FORMOSAT-3 Satellite Thermal Control Design and Analysis *

Journal of Aeronautics, Astronautics and Aviation, Series A, Vol.39, No.4, pp.287-292 (27) 287 Technical Note FORMOSAT-3 Satellite Thermal Control Design and Analysis * Ming-Shong Chang **, Chia-Ray Chen, Jeng-Der Huang, and Jih-Run Tsai Mechanical Engineering Division National Space Organization (NSPO) National Applied Research Laboratories (NARL), Taiwan, R.O.C. 8F, 9 Prosperity 1st Road, Hsinchu Science Park, Hsinchu 378, Taiwan, R.O.C. ABSTRACT FORMOSAT-3 program consisting of six small-satellites are used for weather prediction, atmospheric studies, and space weather monitoring. The satellites were injected into the parking orbit of 516 km and being transferred to final orbit of 8 km individually. Because electrical power is quite critical for micro-satellite, passive thermal control hardware is used as much as possible to save heater power and heaters are only designed and used to warm up critical components. All worst hot and cold cases are analyzed with corresponding beta angles, orbit altitudes, thermal environments, and operating modes. After conscientious trade-off studies among all worst hot and cold cases, radiator area and heater power are determined and finalized to meet all design requirements. Keywords: FORMOSAT-3, Satellite, Thermal control, Thermal design and analysis I. INTRODUCTION FORMOSAT-3 is the third major satellite program of NSPO (National Space Organization) and it s a joint cooperation between NSPO and USA University Corporation for Atmospheric Research (UCAR). The primary goal of FORMOSAT-3 mission is to launch a constellation of six micro-satellites into Low-Earth Orbit (LEO) to collect GPS radio occultation data for weather prediction, atmospheric studies, and space weather monitoring. Each satellite has GPS Occultation Experiment (GOX), Tiny Ionospheric Photometer (TIP), and Tri-Band Beacon (TBB) payloads. For satellite thermal design criteria is the temperature of the components and payloads shall be properly controlled within their allowable limits through the thermal design when the satellite exposes to the outer space environment on orbit mission operation. Because of the limited electrical power generated from solar panels for micro-satellite, the thermal control is achieved mainly through passive elements, such as MLI(Multi-layer insulation) thermal blankets, thermal fillers, and surface finishes. However if the thermal requirements cannot be met by the passive control, the active element, i.e., heater, may be used to maintain the unit temperatures above their lower limits. II. SATILLITE CONFIGURATION The dimension of each micro-satellite is 116 cm in diameter with thickness of 16 cm as shown in Figure 1. Each satellite weight is estimated about 62 kg (with propellant) and the average power consumption is about 81W. The +Y and Y solar panels with 31 o cant angle are rotated with the Y-axis to track the sun to the maximum extent in the normal operating mode. The TBB antenna boom will be deployed in the +Z direction after the satellite separation from the launcher. Figure 2 shows the satellite orientation in the orbit with different beta angles. Beta angle is the angle between orbit plan and sun vector. When the beta angle is 9 o, i.e., the sun vector from the top of orbit plane, Y side solar panel is shaded by the satellite and receives no solar energy. When the beta angle is o, the sun vector * Manuscript received, Mar. 28, 27, final revision, Jul. 6, 27 ** To whom correspondence should be addressed, E-mail: mschang@nspo.org.tw

288 Ming-Shong Chang Chia-Ray Chen Jeng-Der Huang Jih-Run Tsai is parallel to the orbit plane. When the beta angle is less then o, the satellite performs flip-flop and will turn 18 o along the Z-axis (pirouette maneuver) in order to get the maximum solar energy for power generation. other to maintain power-safe and thermal-safe conditions with the non-essential loads off. FORMOSAT-3 satellite thermal control design for external surfaces are achieved by using MLI and radiators (Silver-Teflon, Second Surface Mirror) as shown in Figure 3. All spacecraft outside surfaces are applied with MLI except that the Y side bus panel is applied with radiator to radiate the internal waste heat to the deep space. Besides, another small radiator is located on the outside of Z side bus ring to keep the solar array drive temperature below its hot limit. External geometric model Solar Array MLI Figure 1 Satellite configuration Silverized Teflon Radiator Total Area=.24 m 2 32% of Available Area X Z Y Magnetometer External TRASYS 343 Surface Nodes and SINDA 538 Nodes Z Y X Figure 3 Satellite external thermal design Figure 2 Satellite orientation in normal mode III. THERMAL DESIGN APPROACH The main purpose of thermal control design is to maintain all