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High accuracy measurements of the earth’s gravitational field by satellites are affected by noise due to micro-vibrations caused by the environment and internal sources. In order to derive the best possible gravity field model the effects of non-gravitational accelerations have to be avoided and, where this is not possible, compensated or minimized. Among many other sources of those undesired disturbances are also classical thermal hardware items. This paper describes the development test programme established within the GOCE project to characterize the performance of MLI blankets, which were identified as potential micro-disturbance sources. The outcome of these tests is briefly discussed and the resulting selection of the flight thermal hardware is presented.

The Envisat spacecraft has a launch mass greater than 8.0 tonnes and external dimensions of 10.0 metres x 2.8 metres x 2.1 metres. Due to it’s large size, it was necessary to perform the thermal balance and thermal vacuum testing in two stages. Firstly, there was the testing of the Service Module and, secondly, the testing of the Payload Module (PLM). This paper discusses the thermal balance testing of the PLM. The PLM, itself, is 7.5 metres tall; too large to fit into a test facility solar beam. Originally, it was intended to conduct two solar beam tests; one for the upper part and the second for the lower part. Following a revision and re-scheduling within the programme, it was decided to perform a single, non-solar beam, thermal balance test. The thermal balance test would be performed using test specific, electrical, heaters and test facility shroud control.

The first use of the Polar Platform (PPF) is for the Envisat/PPF mission. The Envisat/PPF spacecraft has a launch mass of 8.5 tons and external dimensions of 10.0 metres x 2.8 metres x 2.1 metres. Due to it's large size it was necessary to perform the thermal balance and thermal vacuum testing in two modules. The first test was for the Service Module (SM) and the second for the Payload Module (PLM). This paper discusses the thermal balance testing and subsequent correlation of the Polar Platform Service Module thermal mathematical model.

The Polar Platform is a multi-mission platform designed to operate from low earth polar orbits. Its purpose is to accommodate and support various instrument complements dedicated to meteorology, earth observation and science. In its first mission configuration, ENVISAT-1, dedicated to earth observation, the lower three sections of the Payload Module represent the Payload Equipment Bay (PEB), which houses the payload support subsystem units and various payload instrument electronics equipment. All units are mounted on to the internal faces of the PEB CFRP honeycomb side panels. The PEB thermal design is basically passive, assisted by heaters and in general uses conventional techniques. However, thermal doublers, used to spread high heat fluxes over a greater area are made for the first time from carbon/carbon. This material matches the low coefficient of thermal expansion of the CFRP panels and at the same time provides the high thermal conductivity required for this application.

The ESA Polar Platform, as part of the ESA Columbus Development Programme, is scheduled to be launched as single passenger by an Ariane 5 vehicle in mid 1998. The multimission platform is designed to accommodate a wide range of payload complements to be flown on a series of missions in order to satisfy the growing future earth observation needs in continuation of the current ERS programme. Multi-mission capability is achieved by design modularity wherever feasible and cost-effective. This paper describes the thermal control design of the Polar Platform which follows its modular configuration and which has to cope with a wide range of generic performance parameters, whilst being adaptable to provide optimised performance for specific missions. Special thermal control features are highlighted as the software and hardware controlled heater systems, thermal doublers using carbon / carbon material and the battery compartment heat pipe radiator.

The “new generation” of large satellites like ERS-1 requires modular thermal balance testing due to the physical size. The purpose of this paper is to outline the experience gained from the ERS-1 Payload Thermal Balance Test. The first part of the paper highlights the test set-up, the earthshine compensation and the selected test phases. The second part describes the temperature uncertainty approach and test correlation criteria defined for the thermal analyses and tests. The third part concentrates on the test correlation with emphasis on the thermo-optical properties of the Optical Solar Reflectors (OSRs) in the Xenon light of the simulated sun and the temperature dependent linear conductance of the honeycomb core material which played a crucial role in explaining a temperature level offset. The paper is understood as complement to the paper presented in 1987 - Thermal Control and Design of the ERS-1 Payload -.

ERS-1 is the first European Remote Sensing Satellite scheduled to be launched early in 1990. While the Platform is based on the Multi-mission Platform developed in the frame of the French SPOT program, the Payload is a new mission specific development. This paper describes the severe requirements imposed by the ERS-1 mission on the PAYLOAD thermal control design and the adopted design solutions concentrating on several interesting features of its most important element the Payload Electronics Module which houses the bulk of the electronic equipment of the instruments. The payload thermal design has been completed and awaits verification by thermal balance testing early in 1988.