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
Chris Hall
Signature
Subsystem
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
Propulsion
Chris Karlgaard
Date
Science
VT-ICD-C-1 Issue Date: 21 July, 2000
1
2.0
History of Revisions
Revision
Date
VT-ICD A-1
25 February 2000
VT-ICD A-2
22 March 2000
VT-ICD A-3
30 March 2000
VT-ICD B-1
19 June 2000
VT-ICD B-2
26 June, 2000
VT-ICD B-2b
12 July, 2000
VT-ICD C-1
21 July, 2000
VT-ICD-C-1 Issue Date: 21 July, 2000
Comments
Initial input from subsystem leads.
Compiled by Adam C. Harvey.
Includes changes from ICD Review on 4
March 2000.
Compiled by Adam C. Harvey.
Includes changes from ICD Review on 23
March 2000.
Compiled by Adam C. Harvey.
Configuration Freeze
Compiled by Adam C. Harvey.
Updates from Power Subsystem.
Adam C. Harvey
Includes changes from Andrew Turner.
Adam C. Harvey
Includes updates for HokieSat CDR.
Includes Thermal Subsystem section.
Adam C. Harvey
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
9.1.4
9.1.5
10.0
Thermal Requirements .................................................................................. 23
I/O Card ........................................................................................................ 24
Signal Definitions ......................................................................................... 24
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
12.2.2
12.2.3
12.2.4
13.0
Earth Sensor .................................................................................................. 41
Magnetometer ............................................................................................... 43
Torque Coils.................................................................................................. 45
Rate Gyros .................................................................................................... 49
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
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4.0
3CS
ADCS
AFRL
APL
BIST
CDMA
CPU
DI
DOD
FOV
FPGA
FSK
GPS
GSFC
HH
HFC
ION-F
JHU
MLI
MSDS
Nanosatellite
OS
Payload
PCM
PPT
PPU
RAM
SHELS
SRAM
Stack
UHF
UNP
USU
UW
VT
VT-ISMM
List of Acronyms and Definitions
Three Corner Sat
Attitude and Determination Control Subsystem
Air Force Research Laboratory
Applied Physics Laboratory
Built In Self-Test for computer subsystem
Code Division Multiple Access
Central Processing Unit
Discharge Initiation
Depth of Discharge
Field of View for cameras
Floating Point Gate Array
Frequency Shift Keying
Global Positioning System
Goddard Space Flight Center
Hitchhiker
HokieSat Flight Computer
Ionospheric Observation Nanosatellite Formation
Johns Hopkins University
Multiple Layered Insulation
Multiple Satellite Deployment System
One of three ION-F satellites
Operating System
Two satellite stacks onboard the MSDS platform
Pulsed Code Modulation
Pulsed Plasma Thruster
Power Processing Unit
Random Access Memory
Shuttle Hitchhiker Ejection Launch System
Static RAM
The ION-F stack of three satellites
Ultra High Frequency
University Nanosat Program
Utah State University
University of Washington
Virginia Tech
Virginia Tech Ionospheric Scintillation Measurement Mission
VT-ICD-C-1 Issue Date: 21 July, 2000
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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
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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
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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
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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
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Table 8.1: External Layout, Side 1
Side
1
1
1
1
Component
Solar Cell String 1A
Solar Cell String 1B
PPT A1
Camera 1
Table 8.2: External Layout, Side 2
Side
2
2
2
2
Component
Solar Cell String 2A
Solar Cell String 2B
PPT A2
C/L Patch 1
Table 8.3: External Layout, Side 3
Side
3
3
3
3
Component
Solar Cell String 3A
Solar Cell String 3B
Camera 2
PPT B3
VT-ICD-C-1 Issue Date: 21 July, 2000
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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
6
6
6
Component
Solar Cell String 6A
Solar Cell String 6B
C/L Patch 3
Figure 8.2
Front view of the side panel external configuration
VT-ICD-C-1 Issue Date: 21 July, 2000
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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
Nadir
Nadir
Nadir
Component
U/L Antenna
D/L Antenna
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
Nadir
Nadir
Zenith
Side 1
Side 1
Side 2
Side 3
Side 3
Component
Electronics Enclosure
Battery Enclosure (Not Pictured)
Hexagonal Torque Coil
Earth Sensor
PPT 1A (Not Pictured)
PPT 2A (Not Pictured)
Earth Sensor
PPT B1 (Not Pictured)
VT-ICD-C-1 Issue Date: 21 July, 2000
14
Side 3
Rectangular Torque Coil 1
Side 4
Side 4
Side 4
Side 5
Side 6
Rectangular Torque Coil 1
PPT B2 (Not Pictured)
Rectangular Torque Coil 2
Earth Sensor
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
Side
1
1
Isometric views of the isogrid side and end plates
Component
PPT A1
Camera 1
10.425”
13.725”
9.0”
VT-ICD-C-1 Issue Date: 21 July, 2000
16
Side
Component
PPT A2
2
Side
3
3
Side
Component
PPT B4
4
Side
5
Side
6
Figure 8.7
Component
Camera 2
PPT B3
Component
Camera 3
Component
Data Test Port
Interior views of the side panels.
