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Thermal Control Design of TUSAT

Murat Bulut1, Nedim Sozbir2 and Senol Gulgonul3

Turksat International Satellite and Cable Operator, Golbası, Ankara, 06839, Turkey

TUSAT is Turkish Communication Satellite, providing mainly communication and TV services using C-band and Ku-band channels. This paper describes the thermal control design which uses passive and active concepts. The active thermal control is based on heaters regulated by software via thermistors. The passive thermal control composes of heat pipes, multi-layer insulation (MLI) blankets and radiators, paints, surface finishes to maintain temperature level of the overall carrier components within an acceptable value. The thermal control design is supported by thermal analysis using thermal mathematical models (TMM).

Nomenclature BOL = beginning of life CCHP = constant conductance heat pipe CDR = critical design review CM = communication module CSS = coarse sun sensor EOL = end of life EPC = electrical power conditioner EQ = equinox GEO = geostationary MLI = multi-layer insulation N/S = north/south OMUX = output multiplexer OSR = optical solar reflector PCU = power conditioning unit PDR = preliminary design review PFLDIU = payload platform distribution unit RW = reaction wheel S/C = spacecraft SM = service module SMU = satellite management unit SS = summer solstice STR = star tracker TCR = telemetry, command, ranging TCS = thermal control subsystem TMM = thermal mathematical model T/O = transfer orbit TWT = traveling wave tube UPS = unified propulsion system WS = winter solstice

1 Satellite Thermal Specialist, Satellite Design and R&D Center, Konya yolu 40.Km Golbasi, Ankara. 2 Consultant, Assistant Professor, Satellite Design and R&D Center, Konya yolu 40.Km Golbasi, Ankara. 3 Satellite Design and R&D Center Director, Satellite Design and R&D Center, Konya yolu 40.Km Golbasi, Ankara.

6th International Energy Conversion Engineering Conference (IECEC)28 - 30 July 2008, Cleveland, Ohio

AIAA 2008-5751

Copyright © 2008 by Nedim Sozbir. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.

I. Introduction USAT Thermal Control Subsystem (TCS) consists of active and passive thermal control elements to maintain the spacecraft components and structure within a controlled range of temperature during all the mission of the

spacecraft from Beginning of Life (BOL) to End of Life (EOL). The payload equipments and their operational requirements were a main drive to develop TUSAT thermal control design, and analysis.

T TUSAT is a geostationary (GEO) communication satellite. It has a maneuver lifetime of at least 16 years and has an operational lifetime at its nominal location of at least 15 years. TUSAT uses a three-axis stabilized type satellite platform and it consists of communication module (CM) and service module (SM). TUSAT payload configuration provides maximum 16 active in Ku-band channels and maximum 4 active C-band channels. The total number of redundant transponders is 4 Ku-band and 1 C-band. There are three distinct coverage areas named as East, West and Turkey coverages. East and West coverages are in Ku-band and both transmit and receive. Turkey coverage is in both Ku and C-bands. Ku-band Turkey coverage is transmit only. C-band Turkey coverage is both receive and transmit.

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As seen from Figure 1, the antenna subsystem features two main shaped antenna reflectors deployed on the east and west face of the spacecraft and two Gregorian antenna located at earth deck. Main features of the satellite are:

• Launch mass with maximum dry mass: 2553.5kg

• Satellite nominal dry mass: 1140 kg

• Overall dimensions of the main body: 2200 x 2000 x 2825 mm

Figure 1. TUSAT in-orbit configuration

• Solar array wingspan: 14564.5 mm • Payload Consumption (EOL): 3676 W • Satellite Power Consumption (EQ-EOL): 4828 W • Satellite Solar Array Power (EOL): 5606 W

II. Program Status TUSAT is the first Turkish Communication Satellite which is designed by Turkish engineers in cooperation with

Thales Alenia Space in France. Presently TUSAT is at PDR level; it is expected to be ready for critical design review (CDR) and completely thermally equipped and launched in 2015.

III. Spacecraft Subsystem Thermal Design TUSAT satellite is functionally divided into two main subsystems which are payload and platform. The payload

ensures the communication mission. It includes the repeater subsystem, the antenna subsystem, and the telemetry, command, ranging subsystem (TCR). The platform ensures the mission control and stability. It consists of the avionics, the unified propulsion subsystem (UPS), the solar array subsystem, the structure subsystem, the thermal control subsystem, the harness, and mechanism subsystem.

