HokieSat Thermal System
Thermal Model Development and Analysis
December 19, 2002
Michael Belcher
Thermal Systems Lead
mbelcher@vt.edu
Model Architecture
A detailed thermal model of HokieSat has been completed. The HokieSat
geometry is modeled in SSPTA (Simplified Space Payload Thermal Analyzer), a
program available on an educational/evaluation basis from Swales Aerospace.
SSPTA is actually comprised of several programs, which individually calculate
shadowing effects, internal and external radiation couplings, simulates the
spacecraft in its orbit, and calculates heat absorbed from the space environment.
The radiation couplings and absorbed heat loads output from SSPTA are used
as input data to SINDA (Systems Integrated Numerical Differential Analyzer).
SINDA is a computational software package that uses finite difference method to
solve systems of equations for nodal temperatures. SINDA is available in many
formats and versions. The program currently in use is an older version of
SINDA-85, which is freeware.
The external and internal geometry of HokieSat is modeled in SSPTA.
The external model consists of the side skin panels, the nadir and zenith isogrid,
the honeycomb standoff, the magnetometer mount and the nadir solar cell
mounting plate. The external model provides a simplified method of calculating
radiation couplings and orbital fluxes. In addition, the external model allows for
verification of the CEE and battery box models independent of the rest of the
model.
The CEE is modeled in SINDA. An Excel spreadsheet calculates thermal
conduction couplings based on the existing geometry and interface properties.
The CEE model can predict either steady-state or transient temperatures. The
steady-state configuration calculates the temperatures that the CEE would reach
if it were switched on and allowed to thermally stabilize. The transient
configuration takes into account orbit absorbed fluxes and duty cycles of the
electronics. The input parameters to the CEE model are given in Attachment 1.
The battery box model is similar to the CEE model. The battery box input
parameters are given in Attachment 2.
The overall model of HokieSat consists of all components, including the
CEE, battery box, GPS boxes, communications relays, cameras, thrusters, and
rate gyros. The internal SSPTA geometry model calculates radiative heat
exchange between components. The conductive heat exchange is calculated in
Excel from knowledge of the interfaces. The external geometry model provides
orbital flux and radiation heat exchange data. The internal dissipations are
calculated from the power budget. This data is input to SINDA and used to
predict the temperatures of all HokieSat components. The internal thermal
model, not including the CEE or battery box, can be found in Attachement 3
Each of the components are assigned appropriate surface optical
properties of emissivity and absorbtivity. These properties are input into SSPTA
and are used to generate radiation couplings. Table 1 gives the optical
properties used throughout the HokieSat thermal model. A picture of the external
geometry model, taken from SSPTA, is shown in Figure 1.
2
Material
Aluminum 6061
Irridited Aluminum
Delrin AF
GaInP2/GaAs/Ge
Used on
Electronics housings
Zenith and Nadir Panels
Cameras
Solar Cells (Side Panels)


0.2 0.035
0.01 0.11
0.96 0.87
0.92 0.85
Table 1 - Optical Properties
Figure 1 - External Geometery (SSPTA)
Figure 2 - Internal Geometry (SSPTA)
3
Figure 3 - Internal Configuration with Side Panels
SINDA
Geometry
Internal
Rad. Couplings
SSPTA
Surface Properties
Input Data
Orbit Parameters
Int. Dissipations
External
Absorbed Loads
CEE
HokieSat
Cond. Couplings
Output Data
Temperatures
Battery Box
Figure 4 - Model Architecture
4
The internal geometry as modeled in SSPTA is given in Figures 2 and 3.
Figure 4 is a diagram showing how the software works and the interactions of the
models.
Assumptions and Calculation Methods
The thermal mass of each component, or node, is calculated from the
specific heat of its material and its mass. If no mass is available for a given
component, a volume and density estimate is used. The isogrid side panels are
modeled in a similar fashion. A large temperature gradient is not expected across
a side panel. Therefore, each side panel is modeled as one node having a
thermal mass equivalent to a solid piece of aluminum of the same mass.
Conduction couplings are calculated from information about the interfaces
between components. Table 2 gives the different types of interfaces present in
HokieSat and the equations used to calculate conduction.
Bolted
Pressure Contact
Contact
Coefficient (h),
W/m2-K
300 – 1000
15 – 30
Node to Node
N/A
Epoxy/Filler
10,000-15,000
Interface Type
Conduction
Equation (G),
W/K
G  4hAhead Nbolts
G  hAcontact
kAx sec tion
L
G  hAcontact
G
Table 2 - Conduction Equations
In these equations, k is the material thermal conductivity (W/m-K), A is the
applicable area (m2), and L is the path length (m). The values of the contact
coefficients are taken from thermal design references as well as suggestions
from USU and Swales.
