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A numerical solution is attempted for the steady state temperature distribution in a two dimensional rectangular radiating fin with single heated footprint. The solution is presented for problems similar to spacecraft doubler plate with single heated footprint. The spacecraft thermal doublers are used to spread the localized heat dissipation by some electronic packages over large area and thus ensure the operating temperature of the package within the specified maximum temperature. The results of this analysis are presented as the design charts for the design of doubler plates in a spacecraft environment. Basically, design of doubler plate is meant to find out the required radiator area and the thickness of the plate for given heat dissipation and the maximum operating temperature of the electronic package.

INSAT-2A spacecraft carries on-board a Very High Resolution Radiometer (VHRR) as one of the major payloads for the purpose of meteorological observations. The thermal infrared (IR) detector of the Radiometer requires cooling to cryogenic temperature for its operation at any of the set points in the range of 105 K to 115 K for optimum performance. A passive radiant cooler has been designed and realised indigenously to meet this requirement. A computer oriented mathematical model has been developed which simulates the on-orbit thermal environment and predicts the performance of the cooler for various sun illumination conditions. INSAT-2A spacecraft has completed more than one year of its successful operation on-orbit now. It has been observed that a minimum patch control power margin of 4.7 mW exists for beginning of life (BOL) condition i.e. 5th August, 1992.

Thermal control materials of different types are being used on the spacecraft for temperature control viz. Thermal control tapes, Thermal control paints, Anodised surfaces and Optical solar reflectors. This paper covers the measurement of emittance at different temperatures on various Optical Solar Reflectors ( OSR ) both rigid and flexible types. The rigid OSRs evaluated are OCLI - USA, PPE - UK and ISRO - INDIA. The flexible OSR is from SHELDAHL - USA. The emittance of the OSRs were measured in the temperature range 223K - 373K using steady state calorimetric method. Measurements were done on a square sample with OSRs affixed on each face. The sample has a built in heater and an evacuated low temperature radiation receiver is used for the test environment. Experimental results, temperature dependence of the total hemispherical emittance relation for each of these OSRs and error estimation are presented in this paper.

Prior to successful launch of IRS-1A, two different thermal balance tests were carried out, one on the Structure Thermal Model using heater pad technique and second on Engineering Model, using solar simulation technique. The temperature match was found to be better for extreme cold case in STM test and for EOL case in solar simulation test and not so good for other cases. An attempt was made to find out the reasons for this mismatch, but the exact reason could not be ascertained. The flight temperatures were predicted, based on solar simulation test data, which matches well with on orbit data.

The first Indian Remote Sensing Satellite, IRS-1A, has a mono-propellant Reaction Control System(RCS) for orbit and attitude correction. In early phase of spacecraft life some anomaly in fuel mass calculations was observed and one group of tanks were emptied much earlier than expected. This paper explains how heating/cooling rates of propellant tanks and comparison with theoretically calculated temperature response were used for sorting out the anomaly.