Thermal Design and On-orbit Validation of the First

47th International Conference on Environmental Systems ICES-2017-130 16-20 July 2017, Charleston, South Carolina

Thermal Design and On-orbit Validation of the First

Philippine Micro-satellite: DIWATA-1

Delburg Mitchao1 Tsuyoshi Totani2 Masahi Wakita3 Harunori Nagata4

Hokkaido University, Sapporo, Hokkaido, 060-8628, Japan

and

Yuji Sakamoto5

Tohoku University, Sendai, Miyagi, 980-8579, Japan

A thermal design procedure is presented based on methods applied to DIWATA-1

without performing arduous thermal tests. DIWATA-1, the first Philippine-made satellite,

was deployed through the International Space Station (ISS). It follows a near-circular orbit

with a nominal altitude of 400 km and an inclination of 51.6º. The micro-satellite has a

volume of 55 x 35 x 55 cm3 with no solar paddles and uses attitude control during

communication, download, and observation modes only. When the satellite is under one of

these modes, the typical average power consumed is greater than generated power. Given

the surface area and power limitations, passive thermal control was deemed necessary to

guarantee a successful mission under extreme environments resulting from the varying orbit

beta angle. Conductive insulation between inside panels and outside solar-cell-mounted

panels with Kapton polyimide surface finish was implemented on two thermal models: a

multi-nodal and a two-nodal model. Surface finish area of the external panels were varied

until temperature of components satisfied the allowable operating temperature ranges based

from pre-flight analyses using an originally developed thermal analysis tool and Thermal

Desktop/SINDA/Fluint. By applying these steps, development time was reduced into around

one year. After its deployment in space, the on-orbit temperature and pre-flight analyses

results using the same orbital elements were compared to validate the design approach and

determine the sufficiency of the numerical model.

Nomenclature

b = Beta Angle

δs = Declination of Sun

Ωs = Right Ascension of Ascending Node of Sun

Ω = Right Ascension of Ascending Node of Satellite

w = Argument of Perigee

h = Altitude of Satellite from the Surface of the Earth

Re = Radius of the Earth

RI = Orbit Inclination

TLE = Two Line Element

TMM = Thermal Mathermatical Model

1 Graduate Student, Department of Mechanical and Space Engineering, [email protected] 2 Associate Professor, Faculty of Engineering, [email protected] 3 Assistant Professor, Faculty of Engineering, [email protected] 4 Professor, Faculty of Engineering, [email protected] 5 Associate Professor, Department of Aerospace Engineering, [email protected]

International Conference on Environmental Systems

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I. Introduction

or the past few years, many countries have started to notice that space is an all-encompassing field. It may

include fundamental things such as agriculture, disaster management, environmental management, and national

security. This fact has sparked the interest of many to build their own satellite suiting their needs. Nowadays, even

third world countries have started venturing out in space.

One example is DIWATA-1, the first satellite made by the Philippines. DIWATA-1, whose development started

in December 2014, was an initiative by the Philippines’ Department of Science and Technology (DOST) in

partnership with Tohoku University and Hokkaido University. Its mission was to observe the Philippines by

covering various remote sensing applications including calamity assessments, oceanographic studies, agricultural

productivity, and urban planning. To address these, it has four cameras onboard: a 3-meter resolution NIR High

Precision Telescope (HPT), 80-meter Space-borne Multispectral Imager (SMI) with two liquid crystal tunable filters

(LCTF), a 7-km resolution Wide Field camera, and a middle-field wide-view CCD camera. The latter is used to

assist in attitude control and target pointing. In addition, a three-axis attitude control system consisting of 4-reaction

wheels, gyro sensors, sun aspect sensors, magnetometers and star sensors are utilized to observe points and locations

of interest1.

On March 23, 2016, the Atlas V rocket was launched from Cape Canaveral, Florida with DIWATA-1 as one of

its payload. A month later, it was released from the International Space Station through the Japanese Experiment

Module (JEM) Orbital Deployer (ISS-50)2. The release from the orbital deployer designed by the Japanese

Aerospace Agency (JAXA) is shown in Figure 1. In effect, the orbit inclination, initial altitude, and eccentricity of

DIWATA-1 is similar to ISS at approximately 51.6º, 410 km and ~0.001, respectively.

Satellites operating at low-Earth (400-1000km) orbits with low to medium inclinations are exposed to extreme

thermal environments. Direct sunlight, an external heat source, primarily influences the temperature rise and drop in

a satellite. A satellite’s exposure time to the sun is highly dependent on the angle between the solar vector and its

orbit plane. For satellites in non-sun-synchronous orbits, this angle varies all throughout the year and the satellite

subjected to sunshine for several days without eclipse may arise. In addition, heat sources such as albedo and Earth

irradiation are found to be significant in heat transfer calculations especially for near-circular orbits3. These factors

might affect satellite operation making thermal control a crucial task. However, replicating the thermal environment

in space on Earth is difficult due to complexity and testing is time-consuming. Therefore, using software to simulate

the thermal environment and analyzing temperatures of objects in space is preferable for a quicker development

time.