The orbit parameters used to calculate incident fluxes are given in Table
3. The orbit is similar to the ISS orbit. Additional information will be needed to
more accurately simulate the orbital conditions directly after separation from the
space shuttle.
Altitude
Inclination
Period
380 km
51°
92.3 min
Table 3 - HokieSat Orbital Parameters
5
A hot and a cold operational case are analyzed. The hot case assumes
that all satellite components are operating within normal parameters, and that the
orbital environment is as listed in Table 4. The cold case assumes the satellite is
operating at minimum mission requirements and that only essential components
are on. A survival case that models the satellite with all components in standby
mode will also need to be developed.
Orbital Fluxes
Solar
Albedo
Earth IR
Hot
1418 W/m2
35%
258 W/m2
Cold
1350 W/m2
25%
216 W/m2
Table 4 - HokieSat Thermal Environment
Blake Moffit from USU has suggested that the nadir and zenith panels be
insulted with Multi-Layer Insulation (MLI). MLI would effectively prevent radiation
from entering the satellite and causing possible unwanted temperature variations
of internal components. The battery box and CEE, which rest on the nadir panel
and would therefore be subjected to albedo and the Earth’s infrared radiation, are
of special concern. Therefore, a SSPTA geometry model has been created that
simulates MLI on the nadir and zenith panels. The results of the MLI model will
be compared to those from the current configuration. The current configuration is
modeled in SSPTA as shown in Figure 3.
Model Results
The CEE and battery box models simulate the steady state temperatures
of the electronics boards and the batteries, respectively. These temperature
predictions are used in the preliminary thermal design phase to identify problem
areas and to provide information regarding the expected temperature gradients
from one component to the next. The steady state results of CEE model are
given in Attachment 1, and the battery box model in Attachment 2.
Early CEE simulations indicated the necessity for an interface filler, such
as indium tape, to provide better heat conduction between the boards and their
surroundings. The model results shown assume that all interfaces have a
conduction coefficient of h = 14,000 W/m2. Without an interface filler, the boards
could reach temperatures almost 100 °C higher than those shown.
The steady state battery box model indicates that the revised thermal
design of the box sufficiently insulates the batteries. As shown, the average
temperature of the batteries varies only 2 °C relative to the battery box frame.
Transient results from the HokieSat thermal model are given in Tables 5
and 6 and the figures in Attachments 3 and 4. The transient results indicate the
temperatures the spacecraft will experience during operation in orbit. The
components will experience the sinusoidal temperature variation depicted in the
figures as the satellite passes through its orbit. The figures depict 5 orbital
periods. The tables give temperature predictions in the form of a maximum, a
minimum, and an average, as well as the operational limits of the components.
6
Cold Case Predictions for Selected Components
Operational Limits
Min
Ave
Max
Cold
Hot
Batteries
43.2
43.8
44.6
5
20
Boards
20.6
24.0
32.0
-40
80
PPU
11.8
19.1
24.7
-55
125
PPT Capacitors
-28.1
-13.1
6.6
N/A
125
Thrusters
-26.8
-10.7
10.4
-40
100
Cameras
-25.2
3.3
53.7
-20
60
Rate gyros
10.3
11.4
12.59
-40
80
Magnetometer
-8.7
20.0
51.2
-40
85
D/L transmitter
16.4
36.3
61.9
-20
70
U/L receiver
6.4
33.3
52.2
-20
70
GPS Filter
-8.3
3.8
20.7
0
50
GPS Isolator
-18.1
-16.2
-14.3
0
50
GPS NCLT
21.4
21.6
21.8
0
50
GPS Preamp
-7.9
4.0
20.1
0
50
GPS Switch
-18.1
-16.1
-14.1
0
50
S-band Preamp
-18.0
-16.2
-14.5
0
50
Table 5
Hot Case Predictions for Selected Components
Operational Limits
Min
Ave
Max
Cold
Hot
Batteries
48.7
49.3
50.0
5
20
Boards
26.0
29.2
37.2
-40
80
PPU
56.4
63.3
68.6
-55
125
PPT Capacitors
-20.4
-6.0
13.1
N/A
125
Thrusters
-20.0
-4.5
16.0
-40
100
Cameras
-21.2
7.3
57.8
-20
60
Rate gyros
15.6
16.7
17.9
-40
80
Magnetometer
-3.1
25.3
56.4
-40
85
D/L transmitter
23.5
42.9
68.3
-20
70
U/L receiver
11.6
37.9
56.5
-20
70
GPS Filter
-4.7
7.2
24.1
0
50
GPS Isolator
-14.2
-12.4
-10.5
0
50
GPS NCLT
45.4
45.5
45.7
0
50
GPS Preamp
-4.2
7.5
23.5
0
50
GPS Switch
-14.3
-12.3
-10.4
0
50
S-band Preamp
-14.2
-12.4
-10.7
0
50
Table 6
7
The results shown are for the case of MLI covering the nadir and zenith
panels. At the time of this report, a SSPTA model of the current configuration
has been developed. However, due to undetermined problems with the software,
the simulation cannot be run and the results are unavailable. Personnel at
Swales have been contacted regarding this problem and a solution should be
forthcoming.
