ICD-C-1 Issue Date: 21 July, 2000                                                   51
13.0           Propulsion Subsystem

   13.1        System Overview

      13.1.1 Description
The purpose of the propulsion system is to provide the impulse necessary for the
formation-flying mission. The flight control software determines the required impulse
for a maneuver. The propulsion system uses two pulsed plasma thruster (PPT) pairs
mounted in the center (vertically) of a sidewall as shown in Figure 13.1. Thrusters 2 and
3 provide translation; thrusters 1 and 4 provide rotation about the yaw axis.


                                                                        1
                                                       2




                                                  3

                                                                        4



                           Figure 13.1    Position of Thrusters




VT-ICD-C-1 Issue Date: 21 July, 2000                                                    52
                                  Spring

                                                                 Teflon Fuel
          Ultem 2300
       Thruster Assembly                                         Bar
                                                                 Mica-paper / Foil
                                                                    Capacitor




                  Boron Nitride
                    Insulator

                            Figure 13.2 Pulsed Plasma Thruster




               Ultem Isolator


        Cathode




      Annode




VT-ICD-C-1 Issue Date: 21 July, 2000                                                 53
                               Figure 13.3 Thruster Housing




      13.1.2 System Operation
The propulsion system operates on an as needed basis. When translation is required for
formation flying, the flight control computer determines the direction and duration of
thrust required to perform the maneuver. The torque coils and thrusters 1 and 4 are used
to point the thrusters in the required direction. The thrusters and torque coils cannot
operate simultaneously since the thrusters produce a magnetic field that would interfere
with torque coil operation. For the same reason, the thrusters cannot be operated
simultaneously with the magnetometer. Specifications for one thruster are given in the
table below.


 Impulse Bit                56 Ns
 Specific Impulse           485 s
 Thrust (1 Hz fire rate)    56 N




      13.1.3 Interfaces
The primary interfaces of the propulsion system are with the structure, computer, power
and attitude control subsystems.

       13.1.3.1        Thermal Requirements
Component                      Heat Production                 Maximum Temp (C)
Capacitor                      20 % of input power             125
Electronics                    20 % of input power             100
Thruster housing               10% of input power              No concern



       13.1.3.2        Computer Interface
The computer must send commands to the propulsion system designating which thruster
is to fire, as well as the actual fire command. Additionally, telemetry data are required to
verify thruster operation. Telemetry requirements are listed below.




VT-ICD-C-1 Issue Date: 21 July, 2000                                                      54
      4 Digital input lines for controlling the thruster
      4 Analog signals for telemetry (specific type of telemetry TBD)

       13.1.3.3        Power Interface
The propulsion system requires 13 watts at 16.5 – 22.5 volts during operation for the
translational thrusters and 6.5 watts at 16.5 – 22.5 volts for the rotational thrusters. While
the system is not in use, it requires no power.

       13.1.3.4        Structural Interface
Thrusters 1 and 4 require two cutouts centered vertically in a sidewall (Figure 13.1),
approximately 1.996 inches (across) by 0.810 inches (tall). Similar cutouts are required
for thrusters 2 and 3, one each on the two adjacent faces measuring 2.235” x 0.810”. All
of the attachment points are on the bottom surface of the thruster. This was done so that
the PPT did not become load bearing. Mounting the thrusters in the center of the
sidewalls requires support which may be provided by a computer box or some other
appropriate component or structure. Additional support is required for the Teflon bars to
prevent bending during launch.

Dimensions of the thruster unit and structural interface are shown below in Figure 13.5.




