Small Satellite Thermal Modeling and Design at USAFA:
FalconSat-2 Applications1
C1C Richard Lyon2
LtCol Jerry Sellers2
Craig Underwood3
2
USAF Academy Small Satellite Research Center
USAF Academy, CO
3
Surrey Space Centre, University of Surrey,
Guildford, Surrey, UK
C02Richard.Lyon@usafa.edu
Jerry.Sellers@usafa.edu
C.Underwood@eim.surrey.ac.uk
Abstract—The US Air Force Academy FalconSat program
is one in which undergraduate cadets design, build, test, and
operate satellites to carry Air Force and Department of
Defense payloads for scientific missions. Currently, cadets
are working on FalconSat-2, designed to carry the Micro
Electro-Static Analyzer (MESA) payload that will
investigate the morphology of plasma depletions in the
ionosphere. The Engineering Model was completed and
tested in April 2001, and cadets will construct the
Qualification and Flight Models in the fall of 2001. To aid
in the development of the satellite, behavioral models of
various spacecraft subsystems have been created using
MatLab and used to simulate projected operational modes of
the satellite and the effects on major satellite subsystems.
One major subsystem that had been overlooked until this
summer was the thermal subsystem. We require a detailed
thermal model to aid in the development and testing of
FalconSat-2 for several reasons. First, we wish to predict
the thermal behavior of the satellite in the various thermal
tests it will undergo in the development process. We also
wish to predict the thermal behavior of the satellite in
various expected operational modes and attitudes. This will
in turn enable us to design and implement any required
thermal control for the satellite. Over the summer, during
research performed at the University of Surrey, a thermal
model of FalconSat-2 was created in MatLab using finite
differential analysis and a lumped-parameter approach. The
FalconSat-2 model was adapted from models developed by
Dr. Craig Underwood, which have been used over the years
in the design and analysis of Surrey’s small satellites –
including, most recently, the UK’s SNAP-1 nano-satellite.
This paper will detail the development process undergone in
creating the FalconSat-2 thermal model, will demonstrate
how the model works, and will validate the results.
Additionally, the paper will describe the thermal control
solutions implemented for FalconSat-2 and how the model
is used in the development process.
TABLE OF CONTENTS
1.
2.
3.
4.
5.
INTRODUCTION
MODEL DESCRIPTION
MODEL VERIFICATION
THERMAL DESIGN DISCUSSION
CONCLUSION
1. INTRODUCTION
The capstone of the United States Air Force Academy
Astronautics curriculum is the FalconSat Program. One
goal of the program, housed within the Academy’s Small
Satellite Research Center, is to give undergraduate cadets
the unique opportunity to “learn space by doing space.” The
program facilitates cadet development of small satellite
mission design through instructor guidance and mentorship.
It allows cadets to gain real-world experience with satellite
design, assembly, integration, testing, and operations within
the context of a two-semester engineering course sequence.
A second goal of the program is to provide a useful
nanosatellite platform for Air Force and Department of
Defense space experiments.
Through FalconSat
participation, cadets receive the hands-on opportunity to
apply the tools developed in a classroom to a real program,
ideally preparing them for the situations they may encounter
as officers and as engineers after they graduate.
The current project, FalconSat-2, is the third satellite to be
developed within the Academy’s program. The satellite’s
primary payload is the Micro Electro-Static Analyzer
(MESA) sensor suite, designed to study plasma depletions
in the F region of the ionosphere. It will be launched on the
Space Shuttle as part of the small payloads Hitchhiker
project. FalconSat-2 will be mounted in a Get Away
Special (GAS) canister with the Hitchhiker Motorized Door
Assembly (HMDA) and will use the Pallet Ejection System
(PES). The satellite is built around a “FalconSat-N”
approach, meaning the spacecraft bus is designed so that it
will be easily adaptable to carry future payloads. As such,
the basic design is one of an outer structural shell upon
which the solar panels and MESA sensors are mounted, and
an inner column around which the other subsystems are
placed in module boxes. The satellite is a 12.5-inch cube,
with the solar panels placed on the +X, -X, +Y, and –Y
facets, the MESA sensors, S-band antenna, and whip
antenna placed on the +Z facet, and the interface ring placed
on the –Z facet for attachment with the Space Shuttle’s Get
Away Special (GAS) canister. Figure 1a shows an external
view of FalconSat-2, and Figure 1b shows an exploded view
detailing key features and components.
