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44th International Conference on Environmental Systems ICES2014-119 13-17 July 2014, Tucson, Arizona
Heat Storage Panel Using a Phase-change Material
Encapsulated in a High-thermal conductivity CFRP for
Micro Satellites
Kouhei Yamada1 and Hosei Nagano2
Nagoya University, Nagoya, Aichi, 464-8603, Japan
Yoshinari Kobayashi3
University of Tokyo, Tokyo, 113-8656, Japan
and
Tsuyoshi Totani4
Hokkaido University, Sapporo, Hokkaido, 060-8628, Japan
Thermal control of small, micro and nano satellites is very difficult mainly due to power/
mass restrictions and small heat capacity. In order to satisfy the thermal requirements for
these satellites, it is required to develop new thermal control devices and methodology. In
Nagoya University, a new thermal control device, named heat storage panel (HSP) is
currently under development. The HSP consists of a phase change material (PCM) and a
thin panel-shaped container. The container is made of a high-thermal-conductivity pitch-
based carbon fiber reinforced polymer (CFRP). PCM is used to increase the apparent heat
capacity with a small mass gain. The high-thermal conductivity CFRP is used to enhance
heat dissipation. In other words, the HSP can be used as a heat absorber/heater around the
phase-change point of the PCM, and also be used as a thermal doubler. In this paper,
concept, design, fabrication, and thermal test results of the HSP are presented.
Nomenclature
HSP = Heat Storage Panel
PCM = Phase Change Material
CFRP = Carbon Fiber Reinforced Polymer
DSC = Differential Scanning Calorimetry
E = Flexural modulus [Pa]
L = Span of the sample [m]
b = Width of the sample [m]
h = Thickness of the sample [m]
ΔF = Difference of load between two points[N]
ΔS = Difference of deflection between two points[m]
N = Number of laminated prepreg
X = Size of the space for PCM (Length of one side; pictured in Fig. 8) [m]
P = Pressure [Pa]
V = Volume [m3]
a = Van der Waals constant (1.35× 10-3) [Pa m6/mol]
b = Van der Waals constant (36.6× 10-6 )[m3/mol]
1 Graduate student, Department of Aerospace Engineering, Furo-cho,Chikusa-ku 2 Associate Professor, Department of Aerospace Engineering, Furo-cho, Chikusa-ku, AIAA member 3 Graduate student, Department of Aeronautics and Astronautics, Hongo, Bunkyo-ku 4Associate Professor, Department of Aerospace Engineering, Kitazyusannzyou-Nishi 8, Kita-ku, AIAA member
International Conference on Environmental Systems
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n = Amount of substance [mol]
R = Gas constant (8.314472) [m2kg/s2/K/mol]
T = Temperature [K]
δ = Deflection [m]
w = Uniform load [N/m2]
l = Length of the beam [m]
I = Moment of inertia of area [m4]
I. Introduction
ecently, requirements for cost reduction have inspired development of smaller satellites, called micro and nano
satellites. On the other hand, thermal control of these satellites is very difficult due to (i) limited power
resources, (ii) small heat capacity, (iii) insufficient radiator area, (iv) high density packing of electronics, and (v)
mass limitation. In order to meet such demands, a new type of thermal control device, which is smaller and needs no
electrical power, is required.
