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Bepi Colombo is an ESA mission targeted to the exploration of Mercury with two spacecraft, a Mercury Polar Orbiter (MPO) provided by Europe and a Mercury Magnetospheric Orbiter (MMO) provided by Japan. The Mission is lead by Astrium Friedrichshafen, with Thales Alenia space Torino responsible for the MPO thermal design. The MPO is a 3-axis stabilized scientific spacecraft in Mercury polar eccentric orbit, with altitude from 400 to 1500 km, with one face planet oriented and pointing Nadir, and housing the apertures of the observation P/L. Studies for this mission were initiated in the late 90ies, and pointed out that one of the main design drivers for the MPO was the thermal environment in orbit, due to the combination of high solar constant (up to 10 solar constants on Earth), infrared and albedo from the planet (maximum IR up to about 4 terrestrial solar constants, albedo up to about 1).

The SMOS satellite is a polar-orbit sun-synchronous Earth observation ESA mission, whose science objectives are to: globally monitor surface soil moisture over land surfaces, globally monitor surface salinity over the oceans, characterise ice and snow covered surfaces The SMOS satellite is composed of the Proteus platform and the Payload module/MIRAS instrument. This paper is dedicated to the thermal design and testing of SMOS payload module (PLM). The PLM thermal design is passive, maximizes the use of proven materials and processes and is supported by heaters. The major drivers for the design are the limitation in heater power allocation and the stringent temperature requirements. The verification of the PLM thermal design is based on a Thermal Balance (TB) testing of a Structural Thermal Model (STM), followed by a TB/TV test on the ProtoFlight Model (PFM). A Thermal Vacuum (TV) test has also been performed for the complete spacecraft.

Fuel cells have become a very important part of the energy supply in past, current and future space activities. In the past they have been used in the Gemini missions - their first application in space - and later in the Apollo programme. Fuel cells will be employed in future American, Russian and European space vehicles where significantly larger electrical power generation is envisaged compared to early applications. This increased capacity is linked to a correspondingly larger amount of waste heat which must be managed in turn. Effective study of how best to accommodate such fuel cell systems into a space vehicle and of the impact on other subsystems - e.g. thermal control - requires a suitable simulation tool.

This paper describes the main changes affecting the APM Environmental Control System (ECS) as a consequence of the Space Station Freedom (SSF) restructuring and Columbus APM overall reconfiguration. The main purposes of this reconfiguration are: minimize the number and complexity of the interfaces with Space Station Freedom (SSF) centralize avionics command and monitoring tasks revisit the failure tolerance concept of some ECS functions unify/standardize similar functions in the two subsystem adjust lifetime requirements and simplify maintenance concept of equipment. The APM ECS consists of the following functions: active thermal control (ATCS) passive thermal control (PTCS) atmosphere pressure and composition control air revitalization and cabin ventilation temperature and humidity control vacuum and venting nitrogen supply fire detection and suppression. The new ATCS configuration provides a cooling capability for a reduced number of P/L racks by means of its moderate loop.

Dedicated technology has been developed to support long-term biological experiments on-board spacecraft. These developments include a microgravity compatible tubular photo bioreactor for the cultivation of micro algae at very high biomass concentrations and with very high gas exchange rates, a microgravity compatible gas / liquid phase separator which also works as a pneumatic low shear-stress pump, a microgravity compatible dehumidifier, and a maltose separating reverse osmosis unit. Integration of these technologies into a partially closed artificial ecosystem form the foundation of the SYMBIOSE concept (System for Microgravity Bioregenerative Support of Experiments).

ESABASE/SUNLIGHT is a software tool to calculate illumination, effective illumination, exposure time, incident electromagnetic power, absorbed electromagnetic energy for spacecraft surfaces during planet orbiting missions considering sun and planet irradiation, effects of eclipse and self-shading, multireflections, transmission, pointing and (variable) geometry. Calculation applies a fast Monte Carlo raytracing algorithm and is based on wavelength dependent spectra and material properties. ESABASE/SUNLIGHT is fully integrated in the CAE-frame ESABASE which offers a powerful geometry specification language, orbit generator, pointing facility and advanced libraries as well as gateway, pre-, postprocessing and display tools with the benefits of standardisation and exchange to other analysis tools.

As it is not possible to directly use commercial liquid flow meters in spacecraft fluid loops, a study was carried out for the European Space Agency to adapt commercial flow meter assemblies for space applications. The various activities (described in detail) eventually led to the selection of two commercial units, which were redesigned/adapted to be used in spacecraft single-phase (water) and two-phase (ammonia) thermal control loops. These flow meter assemblies were tested according to an agreed test programme, that included performance and calibration tests in a test bench (developed during the study), vibration testing and EMC/EMI testing. The results are discussed in order to assess to what extent the study objectives were met. Recommendations for future work are given also.