of the satellite components and payloads temperature within their allowable temperature limits when the satellite exposes to the outer space environment[1] on orbit mission operation. The satellite will suffer from varying external heating of complicated space thermal environment and internal power dissipation from all units in different operational modes during the 5-year design life after launch. The beta angle varies from 9 o to +9 o and the altitude may change from 516 km to 8 km. The operational modes of satellite are defined as follows: - Yaw-steering mode: the mode to gain the maximum sun pointing for solar panels. - Normal mode: the mode to support all science operations and normal operations for data gathering. - Safing mode: Two sub-modes are designed and include one to maintain attitude-safe condition and the All satellite internal components are treated with surface finishes (black paint, Primer, and irridite) as shown in figure 4 to enhance the radiation coupling and mounted on the deck with thermal fillers (Nusil and Silpad) to increase the contact conductance. Battery and the propellant tank and lines are covered by MLI to reduce the radiation interchange with other components. The software-controlled heaters with thermistors are applied on battery, propellant tank and lines, REA valve and payload GOX to maintain the minimum required temperatures when satellite is in the cold conditions. Internal geometric model (with battery MLI) TX Filter RX Filter Diplexer Receiver Transmitter TBB Hybrid Coupler Battery MLI BCR TIP TIP Sensor Stiffener Internal TRASYS 316 Surface Nodes Solar Array Drive Tank Stiffener Torque Rod RWA Avionics GOX MIU/PCM Y Z X Surface Finish Black MLI Irridite Primer Radiator α/ ε.85/.8 5.6/.72.4/.14.5/.8.2/.76 Thermo-optical properties Figure 4 Satellite internal thermal design In order to take into account the errors for modeling assumptions and uncertainties from the thermal physical properties, a thermal uncertainty margin of 8 C before model correlation with thermal balance test (5 C after

FORMOSAT-3 Satellite Thermal Control Design and Analysis 289 model correlation) are included in all analyses so that the maximum or minimum predicted flight temperatures can be determined. If the component has the heritage in the previous programs, the requirement of 8 C uncertainty margin could be waived. Each heater is provided with a ground-command override capability and sized to control the component to 5 C above its minimum allowable temperature or to control the component at its minimum allowable temperature with 25% excess control authority. (W/ o C-m 2 ) is used in the battery-pack interface. Five battery cells are mounted on the battery spreader plate with Silpad and then are mounted to the deck with Nusil as shown in Figure 5. IV. THERMAL PARAMETER Thermal parameters include the solar constant (direct solar flux), earth infrared emission, albedo (reflected solar flux from the earth), and the satellite surfaces properties such as solar absorptivity and infrared emissivity in both (Begin-of-Life) and EOL (End-of-Life). All these parameters are used to calculate the environmental radiation and are listed in Table 1 and Table 2. Silpad Nusil Table 1 Thermal environmental constants Hot Solar constant 1287. (W/m2) 142. (W/m2) Albedo Coefficient.25.4 Earth Emission 19. (W/m2) 262. (W/m2) Beta Angle -9 o to +9 o Parking Orbit: 516km; Altitude Mission Orbit: 8km * Albedo= Albedo coefficient * incident solar flux on earth Table 2 Thermal optical properties Solar Absorptivity (α) EOL Infrared Emissivity (ε) Silver Teflon (Radiator).8.15.76 MLI.4.6.72 Solar cells.721.775.79 Black Paint.85.85.9 Thermal capacitance of each component used for the temperature predictions is assumed to be 92 (J/kg o C). All components are mounted on the deck with two kinds of thermal fillers to reduce the contact resistance in the mounting interfaces. One thermal filler called Nusil with the contact conductance of 31 (W/ o C-m 2 ) is used for all components except batteries. The other one called Silpad with the contact conductance of 775 Figure 5 The installation of batteries and ancillary hardware The effective emissivity ε* of MLI (Multi-layer Insulation) is typically ranged between.1 and.5 which depends on the layers, material, manufacture, and assembly workmanship. There are 13 layers of Kapton films with double-side aluminum coatings used for MLI manufacture. Their thermal insulation performance should be good enough and the ε* value should be low. The effective emissivity ε* is assumed to be.1 and.5, respectively, for the worst cold and hot cases analysis where worst cases studies are adopted to predict the extreme hot and cold temperatures of satellite component units. V. THERMAL MODELING The typical approach for satellite thermal analysis