As a key technology of such device, heat storage using phase-change materials (PCMs) is very effective. PCMs
can absorb or supply heat as latent heat changing its phase. As latent heat is usually several times larger than
sensible heat within spacecraft’s temperature range, a PCM heat storage device has high energy storage density
around the phase change point. Research on PCM’s has flourished in recent years. For example, NASA has studied
several types of PCM heat storage. Johnson Space Center is studying PCM heat storage for manned spacecraft.1
Water is used as PCM for heat storage. 2 For battery thermal control, Jet Propulsion Laboratory proposed a thermal
storage system using dodecane.3 However, these PCM designes are not suitable for small satellites. They are too
heavy and large for the small satellites. The largest challenge of PCM heat storage device is how to compensate low
thermal conductivity of PCM. Therefore, such devices need a heat conduction member in addition to a strong
container which can bear volume change with PCMs' phase change. This is the main reason why conventional PCM
devices tend to be heavier. Lighter and high-thermal conductivity materials enable a light weight PCM device which
suits smaller spacecraft, and some types of pitch based CFRP may enable such PCM devices. These types of CFRP
have very high thermal conductivity. The thermal conductivity is sometimes higher than aluminum alloys or even
copper. For example, a high thermal conductivity CFRP exhibits a thermal conductivity of 420W/mK in plane.6
Combining PCM with high-thermal conductivity CFRP, a new thermal control device is proposed. The device is
a CFRP panel containing PCM, called Heat Storage Panel (HSP). In this paper, the concept and the design of the
HSP, the results of manufacturing, and the results of thermal tests are provided.
II. Concept of Heat Storage Panel
The HSP is a thin CFRP panel, and PCM is injected into it. The high thermal-conductivity CFRP panel
compensates for the low thermal conductivity of PCM. The HSP's description is shown in Fig. 1. Some components
in satellites, such as the communication system, generate heat periodically, and the heat sometimes changes
temperature of these components drastically. The HSP is attached to such components and moderates temperature
change. The inserted PCM can increase the apparent heat capacity of the HSP, and the larger apparent heat capacity
of the HSP minimizes the temperature change experienced by the attached components. In other words, depending
on the temperature of the HSP, it can work as a heater or a heat absorber. Thanks to this flexible function, it can deal
with abnormal heat without any additional radiators, and it can facilitate the thermal design of spacecraft. Additional
oversized radiators are used on some satellites in order to deal with heat generated temporally. The HSP can be
substituted for these additional radiators. This reduction in radiator can reduce excessive heat loss during cold
temperature periods potentially eliminating or minimizing the amount heater power required
Compared to other PCM heat storage devices, the HSP has the following three features, thinner shape, high
specific strength, and high thermal diffusivity. While utilizing these features, the HSP can be used as a honeycomb
face sheet (shown in Fig. 2) or a thermal doubler not only as a heat storage device. Such multi-functional equipment
is a great advantage for small satellites.
R
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III. Selection of Material Used as PCM
For inserting PCM into the HSP, the most
important feature is the amount of latent heat
it provides. The PCM latent heat for the HSP
should be larger than 200J/g in order to obtain
high heat storage density. Another important
point is to have the appropriate phase-change
temperature; the HSP can function as thermal
controller only when its temperature is around
the PCM’s phase-change temperature.
Considering latent heat and phase-change
temperature, three materials, pure water (H2O),
eicosane (C20H42) and sodium acetate
trihydride (CH3COONa ・ 3H2O), are
identified as candidates for the encapsulated
PCM. These three materials have large latent
heat and they have phase-change temperatures
within the allowable temperature range of
batteries (0°C-40°C), which typically require
the most stringent temperature control of all
components in a satellite.9
Their latent heat and behaviors around
their phase-change temperatures were
examined by Differential Scanning
Calorimetry (DSC). DSC60-A by Shimadzu
Corporation was used in this test. DSC60-A is
a heat flux type DSC. In other words, it
measures latent heat and phase-change
temperature from the difference between the
absorbed flux of a reference material (alpha-
alumina) and the test sample when they are
heated at the same temperature. The difference
of the flux is called “DSC”. As samples, 0.53g
pure water, 4.79g sodium acetate trihydride,
2.37g eicosane are used.