The basic purpose of thermal design is to maintain the
temperature of all spacecraft components within desired
limits.
We also wish to minimize the temperature
fluctuation (thermal cycling) that the spacecraft components
are subjected to. FalconSat-2’s internal components, which
are the most thermally sensitive parts of the satellite, are
fairly thermally decoupled from the external heat flux the
satellite is subjected to. This is due to the design with the
inner column and outer structural shell. This allows us to
control the temperature with a passive thermal design
approach. We will modify the thermo-optical properties
(absorptivity and emissivity values) of the external facets of
the satellite so that the satellite and all components are
maintained within the optimal temperature range.
On FalconSat-2, the operational temperatures are limited by
the electronic components within the satellite, and
specifically by the battery. The battery is the most
thermally sensitive of the satellite subsystems because it
cannot be recharged below 0˚C. As a result, the nominal
temperature range targeted for the batteries and internal
components of FalconSat-2 is +5 to +30 deg C. The other
commercial electronics within the satellite have temperature
limits of –40 and +85 deg C. The structural components
and solar panels have much more relaxed temperature
limits. Table 1 lists the temperature limits for FalconSat-2.
Table 1 – Temperature limits for FalconSat-2 subsystems
Subsystem
Figure 1a: FalconSat-2 external view
Antenna
Cadet-built
Solar Panel
Electronics
Modules
MESA Sensors
Orthogrid Al
Structure
Adapter Ring
Battery
COTS
Solar Panel
Figure 1b: Exploded view of FalconSat-2 showing key
features and components
FalconSat-2 has been in development since Fall 2000, with
the Engineering Model completed and tested in Spring
2001. Prior to Summer 2001, however, no work had been
done on the thermal subsystem of FalconSat-2. This need
was addressed over the summer of 2001 through research
conducted at the University of Surrey.
Battery
EPS
MESA
Data Handling
Comm
Solar Panels
Structure
Minimum
Temperature
(˚C)
0
-30
-40
-40
-40
-100
N/A
Maximum
Temperature
(˚C)
50
+50
+85
+85
+85
+110
N/A
To design the thermal subsystem and ensure that FalconSat2 will meet these temperature limits, we had to first simulate
the thermal behavior of the satellite. This will allow us to
see how the satellite will behave without any thermal
control implemented, which will in turn show us what
design we must implement to meet the temperature range
requirements. In order to simulate the satellite’s thermal
behavior, a model had to be created.
We require a detailed thermal model of FalconSat-2 for
several reasons. Primarily, we need to simulate expected
on-orbit thermal behavior of the satellite and ensure that no
spacecraft components exceed their maximum or minimum
temperature limits. We also need to ensure that the
temperature fluctuation (thermal cycling) of all spacecraft
components is minimized. By simulating varying on-orbit
scenarios, including varying attitude modes and varying
subsystem operation modes, we can also simulate worstcase hot and worst-case cold temperature scenarios.
Furthermore, we wish to use the thermal model to simulate
testing environments that we will subject the satellite to at
various phases throughout the development. Furthermore,
we wish to integrate the thermal model into an overall
behavioral model of the satellite to assess the interaction of
the thermal design with the rest of the satellite.
2. MODEL DESCRIPTION
The computer thermal modeling tool was created in MatLab
using Simulink to coordinate the programming. It uses
finite difference analysis to calculate the change in
temperature at each node at every time step. The overall
thermal model is broken up into two parts. The first part
compiles a history of the external flux inputs to the satellite
for a single orbit. The second part of the model then uses
these flux inputs along with the physical makeup of the
satellite to actually perform the finite difference analysis to
calculate the thermal behavior of the satellite throughout the
orbit.
Flux History Calculation Model
The flux history calculation model calculates the heat flux
coming into the satellite due to insolation (direct solar
radiation), Earth infrared radiation, and albedo (solar
radiation reflected off of the Earth). The model calculates
this flux and compiles a history as a function of time for
each of the six faces of the spacecraft.
The inputs to the flux history calculation routine are the
satellite’s epoch classical orbital elements, epoch date and
Universal Time, the satellite’s attitude control method (Suntracking, velocity tracking, tumbling, or quaternions), and
the time of flight taken from the simulation clock. The
outputs are insolation, Earth infrared, and albedo fluxes for
each face with respect to time for an entire orbit.