Figure 1. Release of DIWATA-1 from the ISS taken by Astronaut Tim Peake4

II. Structure and Component Layout

DIWATA-1 has a dimension of 55 cm x 35 cm x 55 cm, with a mass of around 53 kg. Its structure consists of

milled panels made of aluminum alloy A7075-T7351. Figure 2 shows the configuration of internal and external

panels of DIWATA-1. Inside, two panels are fixed to a 25-mm thick main central panel. Whereas on the outside, six

10-mm thick external panels are interconnected by four guide rails. DIWATA-1 has no solar array paddles. Thus,

the satellite relies solely on 19 strings of body-mounted photovoltaic (PV) cells for electrical supply. These PV cells

are capable of generating an average power of 39 W over the average sunshine time of 54.6 min/rev.

F


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Figure 2. DIWATA-1 Panel Structure.

All internal components are shown in Figure 3. These are categorized into four systems mainly Communication,

Payload, Power, and Attitude Determination and Control. Most of these components were mounted to the three

internal panels to increase the effective radiative area internally. Whereas, Figure 4 presents the different

components mounted on the outside panels. The UHF, S-band, and X-band antennas located at +Z Panel and -Z

Panel are used to communicate with the satellite, send commands, and download data.

Figure 3. Internal component layout of DIWATA-1.

Figure 4. External component layout of DIWATA-1. S: S-band; X: X-band; U: UHF; ANT: Antenna


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III. Thermal Design and Parameters

A. Space Environment

The principal forms of environmental heating in space are direct sunlight, sunlight reflected off the Earth, and

infrared (IR) energy emitted from the Earth. These forms of heat may vary by season. Table 1 summarizes the

possible worst case parameters a satellite may experience on-orbit. Maximum values were associated to hot case and

minimum to cold case.

Table 1. Space thermal environment. 5-6

Worst Hot Case Worst Cold Case

Solar constant, Gs [W/m2] 1414 1326

Albedo factor, a 0.4 0.25

Earth IR radiation, qIR [W/m2] 258 216

Initial temperature [˚C] 25 10

Initial position N/A Shadow Entry

B. Orbital Parameters

For low-earth orbit and non-sun-synchronous satellites, orbit beta angle was introduced. It is the minimum angle

between the orbit plane and the solar vector. Using the software, Systems Tool Kit (STK) by Agility Graphics Inc.7,

the beta angle history of the International Space Station was identified. From Equation 18, it was established that at

|b| >70.2º, no-eclipse phase starts. This phenomenon may last for two to four consecutive days and was treated as

worst hot case.

b = sin-1[Re/(Re+h)] (1)

Figure 5. Beta Angle History of International Space Station (ISS).

Using beta angle, orbital elements such as right ascension of ascending node, argument and periapsis were

estimated. Meanwhile, inclination, altitude, and eccentricity were taken from the International Space Station’s TLE9.

Table 2. Orbital parameters.

Worst Hot Case Worst Cold Case

RI, deg 51.6

Ω, deg 180 90

ω, deg 0

Beta Angle, deg ~75 ~15

Altitude, km 400

Eccentricity 0.0001

Date of Simulation 2016 Dec 21 2016 Jun 21

C. Internal Heatload

In space, thermal conditions of DIWATA-1 may differ due to the varying power consumption. The total power

expended is highly dependent on the mode used. The various operation modes are summarized in Table 3 and are

grouped into four categories: Communication, Download, Observation, and Standby. During communication,

normal mode with attitude control is used. For observation mode, payloads such as High Precision Telescope (HPT),


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Spaceborne Multispectral Imager (SMI), Wide Field Camera (WFC), etc. are additionally turned on. In download

mode, X-band transmitter may be activated as supplement to the S-band transmitter to achieve faster download

rates. Meanwhile, the satellite runs under standby mode when not in contact with the ground station. In this mode,

attitude is not controlled causing the satellite to tumble. Tumbling was approximated at around 0.8 deg/s. Given this

value and for simplification of the thermal design, an effective internal heatload was estimated. For WHC, a

constant 25 W heat load was applied to the internal panels for the whole duration of simulation. On the other hand, 5

W was used for WCC calculations.

Table 3. Power consumption of DIWATA-1 at different modes.

D. Thermal Design

Passive control was implemented by conductively isolating the outer panels from the inside panels. It was

employed to mitigate large temperature variations on the internal components. Glass epoxy with a conductivity of

0.471 W/m·K was used as insulation10. In addition, entire outside surface of solar-cell-mounted panels were initially

covered with Kapton polyimide. Whereas, surface finish of external panels’ inner surfaces and internal panels were

kept as manufactured. Unlike paints which are permanent, final absorptivity and emissivity of each panel may be

adjusted by reducing the total area of Kapton polyimide. The final area of Kapton polyimide was estimated based on

the simulations using the different parameters discussed in this section.