Figures 3 to 5 show the DSC test results of
pure water, sodium acetate trihydride, and
eicosane. The broken line in each figure
indicates a temperature program where the
raising/dropping rate is 2°C/min. Because
sample materials attempt to keep their
temperature and increase heat flux drastically,
0 5000 10000-50
0
50
100
-5
0
5
10
Time[sec]
Tem
per
atu
re[℃
]
DS
C[m
W]
Temp. program
DSC
Figure3. Result of the DSC test; Pure water
0 2000 4000 6000 8000-50
0
50
-15
-10
-5
0
DS
C[m
W]
Tem
per
atu
re[℃
]
Time[sec]
DSC
Temp. program
Figure 4. Result of DSC test; Sodium acetate trihydride
0 2000 4000 6000 8000
-40
-20
0
20
40
-10
0
10
DS
C[m
W]
Tem
per
ature
[℃]
Time[sec]
Temp. program
DSC
Figure 5. Result of the DSC test; Eicosane
Figure 1. Description of the HSP
Casing Layer
Phase-Change Material
Cover Layer
Cover Layer
Prepreg
Figure 2. HSP honeycomb panel
Hea
t fl
ow
[mW
] H
eat
flo
w[m
W]
Hea
t fl
ow
[mW
]
International Conference on Environmental Systems
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the DSC curves peak around the phase-change temperature.7 The height of DSC peaks indicates the quantity of the
latent heat, and the temperature at DSC peaks indicates the phase change temperature. The phase-change
temperature and latent heat of each material obtained from this test are shown in Table 1. There are large gaps
between freezing point and melting point of pure water from Fig. 3. The same behavior can also be seen in Fig. 4. It
is considered that such phenomena occurred because of supercooling. Furthermore, sodium acetate trihydride
dehydrated at high temperature. Based on these results, it is difficult to make pure water or sodium acetate trihydride
change phase at accurate/consistent temperatures. Eicosane also exhibits supercooling. However, the supercooling of
eicosane is much smaller than that of the others. The eicosane phase-change temperature was more constant and
precise than others. From this result, eicosane was adopted as the PCM to be inserted into the HSP. The thermal
conductivity of eicosane is about 0.42W/mK around ambient temperature.8
Table 1. Phase change point and latent heat10 11 12
Material Phase-change temperature[°C] Latent heat[J/g]
Solid→Liquid Liquid→Solid Reference Solid→ Liquid Liquid→Solid Reference
Pure water -0.720 -20.5 0.00 269 349 334
Sodium acetate
trihydride
59.3 None 58.0 272 None 264
Eicosane 36.6 34.9 36.4 232 221 247
IV. Structural Design
Manufacturing Process of the HSP A.
Before discussing the structural design of the HSP, the manufacturing process is described below.
First, CFRP panel for the HSP is molded by an autoclave method. For the HSP, NT91500-525S prepreg
produced by Nippon Graphite Fiber Corporation was used in this study. This prepreg was made of mesophase pitch
carbon fibers and epoxy resin. The thermal conductivity of this prepreg was measured by AC calorimetry13 to be
347W/mK in the fiber direction and 3.0W/mK in the cross-fiber direction. As shown in Fig. 1, HSP CFRP structure
mainly consists of three parts: two cover layers, and one casing layers. One cover layer and the casing layer are
integrally molded and the other cover layer is molded alone. In this step, one dish-shaped panel and one flat square
panel, as shown in Fig. 6, are manufactured.
Next, injection needles are embedded at the edge of the casing layer. After that, the panel with injection needles
is covered by the flat square panel as shown in Fig. 7. Both needles and covering panel are fastened by epoxy resin.
Finally, eicosane in liquid state is injected through these needles. Once the eicosane has been injected, the
needles attached to the panel are cut at the edge of the HSP, and the opening is sealed by epoxy resin, DENATITE
XNR/XHR6815. The HSP was exposed to vacuum for a time period of 220hours. During that test, the weight of the
sample did not decrease. The HSP were able to be heated to 80°C without seeing any impact on the sealing.
Process of Structural Design B.