The results from the hot case analysis yield several observations. First,
the batteries are about 30 °C above their hot operational limit. Also, the
temperatures of some of the GPS boxes will actually go below the operational
cold limit. Camera 6 may also operate below the cold limit, as shown in the
relevant figure.
The cold case predictions yield similar conclusions. The temperature of
the GPS boxes, with the exception of the NCLT, will likely dip below the cold limit
over the course of an orbit. Additionally, the average battery temperature will be
above the operational hot limit by about 24 °C. However, the model yields
favorable results for the other spacecraft components.
Parametric Calculations
Parametric calculations are performed to determine the relative accuracy
of the model. The results discussed previously assume values of h = 300 W/m2
for bolted interfaces and h = 15 W/m2 for pressure contact. However, these
values are estimates and represent the low end of a range of possible values.
Therefore, hot and cold case analyses were performed with h = 1000 W/m2 for
bolted interfaces and h = 30 W/m2 for pressure contact. The results are given in
Tables 7 and 8. These results also assume the interface between the battery
box frame and the nadir panel is filled with thermal epoxy.
The resulting variance in predicted temperatures is minimal for most
components. Most component temperatures will vary by +/- 2 °C with a change
in h values. Some low thermal mass components, such as the PPT capacitors,
will experience a larger variation in temperature. For the most part, however,
using lower values of h appears to give a more conservative estimate of
temperatures, as component temperature variations will tend to stay within the
operational envelope.
8
Cold Case Predictions for Selected Components, High h Values
Operational Limits Low h Comparison
T (°C)
Min
Ave
Max
Cold
Hot
Old Ave
Batteries
13.1
17.8
22.7
5
20
43.8
-26.0
Boards
21.6
24.4
31.9
-40
80
24.0
0.4
PPU
6.1
17.0
26.3
-55
125
19.1
-2.1
PPT Capacitors
-22.8
-4.6
25.3
N/A
125
-13.1
8.5
Thrusters
-22.1
-3.4
27.4
-40
100
-10.7
7.3
Cameras
-22.6
5.6
55.7
-20
60
3.3
2.3
Rate gyros
6.87
10.0
12.82
-40
80
11.4
-1.4
Magnetometer
-5.1
20.4
50.7
-40
85
20.0
0.4
D/L transmitter
7.0
30.2
58.6
-20
70
36.3
-6.0
U/L receiver
-3.0
31.6
56.0
-20
70
33.3
-1.7
GPS Filter
-7.8
6.3
28.9
0
50
3.8
2.5
GPS Isolator
-19.1
-16.3
-13.8
0
50
-16.2
-0.1
GPS NCLT
22.9
23.1
23.3
0
50
21.6
1.5
GPS Preamp
-7.7
6.4
29.2
0
50
4.0
2.3
GPS Switch
-19.1
-16.3
-13.7
0
50
-16.1
-0.2
S-band Preamp
-19.1
-16.3
-13.8
0
50
-16.2
-0.1
Table 7
Hot Case Predictions for Selected Components, High h Values
Operational Limits Low h Comparison
T (°C)
Min
Ave
Max
Cold
Hot
Old Ave
Batteries
18.4
23.1
27.9
5
20
49.3
-26.2
Boards
26.6
29.5
36.9
-40
80
29.2
0.2
PPU
31.0
41.6
50.8
-55
125
63.3
-21.7
PPT Capacitors
-17.5
0.3
29.6
N/A
125
-6.0
6.4
Thrusters
-17.3
1.1
31.4
-40
100
-4.5
5.6
Cameras
-18.8
9.7
60.7
-20
60
7.3
2.5
Rate gyros
12.3
15.4
18.2
-40
80
16.7
-1.4
Magnetometer
0.2
25.5
55.8
-40
85
25.3
0.2
D/L transmitter
11.9
34.8
63.0
-20
70
42.9
-8.1
U/L receiver
2.9
36.9
61.2
-20
70
37.9
-1.0
GPS Filter
-4.3
9.6
31.7
0
50
7.2
2.4
GPS Isolator
-15.3
-12.6
-10.1
0
50
-12.4
-0.2
GPS NCLT
46.4
46.6
46.7
0
50
45.5
1.0
GPS Preamp
-4.1
9.7
32.0
0
50
7.5
2.2
GPS Switch
-15.3
-12.5
-10.0
0
50
-12.3
-0.2
S-band Preamp
-15.3
-12.6
-10.1
0
50
-12.4
-0.1
Table 8
9
Recommendations
The GPS boxes will require heaters to maintain them above the
operational cold limit. Currently, 3 W have been allocated to the thermal
subsystem for heater power. Further analysis should be conducted to determine
if this is sufficient for thermal control of the GPS boxes. If heater power is not
sufficient, the possibility of relaxing the thermal constraints on the GPS boxes
should be investigated. However, if this is not feasible, it may be possible to
shunt heat from other components to the boxes with some type of high
conductivity thermal strap.