VT-ICD-C-1 Issue Date: 21 July, 2000
17
Side
Zenith
Component
Magnetometer
18.00”
Side
Nadir
Component
None
Figure 8.8
Interior views of the nadir and zenith panels.
VT-ICD-C-1 Issue Date: 21 July, 2000
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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
8.1.2
Top view of wall scabs
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
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8.1.4
Components
Table 8.10: External Layout, Side
Component
Side Panel
Top Panel
Bottom Panel
Wall Scabs
Fasteners
8.1.5
QTY
6
1
1
6
TBD
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
Side Panel
6
8.2.1.1
H
cm, (in)
34.9, (13.725)
Mass Budget of Side Panels
Component QTY
Side Panel
W
cm, (in)
23.1, (9.085)
6
Mass
Each g, (lbm)
521.5, (1.1497)
VT-ICD-C-1 Issue Date: 21 July, 2000
Mass
Total g, (lbm)
3129.78, (6.900)
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
Top Panel
Bottom Panel
1
1
8.2.2.1
DIA
(in)
18.00
18.00
Mass Budget of Top and Bottom Panels
Component
QTY
Top Panel
Bottom Panel
1
1
8.2.2.2
Mass
Each g, (lbm)
1133.5, (2.499)
1133.5, (2.499)
Mass
Total g, (lbm)
1133.5, (2.499)
1133.5, (2.499)
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
Wall Brackets
6
Mass
Each g, (lbm)
51.71, (0.114)
VT-ICD-C-1 Issue Date: 21 July, 2000
Mass
Total g, (lbm)
310.71, (0.685)
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
Slot 6
Slot 7
Slot 8
Slot 9
Power Board 2
Slot 5
Data Bus Backplane
Power Board 1
PDP Board
System
Overview
Slot 1
Slot
2
Slot 3
Spare 1
IO Board
Slot 4
Computer Subsystem
Spare 2
Camera Board
Telemetry Board
CPU Board
9.1
Gyro Board
Rate Gyros
9.0
SPI Bus Backplane
Backplane
The computer system has three commonCommon
boards.
These are the CPU board, the telemetry
Dawgstar
board, and the I/O board. Figure 9.1 shows
various board positions in relation to the
Backplane.
Slot 6
Slot 7
Slot 8
Slot 9
Power Board 2
Slot 5
Power Board 1
PDP Board
Slot 4
Spare 1
IO Board
Slot 3
Data Bus Backplane
Torque Coil Board
Camera Board
CPU Board
Slot 2
Telemetry Board
Gyro Board
Rate Gyros
Slot 1
SPI Bus Backplane
Common Backplane
HokieSat
Figure 9.1
9.1.1
Board Positions in Relation to the Backplane
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
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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 SEGRresistant 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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24
Table 9.2: I/O Board Signal Definitions
Directi
on
Sub
System
Signal
Type
3.3V
ANALOG
3.3V
ANALOG
3.3V
DISCRET
E
Discrete
GPS/Xlink I/O
Uplink/Do Discrete
wnlink
I/O
Discrete
I/O
5V
DISCRET
IO Board E
Discrete
PPT
I/O
Discrete
PPT
I/O
PPT
PPT
A/D
wrt IO Signal
board Name
Backplane
Pin #
OUT
?