The performance of the HSP is going to be demonstrated on a small satellite, HODOYOSHI-4. Therefore, the
size of HSP must be within 150g. Moreover, the shape must be a 15cm by 15cm square. Under such conditions, the
Figure 6. Molded CFRP panel for HSP
(Left; Cover panel / Right; Casing panel)
Figure 7. Molded CFRP panel for HSP
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appropriate number of the laminated prepreg was determined by the results of strength tests. The process is
described below.
First, the thickness of the casing layer was decided from the amount of PCM needed. Because CFRP panels are
molded by autoclave method, the thickness of CFRP panel dictates the number of laminated prepregs. The size of
that part also depends on the amount of PCM. The following was used to estimate the amount of PCM to use. The
HSP is expected to be applied to the thermal control of nanoscale satellites communication subsystems. For example,
the transceiver of modern 1.5U Cubesat, Edison Demonstration of Smallsat Networks (EDSN) satellite, gives off the
maximum 8.2W during 6.3% of orbit on.9 It was determined that the HSP heat storage amount is around 2kJ
referring to some satellites including EDSN. By weight allowance of HODOYOSHI-4 satellite and the heat storage
requirement, the total amount of internal PCM (eicosane) was determined to be 10g.
Next, the thickness of the cover layer is selected to resist excessive flexure caused by expansion of the PCM and
depends on the flexure modulus of the CFRP. The required flexure limit for the cover panel depends on the volume
of the HSP’s internal room. On the other hand, the flexural modulus of CFRP is estimated from the result of three-
point bending tests.
Three-point Bending Test C.
Three-point bending tests were conducted in order to estimate relation between the flexural modulus and the
thickness of the panel, the number of laminated prepreg. A strength test machine, AG5000B manufactured by
Shimazu Corporation, was used in this test. From the test result and equation (1), the flexural modulus on each
number of laminated prepreg is determined.
S
F
bh
LE
3
3
4 (1)
Equation (1) is derived from the relation expression between load and deflection amount.
Three types of samples are prepared for this test; 4, 6 and 8 layers. All samples are dual directional materials.
(0degree and 90degrees) The test result is shown in Table 2. From the test result, an approximate straight line is
calculated by using a least square algorithm. The flexural modulus of each case is derived from the inclination of the
straight line. Hereby, equation (2) is determined to expresses the relationship between flexural modulus and the
number of laminated prepreg is determined.
822.259.20 NE (2)
Thickness of Casing Layer D.
The casing layer is pictured in Fig. 8. The space for Eicosane is divided
into 4 rooms to keep the panel strength. Each two rooms are connected by
5.0mm width passage to facilitate PCM injection. By changing parameter “X”,
the number of the laminated prepreg required to insert 10g eicosane was
calculated. However, air is also encapsulated, and it is difficult to fill up the
HSP only with eicosane. Therefore, the total amount of eicosane is reduced to
9.8g. The size of eicosane space is designed to be 5.6% larger than the volume
of 9.8g liquid eicosane to allow mixing of air into eicosane. When the
parameter X is 5.5cm, the void fraction of liquid eicosane is 5.6%. The result
of this calculation is listed in Table 3. Figure 8. Model of casing layer
X cm
15cm
15cm
Table 2. Lists of the three-point bending test results
Number of prepreg;
N
Width ;
b[mm]
Thickness;
h[mm]
Flexural modulus;
E [GPa]
Average
[GPa]
4 14.4 0.393 74.9 70.5 72.7 72.7
6 13.7 0.587 129 126 126 127
8 13.7 0.826 153 153 153 153
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Thickness of Cover Layer E.