As discussed previously, all interfaces of the CEE should be filled with a
highly conductive material. In addition to filling the interfaces with a thermal
compound such as indium, it will likely also be desirable to bond the CEE to the
baseplate and isogrid with thermal tape, such as Chomerics Thermattach.
The battery temperature can be regulated by likewise filling the interface
between the battery box frame and the nadir panel. Tables 7 and 8 depicts the
predicted battery box temperatures if the h value between the frame and the
nadir panel is increased to 14,000 W/m2. As shown, the predicted battery
temperature is much closer to the hot limit in this case.
The analysis documented here assumes MLI covers the external surfaces
of the nadir and zenith panels. An additional analysis with the nadir and zenith
panels open to space is highly recommended. However, it is likely that exposing
the internal components to space and to incoming radiation will yield larger
temperature variations than those predicted by the current analysis. Therefore,
the possibility of insulating the end panels with MLI should be investigated.
Swales Aerospace manufactures MLI and may be willing to donate scraps of
sufficient size. If not, Swales could advise Virginia Tech on the fabrication of
such blankets.
Further Analysis
A procedure has been outlined for the necessary thermal testing. The
electronics board must be thermally cycled to test workmanship and functionality.
The boards will be placed in a vacuum chamber available to the department.
The testing requirements dictate three cycles, three hours at each plateau, hot
plateau of 70 °C and a cold plateau of –20 °C. The boards will be heated or
cooled with Thermoelectric Coolers (TEC). TECs are small ceramic devices that
remove heat from one side to the other. The TECs must be used in conjunction
with a heat sink. The heat sink will be liquid cooled and pressurized so that it can
operate in the vacuum chamber. The heat sink has been ordered and is ready
for use. Resistance heaters may be used in conjunction with the TECs for board
thermal control. Thermal testing can begin once the vacuum chamber has been
adequately prepared.
The battery box will also be thermal cycled. The cells must be cold
soaked at –20 °C for 2-4 hours and hot soaked at 40 °C for 2-4 hours. TECs
and/or strip heaters will also be used in this phase of testing.
10
If possible, the thermal models of the CEE and battery box should be
verified with thermal vac testing. This will involve placing thermocouples at
various locations on the CEE and battery box and monitoring their temperature
while operating under vacuum. In addition, the conduction values used in the
model should be verified by measuring the temperature gradient across the
interfaces. If necessary, the models will be updated or modified according to the
results of the thermal testing.
Different orbital cases should be examined. The current model simulates
the standard ISS orbit. However, different orbits are possible depending on the
separation environment. Additionally, the survival environment of HokieSat
directly after launch should be investigated. The thermal model of HokieSat
should be integrated into the current model of the space shuttle payload bay and
the existing ION-F models.
The thermal system will need to be integrated into HokieSat. This
includes preparing the interfaces with the necessary thermal materials. Also,
thermocouples should be placed where they are needed. Heaters will need to be
acquired and integrated. Heater logic subroutines must be built into the software
to ensure adequate thermal control.
11
Attachment 1: CEE Model
Attachment 2: Battery Box Model
Attachment 3: Cold Case Orbital Temperatures
Attachment 4: Hot Case Orbital Temperatures
Attachment 5: Internal Model
12