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Figure 6. Sample external panels with Kapton polyimide surface finish.

IV. Thermal Analyses and Results

The thermal environment simulations were carried out using Thermal Desktop and an originally developed

analysis tool to analyze the thermal design of DIWATA-1. By applying these, development cost and time was

greatly reduced.

A. Thermal Desktop Mathematical Model

The thermal mathematical model (TMM), shown in Fig. 7, was created using Thermal Desktop. In the model,

multiple nodes were set on components with complex geometries and large volumes specifically the external panels.

In total, the TMM has 790 external nodes and 526 internal nodes.

Figure 7. Thermal mathematical model of DIWATA-1 in Thermal Desktop.


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The size, mass, and material properties applied in the thermal mathematical model were measured from the

SolidWorks model and data gathered from the Tohoku BUS team. Values of optical properties utilized are shown in

Table 4. These were based from past record in Tohoku University during the development of DIWATA’s

predecessors such as SPRITE-SAT, RISING-2, and RISESAT.

Table 4. Optical properties.

Material Absorptivity Emissivity Ratio (α/ε)

Black 0.67 0.89 0.753

Solar Cells 0.92 0.8 1.15

Aluminum 0.255 0.025 10.2

Kapton Polyimide 0.515 0.76 0.678

B. Originally Developed Analysis Tool

The originally developed analysis tool was used as a supplement to Thermal Desktop. It was created at Tohoku

University’s Space Exploration Laboratory and consists of three nodes, as shown in Figure 8. Two nodes were

allocated to the satellite and one node for space. The tool operated under the assumption that the satellite is

radiatively isolated. This means that the value of emissivities of both the internal panels and inside of external panels

is low and similar.

Figure 8. Two-nodal analysis using an originally developed tool.

The outputs are temperatures of both external and internal panels. Inputs necessary were orbital elements, optical

properties of external surfaces, and conductance between the surface panel and inside structure nodes. The value of

conductance was calculated based on the total area with glass epoxy. The final areas of each material in all panels

were estimated with the help of this tool. Figure 9 shows a sample calculation of the average absorptivity and

emissivity values for each outer panel. These values were eventually applied to the external panels of DIWATA-1.


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Figure 9. Absorptivity and emissivity sample calculation based on material area

C. Simulation Results

Two-nodal analyses were first performed to check the influence of the radiative isolation assumption and to

reduce the number of possible optical properties. Areas of each material were varied in the developed tool until the

range of about 0-40ºC was achieved. An example of results using this tool and the optical properties in Figure 9 is

shown in Table 5.

Table 5. Originally developed analysis tool results

Case Outer Panel Temperature (ºC) Central Panel/Battery Temperature (ºC)

Min Max Min Max

Cold, 5 W internal heat

dissipation -13.9 6.6 -1.9 -0.3

Hot, 25 W internal heat

dissipation 47.1 49.7 58.3 58.6

Optical properties in Figure 9 presented temperatures near the desired range but doesn’t satisfy the allowable

range. In addition, further adjustments resulted to either a much lower or higher temperature than the values in Table

5. This means that the assumption of radiative isolation was confirmed to be ineffective in terms of temperature

prediction for DIWATA-1 despite having an aluminum surface finish on the inside of external panel, and an

aluminum surface finish on the internal panels and on some component casings. However, these results provided a

benchmark on the expected temperature range in performing multi-nodal analyses. To confirm this, a model with

higher number of nodes was generated using Thermal Desktop. In theory, TMMs have larger effective radiative area

and higher effective optical property value due to the addition of components attached to the internal panels.

Analyses results from Thermal Desktop TMM, using the optical properties shown in Figure 9, for both worst hot

case and cold case are summarized in Figure 10. Both minimum and maximum temperatures of components in

different satellite orientations were recorded together with its allowable operating temperature range defined by a

grey region. As predicted, the temperatures of components were within the range set by the two-nodal analysis. But

unlike the two-nodal analysis, Thermal Desktop results showed that all inside components satisfied the operating

temperature ranges. The temperature of battery, one of the most critical components for a satellite’s mission success,

at worst hot condition was found to be around 38ºC and 14ºC at worst cold condition. That being the case, the

Kapton area of each panel were finalized and implemented on the flight model of DIWATA-1. Validation of these

results through comparison with the on-orbit temperatures will be further discussed on the next section.


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Figure 10. Thermal Desktop analyses results

V. Comparison of Simulation and Orbital Data

The thermal analysis results and orbital temperature data were compared to assess the thermal design of

DIWATA-1. Figure 11 shows beta angle history of DIWATA-1 starting from release until February 13, 2017. The

figure was generated from STK. In less than a year, the satellite has already experienced worst hot conditions three

times and cold conditions multiple times.