When eicosane is inserted into the HSP, air can be also injected into internal room of the HSP. As the temperature
of the HSP increases, the air wants to expand putting pressure on the cover panels. The uniform load given by this
pressure may bend the cover panels. Flexural modulus required for the cover panels is estimated from allowable
flexure. A more detailed explanation of the process is given below. By thermal fluid analysis software, “Thermal
Desktop”, the maximum temperature of the air, which is in internal room of the HSP on HODOYOSHI-4, was
examined. According to this analysis, the maximum temperature is 132.6°C. By substituting this result into Eq. (3),
the pressure of the air in the internal rooms can be obtained.
nRTnbVV
anP
)(
2
2
(3)
The cover panels are pressed by internal air pressure. This situation is similar to a beam fastened at both ends
when uniformly distributed load is applied on it. The parameter “X” is corresponding to the beam span “l”. The
description of the beam model is shown in Fig. 9. Thus, the amount of the cover panel deflection is obtained from
Eqs. (3) and (4), which are the relation expression correlation uniformly distributed load and the amount of
deflection.
EI
wl
384
4
(4)
In Eq. (4), parameter “δ” indicates the maximum amount of the deflection which occurs on the center of a beam
in the beam model. This “δ” corresponds to the deflection of HSP panel right over the PCM space. The relationship
between the amount of deflection and the number of laminated prepreg is shown in Table 4. This relationship is
obtained from Eqs. (2), (3) and (4). X corresponds to width in Fig. 8. The HSP is aimed to be fixed on satellites. So,
it is better that the amount of deflection is as small as possible. Thus, the allowable amount of deflection was set at
1mm. As a result, the minimum number of laminated CFRP was calculated as shown in Table 5.
Table 3. Relationship between X and the number of prepreg
X[cm] Number of prepreg
(0.1mm/layer)
2.5 50
3.0 35
3.5 26
4.0 20
4.5 16
5.0 13
5.5 11
6.0 9
x
δ
w
HSP(cross section) Figure 9. Description of beam model
Table 4. Relationship between the amounts of deflection
and the number of laminated prepreg
X[cm] The amount of deflection
0.5mm 1.0mm 1.5mm 2.0mm 2.5mm
2.5 6 5 5 4 4
3.0 7 6 5 5 5
3.5 7 6 6 5 5
4.0 8 6 6 6 5
4.5 8 7 6 6 6
5.0 8 7 7 6 6
5.5 9 7 7 6 6
6.0 9 8 7 7 6
Table 5. Number of laminated CFRP
The amount
of eicosane
Number of laminated prepreg
Casing
panel
Cover
panel Total
9.8g 11 7×2 25
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V. Demonstration on HODOYOSHI-4 The HSP is going to be tested on HODOYOSHI-4, which is a small satellite to be launched in 2014. The aim is
to precisely evaluate HSP's use in space by comparing the orbit temperature to the predictions through the thermal
analysis. During this experiment, the HSP is heated by sunlight and it is planned to examine PCM's effects on the
temperature range by measuring temperature of the HSP through one thermistor attached to back side, opposite to
the side directly heated by the sun.
Originally, the HSP is aimed to be used as an equipment panel and to prevent the temperature of attached
components from changing drastically. However, in this case, the HSP is to be fastened to the outer insulation of
HODOYOSHI-4. In other words, the HSP is outside of the satellite and isolated from HODOYOSHI-4 itself. This is
why the heat generated by sunlight is to be used instead of the heat generated by a component heating in the
satellites. The difference between actual use and demonstration on HODOYOSHI-4 is described in Fig.10.
Therefore, the optical property of the HSP’s surface needs to be adjusted in order to absorb appropriate quantity of
heat. It is necessary for the temperature range of the HSP on orbit to cover the phase-change temperature of eicosane
and not to exceed the allowable temperature range of the HSP. In order to satisfy such requirement, solar absorbance
and hemispherical total emissivity of the HSP were adjusted by choosing the adequate material which covers the
sunlit surface. The detailed determination process of the HSP surface's optical property will be mentioned later.
The HSP flight model is mounted to the aluminum frame with Velcro fastener and this frame is fixed on outer
multilayer insulation (MLI) of the satellite. In Fig. 11, the appearance of HODOYOSHI-4 and the location of the
HSP are shown. The mounting position of HSP is the -Z side of HODOYOSHI-4. This is because heat load by
sunlight on this side can be precisely calculated.