The flux history calculation model is broken into five
modules within MatLab. These modules, along with their
inputs and outputs, are discussed here:
COE Update--This module updates the classical orbital
elements (COEs) from the epoch time to the current
simulation time. Inputs are the epoch COEs, the epoch date
and time, and the time of flight, taken from the MatLab
simulation clock. This module outputs updated COEs for
the satellite and the current Julian date.
Light--This module calculates the sun position vector, the
satellite position and velocity vectors, and whether or not
the sun currently illuminates the satellite. Inputs are the
current COEs and Julian date. Outputs are the satellite
position vector (R), satellite velocity vector (V), sun
position vector (Rsun), illumination flag (Vis) and
satellite/sun Beta angle.
Surface Normals--This module calculates the surface
normal vectors of each of the six faces of the satellite. This
routine is used if the satellite is sun-tracking, velocitytracking, or randomly tumbling. There is a switch where the
user can choose which tracking mode to use. Alternatively,
the surface normal vectors can be calculated using
quaternions from an interface with Satellite Tool Kit. There
is a switch that allows the user to choose which method of
calculating the surface normal vectors they would like to
use. Inputs are the satellite position vector (R), satellite
velocity vector (V), sun position vector (Rsun), illumination
flag (Vis) and satellite/sun Beta angle. Outputs from the
module are the surface normal vectors for each face of the
satellite, the angle from the +K axis to the satellite R vector
(phi), and the angle from the +I axis to the satellite R vector
(theta).
Insolation--This module calculates the insolation flux on
each of the six faces of the. Its inputs are the surface normal
vectors, sun position vector, and illumination flag. It
outputs the insolation flux on each face in Wm-2 in both
graphical and matrix form.
Earth Effects--This module calculates the Earth Infrared and
Albedo flux on each of the six faces of the satellite. This
part of the model takes the longest time, as there is a double
discrete summation to calculate the Earth IR and Albedo
view factors for each face of the satellite. Inputs are the
surface normal vectors for each face of the satellite, the
satellite position vector (R), the sun position vector (Rsun),
the angle from the +K axis to the satellite R vector (phi),
and the angle from the +I axis to the satellite R vector
(theta). It outputs the Earth infrared and Albedo flux on
each face in Wm-2 in both graphical and matrix form.
Figure 2a shows the Simulink interface of the flux history
calculation model. It takes approximately thirty minutes of
computation time to calculate the flux history for the 92minute orbit of FalconSat-2.
Figure 2a – MatLab Flux History Calculation Model Interface
Figure 2b – MatLab Nodal Temperature Calculation Model Interface
Nodal Temperature Calculation Model
The second main part of the overall thermal model is the
nodal temperature calculation portion. This part of the
model actually conducts the finite difference analysis and
calculates the temperature of each node of the satellite
versus time for the entire orbit. The MatLab interface for
this portion of the model is shown in Figure 2b.
The inputs to the nodal temperature calculation routine are
the flux histories for the satellite calculated in the first part
of the thermal model as well as the lumped parameter
definitions of the thermal nodes throughout the satellite and
the conduction links between the nodes. These lumped
parameters include the mass (m), specific heat capacity (c),
cross-sectional area (A), absorptivity (α), emissivity (ε), and
thermal conductivity values (k). The output from the model
is a temperature profile for each node in the satellite.
The nodal temperature calculation model is broken down
into five modules within MatLab. These modules are
described here:
External Qin--This module calculates the external heat
transfer into each node due to insolation, Earth infrared, and
albedo. The inputs are the flux history matrices compiled in
the previous model. The flux on each facet is then
multiplied by the appropriate surface area and absorptivity
or emissivity value for each node to calculate the external
heat transfer into each node in watts, designated as Qext.
Fourier Conduction--This module calculates the heat
transfer between nodes due to Fourier conduction. It
iterates through each node in the satellite and calculates the
heat transfer to or from every other node. The key equation
in this subsystem, with “i” being the current node of
interest, is equation 1, with the subsystem performing this
summation for each of the nodes in the satellite:
Qcond (i) 
max nodes
 k (i, j)  T ( j)  T (i)
(1)
j 1
Inputs are the temperature of each node and the conductivity
between each node. The module outputs the heat transfer
into each node due to conduction in watts, designated as
Qcond.
