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This paper discusses about the software ‘ISCP’ developed for thermodynamic design and analysis of Stirling cycle cryocooler. The model is based on second order analysis method which assumes the various parasitic losses as independent of each other. The specifications of the cooler are arrived at using the Design Module of the software after performing various optimizations. The analysis module performs both dynamic and thermodynamic analyses.

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.

“Tinned Copper Braid” is planned to be used in IRS-1C spacecraft for the thermal control of the detector of the “Middle Infra Red” (MIR) band of LISS III Camera. The thermal conductance of the braid was evaluated experimentally in the temperature range of -65°C to +35°C by employing the calorimetric technique. The measured thermal conductance of the 215mm long copper braid varies from 0.051 to 0.056 W/°C at different temperatures in the specified range. An error estimation has been performed which shows a maximum uncertainty of better than ± 5% on the conductance value measured.

The measurement of coefficient of thermal expansion of CFRP strut which is planned to be used in Indian Remote Sensing Spacecraft (IRS-1C) has been measured from -10°C to +50°C using an electronic digital indicator. A suitable test rig was designed and fabricated to carry out this measurement in the required temperature range. The coefficient of the thermal expansion results obtained with this testing and technique are accurate to ± 13.4% and the repeatability of this measurement is better than ± 1.3%.

The total hemispherical emittance of different Thermal Control Coatings viz. Chemglaze Black Paint Z-306, 3M Black Paint, ECP-2200, Whittakar White Paint APA-247 4 and Astral Aluminium Paint PSG-173 being used in different spacecrafts for thermal control purpose were measured in the temperature range 115K-375K using steady state calorimetrie method. Measurements were done on a disc shaped sample with built in heater and an evacuated low temperature radiation receiver is used for the test environment. Experimental results and error estimation are presented here. The results indicate that the emittance increases with increase of temperature for all the coatings. Along with the experimental results, temperature dependence of the emittance relation for each of these coatings is also presented. Facilities and measurement technique used here are capable of providing the results better than ±3.0 % at the lowest temperature of the measured temperature range.

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.

A real time life cycle test, simulating geosynchronous orbit operation, on two Ni-Cd cells and one Ni - H2 cell is in progress. Some significant observations on their cycle performance are reported and discussed in this paper.

A passive thermal control system is provided for the temperature control of the panchromatic camera of IRS-1C spacecraft. The camera with a spatial resolution of better than 10m and a revisiting capability requires a stringent thermal control system. The thermal design is based on mathematical models. This paper gives the details of the model and the predicted temperature distribution of the camera.