VI. Thermal Analysis of HSP for Flight Model Design
A. Configuration of the Analytical Model
In this chapter, the analytical model of the HSP is described. Using Thermal Desktop, a model was created in
order to predict the HSP temperature on orbit. This model consists of the HSP and a baseplate which simulates the
outer panel of the satellite. The whole surface of the base plate is insulated by MLI. In this model, the complex HSP
assembly of CFRP, eicosane and aluminum frame was simplified. The whole panel including the PCM is considered
to change phase, and the inner structure of the HSP is neglected. However, the heat capacity and radiation area of
this simplified analysis model is the same as that of
the actual flight model. Density, specific heat,
thermal conductivity of this model is calculated by
mixturing rule by volume. The orbit used in this
analytical model is shown in Table 6.
HSP
PCM Heating
Components
CFRP
Actual Use
PCM
Surface
Material
Sunlight
Demonstration
Table 6. Orbit of HODOYOSHI-4
Altitude 636km
Beta angle 17~28degree
Position Sun directed (Cant angle 20degree)
Figure 10. The difference between actual use and
demonstration on HODOYOSHI-4
Figure 11. HSP flight test model attached on
HODOYOSHI-4
(@University of Tokyo Nano-Satellite Center)
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B. Result of Model Analysis
The analysis is conducted to select the surface material for the HSP. Several cases were evaluated. The model of
each case has different surface optical properties. The two design objectives are:
- The temperature range on orbit
must cover the phase-change
point of eicosane (36.9°C).
- The maximum temperature on
orbit must be under 90°C,which is the upper limit
temperature of the HSP. The
maximum limit point is
defined from the glass
transition temperature of
epoxy polymer.
The candidate materials shown
in Table 7 and others were
evaluated. The two thermal
requirements mentioned above are
satisfied when the ratio of solar
absorbance to emissivity, α/ϵ,
ranges from 0.42 to 0.71. The
result is shown in Fig. 12.
According to the result of this
analysis, aluminized Upilex film
is adopted as the material for the
HSP's external surface. In this
case, the phase changing time is
about 540 seconds (from solid to
liquid), and 220 seconds (from
liquid to solid). During 13 % of
one cycle, the temperature of the
HSP is kept constant.
VII. Performance Evaluation (Vacuum Test 1)
Thermal performance of the HSP flight unit was tested in a space chamber. Two thermal test campaigns were
conducted. In the first campaign, two types of vacuum tests were conducted and the amount of heat storage and
hysteresis were evaluated. Moreover, the temperature range of two panels, with PCM and without PCM, was
compared and effect of encapsulated eicosane was examined. In the second campaign, we simulated the
demonstration on orbit, and the result was matched to the computational model. In this chapter, the first vacuum test
is mentioned.
A. Experimental Apparatus and Sample
The aim of the test is to show the effect of injected PCM. Therefore, two samples, with (9.8g) and without
eicosane are tested. The temperature of the samples is measured by copper-constantan thermo couples attached to 9
point (sample1 and 2) per a sample on the surface. Fig. 13 shows the location of the thermocouples and Table 8
indicates their properties. All samples are enveloped by Upilex film and are attached to the base by Velcro. Two
samples of the HSP, as shown in Fig. 14, were prepared for the thermal vacuum tests. This thermal vacuum test is
conducted in space chamber. The shroud of the space chamber used in this test has 950mm internal diameter and
1135mm depth. The chamber can be pumped down to 10E-7 Pa. During a test, the shroud temperature is kept below
-180°C.
15000 20000 25000 30000
-20
0
20
40
60
80
100
120
140
Time[s.]