is done by thermal network lump system. The impact of thermal environments on satellite is calculated by radiation analysis tool, i.e., TRASYS [2]. The satellite orbit parameters, satellite geometry and thermal optical properties are the inputs of analysis tool and executed by TRASYS code. The output of TRASYS code is satellite related radiation parameters including absorbed heat fluxes on each external surface from direct solar flux, earth emission, albedo, as well as equivalent radiation conductors between any two radiation surface nodes. The TRASYS output results can be transformed and directly inserted into the thermal network analysis tool, i.e., SINDA/G [3], for the calculations. In the SINDA/G modeling, the geometries of satellite are converted into a group of thermal nodes with corresponding node capacitances, and these nodes are

29 Ming-Shong Chang Chia-Ray Chen Jeng-Der Huang Jih-Run Tsai connected into a thermal network with the thermal conductance between these nodes [4],[5]. There are totally 376 external and 342 internal TRASYS surface nodes and 546 SINDA nodes in the current thermal model. VI. THERMAL ANALYSIS RESULTS We do thermal design about satellite hot cases thermal analysis for radiator sizing and solar array temperature prediction. Then from the hot case radiator sizing results, we do heater check for cold cases thermal analysis to get the required heater power and we also consider the solar array cold case temperature prediction. The detail method and theoretical thermal design analysis model is following the thermal network lump system[1]. Due to the satellite operation scenario, lots of thermal analysis cases need to be performed for different beta angles, orbit altitudes, operating modes, orbit thermal environments and power dissipations to make sure all components operating within their temperature limit on satellite orbit mission operation. The satellite beta angle is ranged from -9 o to +9 o and the orbit altitude is ranged from 516 km to 8 km. Due to different flight operation requirements, there are three operating modes including normal mode, yaw-steering mode, and safing mode. The satellite power dissipation will be different under various operating modes. All analysis cases and power dissipation tables are summarized in Table 3 and Table 4, respectively. Battery is not a unit with constant heat dissipation and its value depends on the conditions of charge, trickle charge, and discharge. Figure 6 shows the power profile of batteries for one orbit at o beta angle 8 km altitude. When beta angle is 9 o, there is no eclipse and the five batteries power dissipation is 3.3 W under the constant trickle charge. From the preliminary thermal analysis results, the worst hot and cold conditions for satellite bus and solar panels are identified and described as follows: Table 3 Thermal analyzed cases for FORMOSAT-3 spacecraft Normal Operating Mode Yaw-Steering Mode Safehold Mode Damper Before Deployment Altitude (km) 516 8 516 8 8 516 Beta Angle (deg), 3,6, 9, 3,6, 9 6, 9 6, 9 Thermo-optical Thermal Properties Environment EOL EOL Hot Hot Thermal Design Radiator sizing Heater power check Radiator sizing Heater power check Heater power check Damper Temperature Prediction Before Deployment Table 4 Power dissipation Average Power (W) Components / Payloads Normal Orbit β= o β=9 o Safing Orbit Avionics 4.45 5.56 6.2 MIU/PCM 9.2 9.2 2.1 Reaction Wheel.68.68. Earth Sensor.3.3.33 Torque Rod.18.18 1.4 BCR 8.9 8.9 4.32 Magnetometer..11.12 Solar Array driver 1.12 1.72.72 Receiver 8.67 8.98 5.84 Transmitter 32(on 32(on 15min) 15min) 32(on 15min) GOX (GPS/SSR/PC) 19. 19.. Tri-Band Beacon (TBB) 7.8 8.59. TIP Sensor Assembly (TSA) 4.3 4.4. TIP Interface Control Elect. (TICE) 2.2 2.2. Power(W) 6 5 4 3 2 1-1 -2-3 2 4 6 8 1 Time(min) Figure 6 Battery power dissipation profile for one orbit at o beta angle 8km altitude(one orbit is 12 minutes) - The worst hot case for satellite bus: Normal mode, beta angle at 9 o, hot environment, EOL thermal properties, and the altitude of 73 km. - The worst hot case for solar array: Normal mode, beta angle at 3 o, hot environment, EOL thermal properties, and the altitude of 516 km. - The worst cold case for satellite bus and solar array: Safing mode, beta angle at o, cold environment, thermal properties, and the altitude of 8 km. In order not to exceed the upper temperature limits of satellite components, radiator areas of satellite are carefully sized and tailored under the worst hot conditions. According to the analysis results, the required radiator areas in Z and Y sides are about.7 m 2 and.24 m 2, respectively. All the temperatures are controlled within their upper temperature limits with 8 C