Figure 11. Beta angle history of DIWATA-1.


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The first worst hot condition occurred last May 26-29, 2016, approaching summer solstice. DIWATA-1’s +Y

Panel was in direct sight of the sun when attitude control was used. Beta angle calculated was 73.2º. As the theory

suggested in Equation 1, no-eclipse phase occurred and highest temperature were noted. The comparison between

the on-orbit temperature and simulation temperatures are plotted in Figure 12. Simulation temperatures were

obtained from Thermal Desktop using the same orbital parameters on-orbit. The battery temperature from simulation

and on-orbit were found to have almost the same value. Meanwhile, a ±3ºC difference was observed for other

payloads. The difference was estimated without applying any analytical uncertainties.

Figure 12. Simulation results and on-orbit data of components during WHC: May 2016

Major difference was noticed in the DPU chip and no simulation result was recorded for FPGA. Both parts are

housed inside the DPU component. This was mainly due to the discrepancy between the TMM and the flight model

as shown in Figure 13. Initial TMM was assumed to be one FD Solid. Meanwhile, in the flight model, the DPU

component consists of aluminum casing, FR-4 board, DPU chip, FPGA, and other electronical parts. To address the

difference and to have a better prediction on these parts for future extreme cases, the model was modified right after

the first worst hot case condition.

Figure 13. DPU chip and FPGA FM vs Initial TMM

The next worst hot case conditions occurred on July 26-28 and December 22-26, 2016. Worst cold conditions,

on the other hand, were experienced frequently based on Figure 11 but an instance of this condition transpired on

September 13, 2016. Figures 14 and 15 shows the comparison of simulation results and on-orbit temperatures for

July and September, respectively. A maximum difference of ±5ºC was observed on some components for both

cases.


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Figure 14. Simulation results and on-orbit data of components during WHC: July 2016

Figure 15. Simulation results and on-orbit data of components during WCC: September 2016

Figure 16. Simulation results and on-orbit data of components during WHC: December 2016


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The most recent hot case condition happened during last year’s winter solstice. This time around, the -Y Panel of

DIWATA-1 was in direct sight of the sun when attitude control was used. The highest beta angle value of 75º was

logged. Figure 16 shows the comparison between the on-orbit temperature and simulation temperatures.

Unfortunately, the battery temperature exceeded the operating temperature range. Image capture and download were

suspended during the period of no eclipse and the satellite operated in standby mode. The primary cause of the

exceedingly high temperature may be attributed to the high radiative heat transfer between the -Y panel and battery

which is located right behind the -Y panel, and other workmanship errors. A difference of ±7ºC was calculated

between the simulation results and on-orbit temperature.

VI. Conclusion

Passive thermal design and pre-flight thermal analyses of DIWATA-1 were carried out in this paper. DIWATA-

1, deployed from the International Space Station, followed a non-sun-synchronous orbit. The thermal environment

on this orbit was proven to be highly dependent on the beta angle and season. Orbital elements for worst condition

simulation were estimated from this parameter. DIWATA-1 employed a passive thermal design strategy due to

surface area and power limitations. Passive design was realized by placing conductive insulator between the internal

panels and external panels initially covered with Kapton polyimide. The final area of Kapton polyimide was

adjusted until temperature of each component were satisfied based on the results using Thermal Desktop and

originally developed analyses. The thermal design was evaluated through comparing simulation results with on-orbit

temperature data. A ±14ºC difference between hot and cold cases using the originally developed analysis tool.

Meanwhile, the various comparisons using Thermal Desktop showed a maximum ±7ºC difference, lower than

acceptable error of ±10ºC, even without performing time-consuming thermal tests. However, higher battery

temperature than simulation may have been noticed and further adjustments on the final Kapton areas made if some

tests were performed. Despite this, simulation using a numerical model with higher number of nodes and the

established worst case conditions was nonetheless sufficient. Moreover, the design procedure is simple and

promotes reduction in development costs and time. Finally, to further improve the prediction for DIWATA-1 and

the next satellites, thermal conductance on the model will be modified using future telemetry data.

Acknowledgments

DIWATA-1 is under the PHL-MICROSAT project funded by the Department of Science and Technology

(DOST) thru the Philippine Council for Industry, Energy and Emerging Technology Research and Development

(PCIEERD). The project is implemented by University of the Philippines–Diliman Electronics and Electrical

Engineering Institute (EEEI), Institute of Environmental Science and Meteorology (IESM), Training Center for

Applied Geodesy and Photogrammetry (TCAGP), and Advanced Science and Technology Institute (ASTI) in

partnership with Tohoku University and Hokkaido University.

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Thermal Design and On-orbit Validation of the First