Heat Storage Panel Using a Phase change Material ...

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

2

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


3

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

]


4

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


5

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


6

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


7

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)


8

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


9

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


11

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


12

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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