VT-ICD-C-1 Issue Date: 21 July, 2000                                                       55
                  Figure 13.5 Thruster and Side Structure Interface

     13.1.4 Components
The propulsion system consists of:
    Two thruster assemblies, each with
          o Two cathodes (copper)
          o Two anodes (stainless steel)
          o Wiring
          o Two Fuel bars (Teflon)
          o Structure (Ultem 2300)
          o Thrust chamber (Boron Nitride)
          o One Capacitor (Mica-Paper Foil)
          o Two spark plugs
    One power-processing unit
    EMI filter
    Four discharge initiation circuits



VT-ICD-C-1 Issue Date: 21 July, 2000                                  56
     13.1.5 Mass Budget
 Component        Quantity Length, in   Width,      Height, in   Total mass,
                           (cm)         in (cm)     (cm)         lb. (kg)
 Thruster Assy    2        8.6 (21.8)   3.1 (7.9)   1.5 (3.81)   0.8 (0.36)
 Capacitor        2        4 (10.16)    2 (5.08)    1 (2.54)     0.8 (0.36)
 PPU              1        Tbd          Tbd         Tbd          Tbd
 DI circuit       4        Tbd          Tbd         tbd          0.25 (0.11)
 Total

     13.1.6 Power Budget
 Component       Power (W)     Voltage (V)
 System          13            16.5 – 22.5



     13.1.7 Status of System
Awaiting further developments of UW thruster design.




VT-ICD-C-1 Issue Date: 21 July, 2000                                           57
14.0             Thermal Subsystem

   14.1          Subsystem Overview

      14.1.1            Description
Analyses were performed on the satellite using a program called I-DEAS Master Series 8.
A basic form of the satellite was created and used with the thermal modeling software,
TMG, in two scenarios; worst case cold and worst case hot. Results were acquired for a
period of two orbits for each case. A thermal design was then created for the satellite
components using these results.

      14.1.2 Operational States

The maximum and minimum temperatures at which components can function and be
stored constitute its operating and non-operating temperature ranges respectively. These
ranges along with the data from orbit analysis determine whether insulation or active
thermal heating is required. Table 14.1 shows each subsystem’s components and their
temperature ranges.

Table 14.1: Subsystem Component Temperature Ranges

                                           Operating Range
           Subsystem       Component             (C)         Non-operating Range (C)
 Propulsion                capacitators       max: 125                N/A
                            electronics       max: 100                N/A
 Computer                       all           -40 to 85            -55 to 125
 Attitude and Control       rate gyro         -40 to 80            -55 to 100
                           magnometer         -40 to 85            -55 to 125
                             cameras          -20 to 60             -60 to 60
                            Downlink
 Comm                      Transmitter        -20 to 70            -55 to 100

 Battery                    Charging           0 to 45                N/A
                           Discharging        -20 to 60               N/A

 Solar Cells                                 -100 to 100           -100 to 100

 Power                   DC-DC converter     –55 to +125




VT-ICD-C-1 Issue Date: 21 July, 2000                                                   58
        14.1.3 Analysis Description and Results
To begin, a simple hexagonal box was created to represent the spacecraft. Inside, a
smaller hexagonal box was inserted and partitioned into the spacecraft. By partitioning
the two boxes they were joined in a way that I-DEAS would recognize one was inside the
other.

The next step was to mesh the boxes in order to analyze smaller sections and get more
accurate results. The small box was meshed using solid meshing and a null material
property. Then its outside panels were meshed with a shell mesh and null material. The
satellite box was also meshed, but with only a shell mesh on its eight sides. Three
materials were used; aluminum for the top panel, a combination of the solar cells and
aluminum qualities for the side panels, and aluminum painted with white acrylic paint for
the bottom panel. Characteristics of the materials can be found in Table 14.2.

Table 14.2: Material Characteristics
                                             Material 1             Material 2           Material 3     Material 4
                                          Pure Aluminum      Cell/Grid Combination   Painted Aluminum     Null
Mass Density (kg/m^3)                          2702                   2702                 2702          -
Thermal Conductivity (W*K/m)                    237                    237                  237            0
Emissivity                                    0.0346                   0.83                 0.9          -
Specific Heat Below Phase (Cp) (J*K/kg)         903                    903                  903          -
Solar Absorptivity                             0.379                  0.673                 0.26         -



Changing from the meshing task to the TMG thermal analysis, boundary conditions were
defined. First the radiation vectors were defined to point outward and a space enclosure
was created as -269C. This allowed any radiation from the satellite to radiate into space.
Then, to allow TMG to view radiation factors between all the elements, a radiation
request was created for all radiation. Lastly a thermal couple was created that would
simulate the heating of the satellite’s internal box. Radiation for the thermal couple was
given a view factor of .45.

