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Loop heat pipes (LHP) combine the classical heat pipe operation principle with a loop configuration, i.e. evaporator and condenser are located at separate places as dictated by thermal and configurational reasons and are connected lines for fluid and vapor transport. Experimental investigations of such LHP revealed certain temperature oscillations under certain conditions. This paper presents the results of analyses of such temperature fluctuations and offers possible explanations for initiation and termination of such behavior. Two types of temperature fluctuations have also been confirmed experimentally by testing different kinds of LHP.

COLUMBUS represents the European participation in the US Space Station (USSS). As a long term goal this program will provide a comprehensive autonomous European Space Station capability with the requisite in-orbit infrastructure.

The European Retrievable Carrier (EURECA) is the first European reusable experiment platform, which is scheduled to be transported in 1988 into orbit by NASA's Space Shuttle. EURECA is designed for five free-flying missions in a 500 km and 28.5° inclination orbit, each with a 6-month operational and a 3-month dormant phase. The capabilities of the baseline Thermal Control Subsystem design have been investigated for future missions with regard to extended lifetime as well as applying a purely passive design according to the mission requirement. This paper describes the thermal concept for various EURECA missions expected to be realized in the near future. Moreover, a general outlook on future EURECA flights is presented. The work has been performed under the contract of the European Space Agency (ESA).

The most, often utilized method of thermal design verification for past and current types of spacecraft has been by Solar Thermal Vacuum (STV) tests at system level. This approach, however, will require fundamental modifications for future large and complex spacecraft, three types of which are defined; a large communication satellite, a modular earth observation satellite and a large infrared telescope, all exceeding the capabilities of existing STV test facilities. A trade-off is made between analysis only and analysis plus test, considering the influence of temperature limits, analysis uncertainty and correlation of analysis and test, as well as indicating the ways in which it affects the thermal design verification. Alternative test methods - STV testing in parts, infrared radiation tests, thermal canister method, skin heater application - are also discussed with regard to their advantages and disadvantages of applying them to the large types of spacecraft.