Black Body Radiation From Space--This module calculates
the black body radiation coming into each node from the
background of space.
Of course, heat is actually
transferring out of each node to space, so the outputs from
this subsystem will be negative. The heat transfer from each
node to space is calculated using the black body radiation
equation, with the background heat of space assumed to be
4K.
Internal Power Dissipation--This module puts the internal
power dissipated at each node into the matrix form the
model requires. These internal power dissipations are an
input to the overall thermal model.
Finite Difference Analysis--This module calculates the
temperature of each node using finite difference analysis.
The key equation in this subsystem is equation 2:
T 
Qext  Qcond  Qspc  Q int   dt
m  c 
The MatLab thermal model was further verified by
modeling the FalconSat-2 Engineering Model configuration.
This model simulates the Thermal/Vacuum test of the
FalconSat-2 Engineering Model conducted at Kirtland AFB
in Spring 2001. The model uses a 25-node finite differential
analysis model to simulate the thermal vacuum test. A
description of the 25 nodes can be found in Table 2.
Table 2 – FS-2 Engineering Model Nodal Definitions
(2)
Once the Delta T at each node is calculated, the temperature
at each node is calculated by adding the Delta T to the
previous nodal temperature.
3. MODEL VERIFICATION
The MatLab thermal model was verified in two ways. First,
over the summer while it was being developed, it was used
to model Surrey Satellite Technology, Ltd.’s SNAP-1
nanosatellite. The results were then compared to a SNAP-1
thermal model created in Pascal by Dr. Craig Underwood of
SSTL. The model results match exactly. This is to be
expected because the same assumptions, including the
epoch COEs, epoch date/time, tracking mode, and SNAP-1
geometry, were used for both models. Figures 3a – 3b show
the temperature vs. time of each node in SNAP-1 results
from the MatLab model compared with Dr. Underwood’s
Pascal model.
Figure 3a: All 30 SNAP-1 Nodal Temperatures Over 1
Orbit from MatLab Model
Figure 3b: All 30 SNAP-1 Nodal Temperatures Over 1
Orbit from Dr. Underwood’s Pascal Model
#
Node Description
1
-Y Solar Array Panel
2
+X Facet
3
-X Facet
4
+Y Facet
5
-Y Facet
6
+Z Facet (MESA side)
7 -Z Facet (Attachment side)
8
GAS Can Interface Ring
9
+X+Y MESA Sensor
10
-X+Y MESA Sensor
11
-X-Y MESA Sensor
12
+X Module 1
#
14
15
16
17
18
19
20
21
22
23
24
25
Node Description
-X Module 2
+Y Module 1
+Y Module 2
-Y Module 1
-Y Module 2
Batteries
+X Inner Column
-X Inner Column
+Y Inner Column
-Y Inner Column
S-Band Antenna
Whip Antenna
Two main assumptions were used in this model. First, the
aluminum structural facets were assumed to be a single
average thickness (the different thicknesses due to the actual
orthogrid pattern were ignored). This average thickness was
determined from the mass of aluminum in each facet
divided by the density and surface area. The second
assumption was that the heat flux inputs used for the model
were the actual temperature measurements of the thermal
vacuum chamber. The temperature of the chamber was
assumed to be an infinite well at the average temperature of
the eight temperature measurements in the chamber.
The results from the Engineering Model Thermal/Vacuum
test simulation were extremely encouraging. The shapes of
the model’s temperature vs. time curves for each node are
consistent with the shapes of the actual temperature vs. time
curves from the thermal vacuum test.
The actual
temperature values are somewhat off, however. This error
was quantified by calculating the root-mean-square, or
RMS, error of the model as compared to the actual results
for each node. RMS error varied between 2.0 and 3.1 deg C
for each node. The overall model average RMS error was
2.7 deg C. Plots of the model and actual temperature vs.
time curves for three nodes of the engineering model are
found in figures 4a – 4c.
Outer Panel 1 Temperature (deg C)
80
60
RMS =
2.4954
Temperature (deg C)
40
20
0
-20
-40
Actual
Model
-60
0
500
1000
1500
Time (min)
Figure 4a: Model and Actual temperature vs. time curve for
outside of structural panel #1
The model is sufficiently accurate to model and simulate the
thermal behavior of the satellite. It provides a good
conservative snapshot of the thermal behavior of the
satellite, and can now be used to predict on-orbit thermal
behavior and make design decisions.