Tem
per
ature
[deg
.C]
UpilexCFRPAl teflon30%CFRP+70u teflonAu teflon
CFRP Upilex
Al Teflon
30%CFRP+70% Au_Teflon
Au_Teflon
Figure 12. Relationship between temperature and surface material
Table 7. Optical properties5
Materials
Solar
absorbance
α
Hemispherical
total emissivity
ϵ
α/ϵ
CFRP 0.93 0.85 1.09
30%CFRP+70% Au_Teflon 0.43 0.63 0.69
Aluminized Upilex 0.38 0.68 0.56
Gold evaporation Teflon 0.22 0.53 0.42
Aluminized Teflon 0.14 0.65 0.22
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B. Heating Program Samples are heated by a heater attached to the sample’s top surface which is the opposite side to base plate. This
heater has the same size of HSP's top face which it fully covers. By this heater, two types of heat loads, program
heating and cyclic heating, are applied to the samples. The first heat load is hypothetically assumed as a typical heat
program of a satellite, and the program is set intentionally in order to make eicosane change its phase continuously.
The second one is cyclic heating program, which repeats a max 20W sine heating ten times.
C. Result of Tests
The test result of the program heating test is shown in
Fig. 15. This is the temperature of the HSP's center (back
side). The red line indicates the temperature of the HSP
with eicosane and blue line indicates that of the HSP
without eicosane. While the eicosane changes its phase, the
temperature of the HSP stays constant in this test.
Moreover, the temperature range difference between the
maximum temperature to the minimum temperature is
13.3°C with eicosane while it is 36.1°C w/o eicosane.
The cyclic heating test result is shown in Fig. 16. This figure shows the result from the 4th cycle to 6th cycle in
ten cycles. In this test, ten cycles of heating and cooling, no hysteresis effect was observed. In other words, the HSP
changes its phase at the same temperature and consumes the same time for its phase change every time in ten cycles.
Figure14. Samples in space chamber
Table 8. Specifications of samples
Sample1 Sample2
Size[mm] 150×150×27
Weight (except frame and films)[g] 89.6 99.7
Amount of eicosane[g] 0 9.8
Total latent heat[J] (Estimated from DSC) 0 2215
Surface material Aluminized Upilex film (thickness 50μm)
TC 9 points
Table 9. Temperature range of program
heating test
Amount of eicosane[g] 0 9.8
Maximum Temperature[°C] 50.0 36.9
Minimum Temperature[°C] 15.9 23.6
Range[°C](Max.Temp.-Min.Temp.) 36.1 13.3
Difference of range[°C] 22.8 (63%)
Eicosane
Heater
Top (faced to the Sun )
Back (faced to the base)
Figure 13. Location of thermocouples
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D. Calculation of Effective Latent Heat
From the result of the program heating test, the total amount of heat absorbed by encapsulated eicosane was
estimated. A latent heat of PCM can be estimated from the time consumed by phase-change. This method is
introduced in reference 2. In this method, the latent heat is calculated from the data obtained under two different heat
loads. In this test, the phase of eicosane changes solid to liquid under 8 conditions. Therefore, 28 different estimated
values could be deduced from this test. The average of estimated value is 208J/g, which is about 8% lower than the
value from DSC test. Therefore, the total heat storage amount of the HSP is 2047J.
VIII. Flight Model Test and Model Matching (Vacuum Test 2)
The second campaign was conducted in order to examine the HSP temperature change on orbit, and to match the
calculation and the experiment. This test campaign consisted of two types of the tests, one was constant heating and
the other was cyclic heating test simulating on orbit conditions. The results of these tests were compared with the
results of the thermal analysis.
A. Experimental Apparatus and Sample
The sample of this campaign and flight model for HODOYOSHI-4 is much the same. However, the sample of
the thermal test has four thermocouples on its surface, instead of the thermistor on flight model. The experimental
apparatus including the space chamber is almost the same with the thermal vacuum test 1 except: the base plate
temperature is kept constant at 20°C and the thermocouples are covered by aluminized polyester films to prevent
radiation leakage.
2000 4000 6000 8000 1000020
30
40
50
60
0
5
10
15
20
25
30
Time[s.]