OUT
?
Top Rang
pin e
Description
Pin #
[0- Current Monitor for on
N/A 3.3V] I/O board
[0- Overcurrent Interrupt for
N/A 3.3V] on I/O board
OUT
?
[0,3. A/D conversion complete
N/A 3V] (I/O board)
DO_2_XLI
OUT NKPWR
N/A
DO_3_TXP
OUT WR
N/A
DO_4_RXP
OUT WR
N/A
DO_DISC_
OUT 0
N/A
DO_6_FIRE
OUT 1
N/A
DO_7_FIRE
OUT 2
N/A
DI_PPT_IN
IN
1
N/A
DI_PPT_IN
IN
2
N/A
?
0,5V Power on/off# XLINK
?
0,5V Power on/off# transmitter
?
0,5V Power on/off# receiver
?
[0,5V
]
Power on/off PPT board
?
5V? Fire 1
?
5V? Fire 2
?
0,5? Voltage 1
A/D
?
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.
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26
10.0
10.1
Power Subsystem
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
Voltage regulator circuit
Current regulator circuit
Purpose
To regulate voltage from solar cells.
To regulate and remove unwanted current
(Shunt regulator to be placed outside of satellite
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27
Chargeable Batteries
DC to DC converter
UController/Chrg Controller
10.2
to dissipate heat) Prevents battery from being
over charged.
To provide power to satellite during eclipse
To distribute voltages to all sub-systems
(To be built by VPT, Dan Sable)
Determines charge rate of battery
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
Computer
5V and 3V
A/D
S Band Xmit
UHF Rcvr
11-15V
Torque Coils 3.3 V
Cameras
+/- 5V
Magnet
+/- 15V
Rate Gyro
+/- 5V
PPT
28 4V
Relays
28V
GPS
5V
Power
>3.0W (max)
14W
5.4W – 15W
0.25W
0.85W  3
0+.3W
1.2W  3
13W
Comment
Can it take +15V
Can it take 15.4-21V
Can it take 15.4-21V
Can it take 5V
(4A/1 usec pulse)
Can it take 15.4-21 V
1.4W
Table 10.2: Voltage Distribution from DC-DC Converters
DVSA283R3S + 3.3V
T. Coil
Cameras
DVSA2805D  5V
DVSA2815S
+15V
Magnetom
Rate Gyro
DVSA2805D  5V
DVSA2805S
+5V
GPS
DVSA283R3S + 3.3V
Computer
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.
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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 22002290MHz, 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.
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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
L3 ST 802 S
S-Band Transmitter
Hamtronics R451* UHF Receiver
X-Link/GPS
Hardware from APL
Patch Antenna (3)
S-Band Transmitting
Antenna
Loop
UHF Receiving
Antenna
GPS Patch Antenna GPS L-Band Patch
Antenna
Total (approx.)
Assume 5 W for X-Link
Power
14W
1.5W
1.4W
NA
Voltage
28V
11-15V
5V
NA
Mass g, (lb)
130 (0.2866)
105 (0.23125)
750 (1.65)
30, (0.0661)
NA
NA
10, (0.0220)
NA
NA
10, (0.0220)
?
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
Transmitter
Receiver
X-Link Hardware
Antennas
Current Status
Has been decided upon
Has been decided upon, and is being modified
by USU
APL is working on that
Currently being designed
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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
Pin Connectors
J-2 RF Output
J-3 Modulation Input
MDM-15S
1 through 8 Freq. Selection
9 – PWR RTn
10 – Spare
11 – Spare
12 – Spare
13 - +28V
14 – PWR RTN
15 – PWR RTN
SMA Female
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.
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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)
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34
12.0
12.1
Attitude Determination and Control Subsystem
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
VT-ICD-C-1 Issue Date: 21 July, 2000
Camera Placement
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.