The dissipative payload equipments are located on North and South panels and are installed on the main constant conductance heat pipe (CCHP) networks using thermal fillers to improve the thermal contact between units and heat-pipes. CCHP networks are subdivided in separate networks according to the different qualification temperature levels. Payload and platform equipments with CCHP layout is shown on south CM and SM panels in Figure 2. External surfaces of North and South panels are covered by optical solar reflector (OSR). The radiative areas of these panels are sized to radiate the maximum dissipation corresponding to 16 Ku-band and 4 C-band operating channels in full operation. Heaters are implemented to compensate the variation of the payload dissipation versus the repeater operational modes. The structure subsystem keeps inside payload and platform equipment and mates with the launch vehicle. The subsystem is designed to withstand the natural environmental forces for all static and dynamic loads encountered

during ground handling, transportation, ground test, and launch phases. In order to minimize the temperature gradients within S/C, internal surfaces have high emittance finishes to aid radiative heat transfer.

The propulsion subsystem is bi-propellant. Multi-layer insulation (MLI) or low emittance coating is used around propellant tanks and lines. To insure minimal required temperatures, dedicated heaters are used on propellant tanks, UPS lines, 10 N thrusters, and 400 N engine. To protect the satellite from heat flux and plume impingement during firing, high temperature protections and heat shields around thrusters are used. The SM units are located on lower part of north and south panels. Batteries, PCU, and PFLDIU are located on these panels. Li-Ion batteries are used. They are mounted on N/S SM panels. The batteries are radiatively discoupled from internal S/C through MLI blankets. Dedicated radiative areas are designed to reject the heat dissipation. Battery design concept includes thermal filler at baseplate mounting interface. Nominal and redundant heaters are used to maintain batteries at minimum temperatures.

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The solar array subsystem generates the necessary power for full operation of the satellite during sunlight and battery charging for eclipse. Two solar array wings are fixed on the North and South faces of the body. Two types of optical sensors are used for attitude determination and control. These sensors are the star tracker (STR) and coarse sun sensor (CSS). The star trackers (STR) and sun sensor are fixed on SM, upper earth panel, and anti-earth panel respectively. Dedicated thermal design is required for these optical sensors. Dedicated radiative areas and heaters are used. Due to particular exposure to space and sun, MLI blankets are used to insulate the unit except the optical active sensor and radiative areas. The communication mission is achieved using for antennas. Two of them are gregorian antennas and two are deployable antennas. Gregorian antennas are mounted on the earth deck. Deployable antennas are hinged on the east and west panels. In addition, two sets omni antennas are used the Telemetry, Command and Ranging mission. Antennas are thermally insulated from S/C structure. The deployment mechanism I/F with the S/C are insulated by thermal washers. Dedicated control heaters are installed.

Figure 2. South panel layout

IV. Thermal Requirements and Constraints The thermal control is required to ensure acceptable temperature limits during the whole lifetime and for the

various operating modes. The main requirements and constraints are the phase of the mission, temperature requirements for the units, and dissipation of the units. The thermal control is designed during launch, transfer and operational phases. During launch phase solar array and the deployable antenna are folded. In transfer orbit (T/O), the satellite solar array is partial deployment configuration and deployment of reflector antenna. In synchronous orbit the satellite is three axis stabilized with full deployment of solar array and reflector antennas.

In order to ensure the mission during the whole life time of the satellite, thermal control are guaranteed for all units with regard to qualification temperature limits as described in Table 1. Dissipation of the units is one of design constraints when the thermal control is considered. Thermal dissipation of the satellite on CM panel is given in Table 3.

V. Passive and Active Thermal Control Hardware

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The aim of the thermal design of a satellite

is to keep all its components within their specified temperature range. The common design approach is to use a combination of MLI blankets, OSR, CCHPs, heaters, surfaces finishes, paints, and thermistors.

TUSAT thermal control uses the North and South panels to reject the internal heat dissipation and to limit the diurnal variation because of the minimal solar illumination on these faces. The radiative areas are covered with optical solar reflector (OSR). The OSR is considered to have the most stable optical thermal properties in space and is used in situations where a surface finish with a low value of solar absorbtivity α and a high IR emmisivity ε is needed. Typical thermo-optical characteristics are α (BOL) = 0.11, α (EOL) = 0.27 and ε=0.84. Sizing of the radiative areas is performed taking account the worst conditions that are maximum heat dissipation, maximum solar illumination (solstices), end-of-life thermo-optical properties. TUSAT radiative areas on each SM panel and CM panels are summarized in Table 2 and Table 3 respectively. The satellite body is insulated from the space environment by wrapping it with MLI blankets. MLI is externally used on the satellite everywhere except north/south radiative areas as well as on surfaces requested to be free in order to minimize the heat input from solar radiation or the heat leakage. MLI is internally used either to avoid excessive heating of

Table 3. OSR radiative areas.