Inner Panel 1 Temperature (deg C)
60
RMS =
3.0343
Temperature (deg C)
40
20
4. THERMAL DESIGN DISCUSSION
0
Once it was verified to be accurate, the MatLab thermal
model was updated to the Qualification Model configuration
of the satellite. The number of nodes in the satellite
increased to 28 nodes due to the increased number of
subsystems contained in the Qual Model. The Qual Model
nodal definitions can be found in Table 3.
-20
-40
Actual
Model
-60
Analysis of the results of the Thermal/Vacuum test thermal
model indicates that the finite difference analysis method
used provides an accurate representation of the thermal
behavior of the satellite, with a root mean square of less
than three degrees Celsius. The slight discrepancies
between the model and actual data can be explained. The
model’s behavior slightly lags the actual data in most of the
models. This is probably due to the first-order differential
equation approximation used in the model. The lower
temperature limit reached by the model is lower than the
actual lower temperature limit for each node. This is
probably due to the fact that the chamber was modeled as an
infinite well of temperature, when it is actually a finite
surface area radiating to the satellite at a given temperature.
The fact that the model reaches a more extreme temperature
limit than the actual data is acceptable, as we want our
model to be conservative in simulating the behavior of the
satellite.
0
500
1000
1500
Table 3 – FS-2 Qual Model Nodal Definitions
Time (min)
Figure 4b: Model and Actual temperature vs. time curve for
inside of structural panel #1
Transmitter Module Temperature (deg C)
60
RMS =
2.0786
50
40
Temperature (deg C)
30
20
10
0
-10
-20
-30
-40
Actual
Model
0
500
1000
1500
#
Node Description
1 +X Solar Array Panel (Spacequest)
2 -X Solar Array Panel (Spacequest)
3 +Y Solar Array Panel (Cadet-built)
4 -Y Solar Array Panel (Cadet-built)
5
+X Facet
6
-X Facet
7
+Y Facet
8
-Y Facet
9
+Z Facet (MESA side)
10
-Z Facet (Attachment side)
11
GAS Can Interface Ring
12
+X+Y MESA Sensor
13
-X+Y MESA Sensor
14
-X-Y MESA Sensor
#
15
16
17
18
19
20
21
22
23
24
25
26
27
28
Node Description
+X Module 1
-X Module 1
-X Module 2
+Y Module 1
+Y Module 2
-Y Module 1
-Y Module 2
Batteries
+X Inner Column
-X Inner Column
+Y Inner Column
-Y Inner Column
S-Band Antenna
Whip Antenna
Because we wished to predict on-orbit thermal behavior of
FalconSat-2, it is necessary to use both the flux history
calculation and nodal temperature calculation portions of the
MatLab thermal model.
Time (min)
Figure 4c: Model and Actual temperature vs. time curve for
transmitter module box
For the flux history calculation part of the model, several
assumptions were made. First, it was assumed that the
satellite behaves in an attitude mode with the –Z facet
tracking the satellite velocity vector. This is a reasonable
assumption considering the geometry of FalconSat-2 and
the way drag torque will likely affect the attitude. It was
also assumed that the satellite was not spinning at all.
Finally, the epoch COEs were assumed to be the current
COEs of the International Space Station. This is reasonable
given the fact that FalconSat-2 will be carried to orbit by the
Space Shuttle and released in an orbit very similar to the ISS
orbit. The epoch date/time was assumed to be 00:00.00 UT,
21 March 2003, as this is a time frame consistent with the
projected launch date of FalconSat-2.
Baseline Outer Facet Temperatures (deg C)
60
50
40
30
20
10
Side 1 (+Y)
Side 2 (+X)
Side 3 (-Y)
Side 4 (-X)
Side 5 (+Z)
Side 6 (-Z)
0
The nodal temperature calculation part of the model had
several assumptions as well. First, it was assumed for the
first run of the thermal model that the exterior structural
facets of the satellite were bare 6061-T6 aluminum with no
thermal tapes applied. This run would then give us the
baseline thermal behavior of the satellite from which we
could determine the appropriate thermal tape design for
FalconSat-2. The nodal temperature calculation part of the
model used a 12-second time step, and processed 15
iterations of the orbit with an initial temperature of 20 deg C
for all nodes.