Orbit parameters were then defined for each case as shown in Table 14.3.

Table 14.3: Orbit Parameters Used in Analysis
  ORBIT PARAMETERS

  Altitude (nmi)                                               200
  Eccentricity                                                   0
  Right Ascension of Ascending Node                              0
  Inclination                                             51.6 deg
  Solar Declination                                              0
  Albedo:
                                  cold case                  0.22
                                   hot case                  0.36
  Earth emitted IR (Btu/hr ft^2):
                                  cold case                     73
                                   hot case                     81

VT-ICD-C-1 Issue Date: 21 July, 2000                                                                             59
The analysis was then performed, and the results are shown in Table 14.4.

Table 14.4: Analysis Results
Table 3:
Temperature Results

                Hot                     Cold
                Case                    Case
                Min        Max          Min     Max
Satellite       -13.3      +10.9        -13.3   +10.9
Inside          -2.88      0            -2.88   0


These results suggest that for both the hot and cold cases the satellite will not undergo a
large range in temperature change throughout the orbit. They also describe the insides as
having an extremely small temperature change during orbit. Both inside and the satellite
itself’s temperatures center around 0C.

      14.1.4 Interfaces
The thermal subsystem interfaces with the propulsion, computer, power, attitude
determination and control, and communication subsystems.

       14.1.4.1         Propulsion Interface
The capacitor will need a thermistor.

       14.1.4.2         Computer Interface
The computer box will need to be covered with Multi-Layer Insulation (MLI) to prevent
the components from getting too cold. Also, the temperature sensors will connect to the
computer, which will then activate or deactivate the heaters.

       14.1.4.3         Power Interface
The battery box will require a thermal thermistor.

       14.1.4.4         ADCS Interface
The cameras will need to be insulated or actively heated.

       14.1.4.5         Communications Interface
The Downlink Transmitter will need to be insulated or actively heated.




VT-ICD-C-1 Issue Date: 21 July, 2000                                                     60
      14.1.5 Components
The thermal system consists of:
        Thermistors
        MLI (Multi-Layer Insulation)
        White Paint

      14.1.6 Power Budget
Each heater-temperature sensor requires little power.

      14.1.7 Status
Worst case cold and worst case hot analyses will be simulated with interior components
placed in position and able to dissipate/absorb heat to/from the structure. Also, an
alternate way of creating orbit conditions will be attempted in an effort to validate present
results.

   14.2        Component Overview


      14.2.1 Thermistors

       14.2.1.1        Operation of Thermistors
The thermistor will heat the battery box to the minimum required operational temperature
so as not to overheat when the satellite receives solar radiation. Active temperature
control of the cameras and downlink transmitter may also be required.

       14.2.1.2        Status of Thermistors
The Shrink Sleeve Probe, Part #H2049, is currently being considered. Manufactured by
the U.S. Sensor Corporation, this thermistor has an accuracy of  .20C when the
environment remains within 0 to 70C and can operate up to a maximum of +150C.
Other possible thermistors could be purchase from Minco Products Inc.
(www.minco.com).

      14.2.2 Multi-Layer Insulation (MLI)


       14.2.2.1        Purpose of MLI
The MLI will be used to insulate the spacecraft’s insides and the components that are
thermally sensitive. MLI prevents a material from losing or gaining heat at a rapid rate.




VT-ICD-C-1 Issue Date: 21 July, 2000                                                      61
       14.2.2.2        Status of MLI
Awaiting selection of type and manufacturer.

      14.2.3 White Paint

       14.2.3.1                Purpose of White Paint
White paint will be used to cover the outside of the Nadir-facing panel of the satellite.
This will allow greater emissivity and lower solar spectral absorbtivity. By increasing the
emissivity and decreasing the absorbtivity a heat sink panel is created that will allow heat
to be dissipated into space at a rate greater than that of pure aluminum.




VT-ICD-C-1 Issue Date: 21 July, 2000                                                     62

				
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