FORMOSAT-3 Satellite Thermal Control Design and Analysis 291 uncertainty margin except the components with heritage such as Reaction Wheel Assembly (RWA), earth sensors, and solar array. The most critical components for hot case are earth sensor and solar array with 1 C and C uncertainty margin, respectively. In order to meet the thermal control requirement that is to maintain all the bus and payload components above their lower temperature limits, heaters are applied. The heater power requirement and locations are calculated under the worst cold condition. The required heater power, heater duty cycle, heater resistance and number of heater for heater control components are listed in Table 5. All the temperatures are controlled within their cold limits with proper margins. Battery, tank, propulsion line, REA valve and GOX are thermally protected with heaters, so their temperatures are well controlled within the range of corresponding heater thresholds. The most critical components for cold case are battery, tank, propulsion line, REA valve and GOX that are thermally protected with heaters. worst cold case is similar to the one in the worst hot case but scale is different. The minimum temperature is about -95 C that is 5 C higher than the minimum operating temperature limit, -1 C. Temperature( o C) 125 1 75 5 25-25 -5 Solar Array Maximum Temperature for Normal Hot, Operating 4.2% Efficiency, H=516 km, Beta=3 o Hot Temperature Limit -75 Table 5 Required heater power, heater duty cycle, heater resistance and number of heater for safing cold begin of life beta= o at 8Km altitude -1-125 1 2 3 4 5 6 7 8 9 Time(min.) Eclipse One Orbit Figure 7 Temperature distribution of solar array for the worst hot case 125 Solar Array Minimum Temperature for Normal, Operating 11.3% Efficiency, H=8 km, Beta= o 1 75 5 Solar array panels are directly exposed to the space environment and the temperature variations are much significant than the others. Figure 7 shows the transient temperature distribution of solar array under the worst hot condition with 4.2% solar array efficiency (i.e., 4.2% of the incident solar heat flux transformed to the satellite electrical power and the rest becomes heat). Because of the 31 solar array cant angle, the maximum hot temperature occurs at around 3 beta angle in the analysis cases. The temperature curve of the solar array decreases in eclipse and then increases and reach the maximum in the daytime. The maximum solar array temperature is around 15 C that is equal to its maximum operating temperature limit, 15 C. Figure 8 shows the transient temperature distribution of solar array under the worst cold condition with 11.3% solar array efficiency, i.e. more electrical power and less heat dissipation. The overall orbit thermal behavior of solar array panel in the Temperature( o C) 25-25 -5-75 -1 Temperature Limit -125 One Orbit 1 2 3 4 5 6 7 8 9 1 Time(min.) Eclipse Figure 8 Temperature distribution of solar array for the worst cold case VII. CONCLUSIONS FORMOSAT-3 satellite will be under large beta angle range from to 9 which will cause large variation thermal environment. With much effort for

292 Ming-Shong Chang Chia-Ray Chen Jeng-Der Huang Jih-Run Tsai proper thermal design and analysis, FORMOSAT-3 satellites are controlled and maintained in the allowable temperature ranges during the entire mission life. The thermal design by selecting typical thermal hardware such as MLI, radiators, thermal fillers, and heaters is analyzed and verified in order to meet the mission thermal requirements. The required radiator area is about.31m 2. The required heater power is about 14.6W. All components are well controlled within allowable temperature ranges. ABBREVIATION AND NOMENCLATURE BCR: Battery Control Regulator : Begin of Life CAT beds: Catalyst beds EOL: End of Life GOX: Global Position System Occultation Experiment MIU: Mission Interface Unit MLI: Multi-layer Insulation PCM: Power Converter Module RCS: Reaction Control Subsystem REA valve: Reaction valve RWA: Reaction Wheel SSR/PC: Solid State Recorder/Payload Computer TBB: Tri-Band Beacon TICE: Tiny Ionospheric Photometer Interface Control Electronics TR: Receiver TSA: Tiny Ionospheric Photometer Sensor Assembly TX: Transmitter REFERENCES [1] Gilmore, D. G., Satellite Thermal Control Handbook, the Aerospace Corp., 1994. [2] Thermal Radiation Analyzer System (TRASYS) User s Manual, by Lockheed, LEMSCO-22641 and JSC-22964, April 1988. [3] SINDA/G User s Guide, Network Analysis Associates, INC, 1994. [4] Kang, C. S., Huang, J. D., Tsai, J. R., and Ting, L. H., Satellite Thermal Modeling Techniques and Computation Tools, The 4 th Conference on Aeronautics and Astronautics, R.O.C., Taichung, Taiwan, pp. 253-26, December 1998. [5] Tsai, J. R., Thermal Analytical Formulations in Various Satellite Development Stages, 8 th AIAA/ASME Joint Thermophysics and Heat Transfer Conference, ST. Louis, Missouri, USA, June 22.