-10
0
10
20
30
40
50
60
70
80
90
100
Figure 5b – Baseline predicted outer facet thermal behavior
Baseline Module Box and Battery Temperatures (deg C)
28
26
24
22
20
The results of the baseline Qual Model thermal simulation
are shown in Figures 5a – 5c. Figure 5a shows the behavior
of the four solar panels. As can be seen, they vary in
temperature between –10ºC and +70ºC. Figure 5b shows
the behavior of the exterior structural facets, which can be
seen to vary between –7ºC and +55ºC. Finally and most
importantly, figure 5c shows the thermal behavior of the
internal module boxes and battery. The modules vary in
temperature between +15ºC and +27ºC. The batteries vary
between +13ºC and +25ºC.
These results are very
encouraging, as they show that even with no thermal tapes
applied, the satellite’s components behave well within the
targeted temperature ranges.
Baseline Solar Panel Temperatures (deg C)
80
70
60
50
40
30
20
10
0
Side 1 (+Y)
Side 2 (+X)
Side 3 (-Y)
Side 4 (-X)
-10
-20
0
10
20
30
40
50
60
70
80
90
100
Figure 5a – Baseline predicted solar panel thermal behavior
18
TX
RX
SIM
MIB
OBC
PWR
Batteries
16
14
12
0
10
20
30
40
50
60
70
80
90
100
Figure 5c – Baseline predicted module box and battery
thermal behavior
Because the baseline thermal behavior of the baseline Qual
Model design is already within temperature limits, our
thermal tape design does not need to modify the thermooptical properties of the structure very greatly. Another
consideration in determining the thermal design is the
attitude control requirement of inducing a spin to ensure that
the solar panels receive sunlight. To do this, we will put
thermal tapes with differing absorptivity values on each side
in a pattern that will induce a spin due to torque created by
solar radiation pressure.
After running simulations using the thermo-optical
properties of several combinations of thermal tape, we
decided to use a combination of aluminum and Kapton
thermal tapes. Specifically, we will be using Sheldahl
Second Surface Aluminum Polyimide Tape with 966
Acrylic Adhesive (Item # 146520) and Sheldahl First
Surface Aluminized Polyimide Tape with 966 Acrylic
Adhesive (Item # 146385). The aluminum tape has an
absorptivity of 0.14 and an emissivity of 0.09, and the
Kapton tape has an absorptivity of 0.39 and an emissivity of
0.63.
We ran the thermal model with the thermal tape design
implemented on the Qual Model structure. The results are
very encouraging.
The thermal behaviors of most
components are raised slightly in temperature, but are still
within our desired nominal temperature ranges. The results
of the Qual Model thermal simulation with the thermal tapes
implemented are shown in Figures 6a – 6c. Figure 6a shows
the behavior of the four solar panels. As can be seen, they
did not change greatly from the baseline design and vary in
temperature between –10ºC and +73ºC. Figure 6b shows
the behavior of the exterior structural facets, which can be
seen to vary between –5ºC and +60ºC. Figure 6c shows the
thermal behavior of the internal module boxes and battery,
whose temperature has increased from the baseline design,
but still remains within the desired temperature range of
+5ºC to +30ºC. The modules vary in temperature between
+17ºC and +30ºC. The batteries vary between +16ºC and
+28ºC. These results show that our thermal tape design will
maintain all components of the satellite within the desired
limits.
Module and Battery Temperatures (Deg C)
30
28
Temperature (deg C)
26
24
22
TX
RX
OBC
SIM
MIB
PWR
BATT
20
18
16
0
10
20
30
40
50
60
Time (min)
70
80
90
100
Figure 6c – Predicted module box and battery thermal
behavior with thermal tape
5. CONCLUSION
Solar Panel Temperatures (Deg C)
80
The first stage of the thermal design of FalconSat-2 has
been completed. It will be a passive thermal control
approach, using aluminum and Kapton tape on the outer
structural facets. The design will meet the requirements of
the system, which is to maintain the satellite components
within their required temperature limits. This design has
been verified through the use of the MatLab thermal model
that was developed this summer and verified to be accurate.