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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 reinitialization 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
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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
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42
12.2.1.4
Mass Budget of Earth Sensor
Component Mass (lb) Mass (gram)
Lens
0.0276
12.5
Board
0.0165
7.5
Cable
0.073
33
Mounts
0.0882
40
Connectors
Misc
Total
# Total mass (lb) Total Mass (gram)
3
0.0828
37.5
3
0.0495
22.5
3
0.219
100
3
0.2646
120
0.154
70
0.11
50
0.8799
400
The mass of the mounting brackets is undetermined.
12.2.1.5
Component
CMOS Camera
12.2.1.6
Power Budget of Earth Sensor
Current (mA) Power (W) Voltage (V)
170
0.85
5
#
3
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.
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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.
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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
Component
Magnetometer
12.2.2.6
Power Budget of Magnetometer
Current (mA) Power (W) Voltage (V)
20
0.3
6-15
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
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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
wire
Figure 12.9
Aluminum
support
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:
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coil_x_pos
coil_x_neg
coil_x
coil_y_pos
coil_y_neg
coil_y
coil_z_pos
coil_z_neg
coil_z
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.
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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.
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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:
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+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
rate_z_bit
gyro board
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
Component
Rate Gyro
Mounts, wires
Total
12.2.4.5
Component
Rate Gyro
Mass Budget of Rate Gyros
Mass (lb) # Total Mass (lb)
0.1323 3
0.3768
0.1455
0.5423
Power Budget of Rate Gyros
Current (mA) Power (W) Voltage (V)
80
1.2
+/- 5
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#
3
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.
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13.0
13.1
Propulsion Subsystem
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
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Position of Thrusters
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
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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
Specific Impulse
Thrust (1 Hz fire rate)
56 Ns
485 s
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
Component
Capacitor
Electronics
Thruster housing
13.1.3.2
Thermal Requirements
Heat Production
20 % of input power
20 % of input power
10% of input power
Maximum Temp (C)
125
100
No concern
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.
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

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.
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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
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13.1.5 Mass Budget
Component
Thruster Assy
Capacitor
PPU
DI circuit
Total
Quantity Length, in
(cm)
2
8.6 (21.8)
2
4 (10.16)
1
Tbd
4
Tbd
Width,
in (cm)
3.1 (7.9)
2 (5.08)
Tbd
Tbd
Height, in
(cm)
1.5 (3.81)
1 (2.54)
Tbd
tbd
Total mass,
lb. (kg)
0.8 (0.36)
0.8 (0.36)
Tbd
0.25 (0.11)
13.1.6 Power Budget
Component
System
Power (W)
13
Voltage (V)
16.5 – 22.5
13.1.7 Status of System
Awaiting further developments of UW thruster design.
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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
Subsystem
Propulsion
Computer
Attitude and Control
Comm
Battery
Component
Operating Range
(C)
Non-operating Range (C)
capacitators
max: 125
N/A
electronics
max: 100
N/A
all
-40 to 85
-55 to 125
rate gyro
magnometer
cameras
Downlink
Transmitter
-40 to 80
-40 to 85
-20 to 60
-55 to 100
-55 to 125
-60 to 60
-20 to 70
-55 to 100
Charging
Discharging
0 to 45
-20 to 60
N/A
N/A
-100 to 100
-100 to 100
Solar Cells
Power
DC-DC converter
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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
Mass Density (kg/m^3)
Thermal Conductivity (W*K/m)
Emissivity
Specific Heat Below Phase (Cp) (J*K/kg)
Solar Absorptivity
Material 1
Pure Aluminum
2702
237
0.0346
903
0.379
Material 2
Cell/Grid Combination
2702
237
0.83
903
0.673
Material 3
Painted Aluminum
2702
237
0.9
903
0.26
Material 4
Null
0
-
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)
Eccentricity
Right Ascension of Ascending Node
Inclination
Solar Declination
Albedo:
cold case
hot case
Earth emitted IR (Btu/hr ft^2):
cold case
hot case
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200
0
0
51.6 deg
0
0.22
0.36
73
81
59
The analysis was then performed, and the results are shown in Table 14.4.
Table 14.4: Analysis Results
Table 3:
Temperature Results
Satellite
Inside
Hot
Case
Min
-13.3
-2.88
Max
+10.9
0
Cold
Case
Min
-13.3
-2.88
Max
+10.9
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.
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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.
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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.
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