Tem
per
ature
[deg
.C]
Hea
t lo
ad[W
]
w/o eicosane with eicosane heat load program
Figure 16. Result of cyclic heating test
-10
0
10
20
30
40
50
-110
-90
-70
-50
-30
-10
10
30
50
70
90
25000 30000 35000 40000
Hea
t L
oad
[W]
Tem
per
atu
re[d
eg.C
]
Time[s.]
w/o eicosane with eicosane heat load
Figure 15. Result of program heating test
International Conference on Environmental Systems
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B. Measurement of Equilibrium Temperature
The HSP equilibrium temperatures were measured at 7/9/11/15W stationary heat loads. The comparison
between the test and the analysis is shown in Table 10. All of the temperatures indicated in Table 10 are the average
of all thermocouples or nodes. From this comparison, it is confirmed that the amount of heat radiation of the
computational model is almost same with that of the actual flight model. When the 9W heat load is applied to the
HSP, the HSP radiate 82% of total heat load from the
top (the side directed the sun), 10% from the edges and
the other 10% from other sides including the back (the
side directed toward the base). In this analysis, the heat
conducted to the base plate is neglected, because the
HSP’s surface do not make contact with the base plate
and the heat is conducted only by Velcro. That heat is
accounted for less than 1 % of heat load.
C. Cycle Test
The on-orbit heating by sunlight is
calculated and the same heating program
is applied to the sample. The result of this
test is shown in Fig. 17. The red broken
line is the result of analysis, which
simulates the thermal vacuum test. The
curve obtained from the test is generally
corresponding to that from analysis. This
test has shown that the calculation model
used for flight model design is correct.
IX. Conclusion
A new thermal control device, Heat Storage Panel (HSP), for small satellites application is proposed in this
study. The HSP is combined with a phase-change material, eicosane, and a high-thermal-conductivity carbon fiber
reinforced polymer. The HSP absorbs the peak heat of satellites’ components and moderates temperature change
around the phase–change point.
The HSP design was conducted in two steps, the selection of PCM and the panel structural design. Eicosane
was adopted as PCM for HSP through DSC testing, because its phase change temperature was the most stable of that
of three candidate materials and its latent heat met design targets. In the structural design, the thickness of the HSP
was optimized in order to achieve sufficient strength to bear the encapsulated air expansion yet minimize mass.
After that, the HSP was manufactured and tested in the space chamber. In the thermal vacuum test, the HSP’s
effectiveness at moderating temperatures and storing heat was evaluated and compared with the panel without PCM.
According to this test, the total heat storage capacity was estimated at 2047J and the HSP with encapsulated
eicosane provided clear advantage over the HSP without.
The HSP’s performance will be demonstrated on the small satellite, HODOYOSHI-4 to be launched in 2014.
The demonstration will rely on solar heating to heat the HSP. Consequently, computational thermal analysis was
used to chose aluminized Upilex as the surface material and evaluate the HSP temperature profile on orbit. That
computational analysis matched data from thermal vacuum testing of the HSP flight test unit.
Acknowledgments
This study was supported by “Funding Program for World-Leading Innovative R&D on Science and
Technology” , which was planned by the Council for Science and Technology Policy in the Cabinet Office, through
the Japan Society for the Promotion of Science. Advice and comments were given by Mr. Ichiro Mase (Next
15000 20000 25000 30000-20
0
20
40
60
Time[s.]
Tem
per
ature
[deg
.C]
Analysis Experimental
Figure 17. The comperison betwee calculation and experiment
Table 10. The HSP equilibrium temperature
Average value of all points 7W 9W 11W 15W
Experimental results[°C] 5.5 23.1 38.4 64.2
Caluculation results [°C] 5.6 23.0 37.7 62.1
International Conference on Environmental Systems
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generation Space system Technology Research Association; NESTRA) and Mr. Koji Yamaguchi (Orbital
Engineering Inc.) has been a great help in this study. We are deeply grateful to them.
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