70
60
Temperature (deg C)
50
40
30
20
10
0
Side 1 (+Y)
Side 2 (+X)
Side 3 (-Y)
Side 4 (-X)
-10
-20
0
10
20
30
40
50
60
Time (min)
70
80
90
100
Figure 6a – Predicted solar panel thermal behavior with
thermal tape
Additionally, the thermal model will continue to be used in
the design process. The model itself will be integrated with
behavioral models of the rest of the satellite subsystems.
Because the thermal behavior affects the behavior of other
subsystems, and vice versa, we will be able to get a more
accurate prediction of the behavior of the overall system.
Outer Facet Temperatures (Deg C)
60
50
40
Temperature (deg C)
The thermal design process will continue as the FalconSat-2
program progresses. The next step will be the actual
assembly and integration of the thermal tape when the Qual
Model is constructed this November. We will then conduct
the Thermal/Vacuum test for the completed Qual Model,
and see if the results agree with the predicted results from
our model.
30
REFERENCES
20
[1]
10
Side 1 (+Y)
Side 2 (+X)
Side 3 (-Y)
Side 4 (-X)
0
-10
0
10
20
30
40
50
60
Time (min)
70
80
90
[2]
100
Figure 6b – Predicted outer facet thermal behavior with
thermal tape
[3]
Gillmore, David G. Satellite Thermal Control
Handbook. The Aerospace Corporation, 1994.
Gomes, Luis, Prof. Martin Sweeting, and Alex de
Silva Curiel. “Thermal Analysis and Design at
Surrey: UoSAT-12 MiniSat Test Case.”
International Astronautical Federation Congress,
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Stanton, Stuart and Lt. Col. Jerry Sellers.
“Modeling and Simulation Tools for Rapid Space
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Applications.” IEEE Aerospace Conference, 2001.
[4]
Wertz, James R. and Wiley J. Larson, ed. Space
Mission Analysis and Design, 3rd Ed., Boston:
Kluwer Academic, 1999.
Richard Lyon is a first class cadet at the U.S. Air Force
Academy, preparing for graduation and commissioning in
May 2002. Majoring in Astronautical Engineering, he is
the cadet thermal engineer for the FalconSat-2 program.
Jerry Jon Sellers is an active duty Lieutenant Colonel in the
US Air Force. His work experience includes: Guidance &
On-board Navigation Officer, Space Shuttle Mission
Control Center, NASA, Johnson Space Center, Houston,
Texas; Assistant Professor of Astronautics at the USAF
Academy; and Chief, Astronautics for the Air Force
European Office of Aerospace Research & Development,
London, UK. His educational background includes a BS in
Human Factors Engineering from the USAF Academy, an
MS in Physical Science/Astrodynamics from the University
of Houston, Clear Lake , an MS in Aeronautics/Astronautics
from Stanford University and a Ph.D. in Satellite
Engineering from the University of Surrey, UK . Currently
he is the Director of the USAF Academy Small Satellite
Research Center in Colorado Springs, Colorado.
Dr Craig Underwood graduated from the University of York
in 1982 with a BSc(Hons) in Physics with Computer
Science. After gaining a Post Graduate Certificate in
Education (PGCE) from York in 1983, he began a teaching
career at Scarborough Sixth-Form College where he
developed satellite activities. In January 1986, Craig joined
the University of Surrey as a Research Fellow responsible
for the generation and maintenance of software for the
UoSAT Satellite Control Ground-Station. In 1988, as a
Senior Engineer with Surrey Satellite Technology Ltd.
(SSTL), he became responsible for mission analysis and the
thermal design of the UoSAT spacecraft. From 1990 he has
been the Principal Investigator of space radiation effects on
the UoSAT satellites, completing a PhD in this area in 1996.
Craig began Surrey's nano-satellite activities in 1995. In
1993, Craig became a Lecturer in Spacecraft Engineering
advancing to Senior Lecturer in 1999. As well as developing
and teaching Surrey's Spacecraft Engineering post-graduate
and undergraduate courses, Craig also teaches courses on
Astronomy and Astrophysics for the Physics Department at
Surrey. Craig heads the Scientific and Environmental
Remote Sensing (SERS) Group within the Surrey Space
Centre, which has the remit of developing the instruments,
sensors and data processing techniques needed to
investigate the Earth and other planetary environments
from space. His current research interests include the
analysis of the space radiation environment and its effects
on commercial-off-the-shelf (COTS) technologies; optical
and radar satellite remote sensing; space science; machine
vision/ optical navigation sensor development, and nanosatellite technologies.