Panel Description Max

Qualification Dissipation OSR

Requirement OSR

Available

oC W m2 m2 CM EPC Ku and C bottom 65 154.4 0.59 0.77 North TWT Ku and C , OMUX 85 803.9 2.53 2.76

Receiver, Beacon, Output Filter, TM 65 45 0.17 0.26

Total Panel 1003.3 3.29 3.79 CM EPC Ku and C bottom 65 154.4 0.61 0.71 South TWT Ku and C , OMUX 85 805.6 2.62 2.65 TCR 65 82.7 0.33 0.35

Total Panel 1042.7 3.56 3.71 Total 2046 6.85 7.5

Table 1. Equipments qualification temperatures.

Qualification Temperature requirement (oC) Equipments designation

Min Max Receiver -5 65 IMUX -5 65 Docon -5 65 CAMP -15 65 Single EPC -15 65 TWT -15 85

Repeater

OMUX 10 80 TC Receiver -30 70 TCR TM Beacon -30 65 RW -20 70 PCU -30 70 PFLDIU -30 65 SMU -30 65

SM Units

Li-ion Battery 0 40

Table 2. SM radiative areas.

Panel DescriptionOSR

Requirement OSR

Available

m2 m2 SM PCU 0.81 1.02 North Battery 0.11 SM PFLDIU 0.81 1.01 South Battery 0.11

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components located on the internal structure or around batteries and propulsion components to emphasize heating efficiency by reducing heat leakage. MLI is also internally used to isolate high temperature radiative areas from internal units.

Natural surface finishes are used under the insulation blankets for all graphite epoxy structures. Dissipative units are black painted. Black paint is used on the inner faces of the north and south panels. Thermal coating is not applied to equipment units with low or zero dissipation if these equipments are provided with a low emissivity.

The high dissipative units of the communication and the service module are located on CCHP assembled in specific networks dedicated to radiator panels. CCHP networks are installed on the inner face of the north/south panels. The main heat pipes, directly in contact with the high dissipative units, are connected by crossing ones to spread the heat dissipation over the whole heat pipe networks. CCHP are used to spread the high concentrated heat dissipation from repeater units as TWT, OMUX, EPC, PCU and PFLDIU. Dissipative units are mounted on CCHP using thermal filler to ensure a good thermal transfer between units and CCHP. CCHP networks are subdivided into separate networks according to the different required temperature levels. CCHP layout and number have been designed in order to allow the failure of any heat pipe without consequent degradation of the satellite performance. Variations in environment and component heat-generation rates, along with the degradation of surface finishes over time, can drive temperature variations in a passive design to range larger than some components can withstand. Heaters therefore are sometimes required in a thermal design- to protect components under cold-case environmental conditions or to make up for heat that is not dissipated when an electronic box is turned off1. TUSAT uses heaters and thermistors as the part of active thermal. TUSAT thermal control uses heaters to provide temperatures control during nominal operation of several platform equipments such as optical sensors, thrusters and batteries and to control the payload equipments by compensating its power dissipation variation according to the operating modes and the effects of seasonal sun exposure. Thermistors are used instead of mechanical thermostat because it is very easy to modify the threshold temperatures from the ground. The active thermal regulation is ensured by software control automatically ON/OFF heater switching according to the temperatures provided thermistors by comparing the read value to the predefined temperature limits2. The heater regulation and the thermal control monitoring of units are done through thermistors.

VI. Verification of Design by Thermal Analysis The aims of the thermal analysis are to size the radiative areas and heating budget at PDR phase. In order to

maintain unit temperature (hot and cold cases) within non- operating and operating with specified temperature ranges, the heating budget needs to be determined.

TUSAT thermal analysis is concerned with predicting the temperature of a satellite in a known or assumed heating environment. The main aims of the Thermal Control Subsystem (TCS) at Preliminary Design Review (PDR) level is to validate that sufficient radiative surfaces are to keep equipments temperatures below their hot operational limit and to confirm that the available heating power budget is sufficient to maintain equipments within their operational temperature limits with regard to the payload drive level. Solar fluxes, satellite lifetime and satellite operational configuration are taken into account in analyzing S/C under the worst case scenario. Solar fluxes are taken into account during winter solstice (WS), summer solstice (SS) and equinox (EQ). Satellite operational lifetime is 15 years. Satellite operational configuration is based on number and location of channels ON/OFF, repeater operation level of drive from no drive to full drive.

The satellite thermal design was based on the analysis of the critical cases that exposed the equipment to extreme thermal conditions. For the external environment point of view, two critical cases were identified. These critical cases are hot case and cold case.

Hot case and cold case analyses are adopted to define upper and lower bounds on predicted temperatures. Hot case and cold case analyses are done for CM, SM, and external equipments in different condition. The power profile for a hot case analysis corresponds to an operation in which components’ activity results in high dissipation, while the orbit is such that the radiators are exposed to considerable solar fluxes3. The input data for the cold case are selected to result in a calculated lowest temperature.

For design purposes, all temperature predictions also include a 5 oC uncertainty. During normal operation, there is at least 15 oC margin between the qualification temperature limits and predicted temperatures at both the hot and cold end4.

Overall S/C Thermal Mathematical Model (TMM) using ThermXL was established since the beginning of the program and continuously updated if needed. The TMM was built up to a frozen version used for PDR level. Each modification of the satellite thermal model introduced in the TMM was justified, monitored and recorded in order to

the TMM evolution under control. A complete set of analysis cases covering main payload operational combinations and all seasonal and mission conditions was performed to demonstrate overall thermal compliance.

Given the thermal control design applied in each subsystem and system, four thermal analyses were carried out to the payload operational scenario. These scenarios were not the most possible ones and conceived only for assuring the thermal margin. The four cases are as follows4:

1. Equinox at BOL: This case represents the extreme cold environments of the components. 2. Summer solstice at EOL: This case represents the extreme hot environments for the components on the

north panel. 3. Winter solstice at EOL: This case represents the extreme hot environments for the components on the

south panel. 4. Equinox at BOL: This case represents the worst total power requirement at equinox.

On-orbit temperatures for each component and heater power consumption for the spacecraft are predicted. Main payload equipment temperature predictions resulting from PDR analysis are shown in Table 4 in GEO. The predicted heater power consumption for payload during GEO at EOL is 100 W.

VII. Conclusion for Thermal Design and Analysis

The TUSAT thermal control design and analysis were presented. This design has been verified through the use of ThermXL. The thermal design concept in system and subsystem to maintain the spacecraft payload components within a controlled range of temperatures was presented. The thermal analysis based on the normal and worst-case payload operational scenarios were provided. All the payload component temperatures can be maintained within operating limits for GEO phase. Most of the predicted temperatures, as shown in Table 4, were not centered in the allowed range. The temperature variation was within the required range in all. The predicted heater power consumption for payload meets TUSAT power budget.

Table 4. Payload equipment temperature predictions in GEO phase.

Module UNITS Qualification Extreme Calculated Temperatures Temperatures

Tmin(oC) Tmax(oC) Tmin(oC) Tmax(oC) EPC -15 65 -7.16 39.24 NORTH CAMP -15 65 -6.87 39.2

CM TWT -15 85 11.83 46.40 OMUX 10 80 13.54 45.00 TCR -30 65 3.05 33.79 EPC -15 65 -7.01 37.83 SOUTH CAMP -15 65 -6.49 37.78

CM TWT -15 85 11.72 48.49 OMUX 10 80 12.46 47.31 Rx -30 65 -0.18 35.02

The optimum radiator areas were obtained using the simplified model. The next step is to optimize the radiator areas and implement them in the detailed model to maintain the temperatures throughout the mission life. Extensive detailed thermal analysis of the spacecraft and all of its components under worst-case hot and cold conditions will be performed.

References 1Gilmore, D. G., Spacecraft Thermal Control Handbook Volume I: Fundamental Technologies, 2nd ed., The Aerospace Press,

California, 2002. 2Koedinger, M., and Brissonnaud, T., “SPACEBUS 3000- Thermal Control ARABSAT 2,” Proceedings of the Sixth

European Symposium on Space Environmental Control Systems, European Space Agency , Nordwijk, the Netherlands, 1997, pp. 57-66.

3Karam, R. D., Satellite Thermal Control for Systems Engineer., Vol.181, AIAA, Reston, VA, 1998. 4Hwangbo, H., and Kim, W. C., “Design of Thermal Control Subsystem for Koreasat 3 Communications Satellite,” 33rd

Thermophysics Conference, AIAA 99-3555, Norfolk, VA, 1999.

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