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This part of the manual presents methods for arriving at a solution to the problem of spacecraft inflight equipment environmental control. The temperature aspect of this problem may be defined as the maintenance of a proper balance and integration of the following thermal loads: equipment-generated, personnel-generated, and transmission through external boundary.

The ROSETTA mission is a European Space Agency spacecraft which will rendezvous with the comet Wirtanen. The spacecraft will deploy a Surface Science Package (SSP) to land on the comet. ...Thus, the heater power demand at 5.25 AU has to be low, when the spacecraft is in Deep Space Hibernation mode. Prior to the comet rendezvous, the spacecraft is woken up. ...With a fully operational payload, the ROSETTA spacecraft establishes an orbit around comet Wirtanen. The spacecraft follows the comet during its approach to the sun.

Spacecraft have previously used passive cooling or pumped liquid loops to maintain proper thermal environments. ...Future large spacecraft, such as the Space Station, will have higher power levels and longer heat transport distances.

A high-temperature, low-power electrochemically-driven fluid cooling pump is currently being developed by Lynntech, Inc. With no electric motor and minimal lightweight components, the pump is significantly lighter than conventional rotodynamic and displacement pumps. Reliability and robustness is achieved with the absence of rotating or moving components (apart from the bellows). Lynntech has recently demonstrated the feasibility of long term pump operation at temperatures of up to 100 °C, and extended storage at temperatures as low as -60 °C. Characteristics of the electrochemically-driven pump are described and the benefits of the technology as a replacement for electric motor pumps in mechanically pumped single-phase fluid loops (MPFL), such as that used in the Mars Pathfinder (MPF), is discussed.

A design study of pumped two-phase and capillary-pumped thermal control systems (TCS) was conducted on a typical, advanced earth-orbiting spacecraft (AEOS). NASA's Upper Atmosphere Research Satellite (UARS) size and baseline design were chosen as the configuration for AEOS; however, the power requirements were increased and the allowable temperature range was decreased to represent the requirements of a more advanced spacecraft. ...NASA's Upper Atmosphere Research Satellite (UARS) size and baseline design were chosen as the configuration for AEOS; however, the power requirements were increased and the allowable temperature range was decreased to represent the requirements of a more advanced spacecraft. For peak-power dissipations of 1 or 2 kW, the capillary-pumped system was lightest. ...In addition to the weight advantage, the more nearly isothermal spacecraft obtainable with the two-phase system permits less compensation for thermal distortion; a simpler, less costly structure results.

Not only the detectors are cooled, but also major subsystems and systems of the spacecraft’s. The Centre Spatial de Liège (CSL) is involved in the testing of several parts of the spacecraft’s, starting from optical tests on the mirrors or on the telescopes, going on with cryogenic vibration testing of scientific focal plane instruments, ending with the full Planck spacecraft testing. ...The Centre Spatial de Liège (CSL) is involved in the testing of several parts of the spacecraft’s, starting from optical tests on the mirrors or on the telescopes, going on with cryogenic vibration testing of scientific focal plane instruments, ending with the full Planck spacecraft testing. ...The Centre Spatial de Liège (CSL) is involved in the testing of several parts of the spacecraft’s, starting from optical tests on the mirrors or on the telescopes, going on with cryogenic vibration testing of scientific focal plane instruments, ending with the full Planck spacecraft testing. Each test requires temperature lower than 20 [K], in volumes ranging from 1 [m3] to 60 [m3], cooling down several kilograms to more than one ton, and withstanding heat load up to 150 [W] in stabilization.

The technique uses the standard SINDA thermal-analysis program and thereby extends the capabilities of SINDA to complex, active spacecraft thermal-control systems. This paper provides sufficient detail that a current SINDA user will be able to apply the technique by reference to this paper alone.

Multilayer Insulation (MLI) is used for thermal control of spacecraft (S/C). Built-in discontinuities degrade the performance of MLI blankets. In this paper the blanket properties are considered by defining an effective emissivity for the MLI covered S/C surface.

A computer-based design system has been developed for conceptual design of spacecraft active thermal control systems (ATCS). Its purpose is to aid the thermal systems engineer in selection of a working fluid and thermal cycle for a given set of requirements. ...It is found that reductions in power system weight penalty result in considerable weight savings for vapor-compression systems, making them more attractive for future spacecraft. The advantage of the present approach lies in the capability to accomodate quantitative technical information as well as environmental and safety considerations in a single model.

This paper describes the development of a microcomputer based control system for a heat pump containing an electrical variable speed compressor drive and a motorized expansion valve. It is designed to operate under very much varying load conditions with minimum power consumption. Difficulties that were encountered during engineering tests could finally be overcome by a relatively simple, practical regulator configuration. It operates near optimum efficiency by regulating a certain temperature difference in the evaporator. Experimental data on operating characteristics and performance are included in the paper.

These models took advantage of the submodel capability of TSS and SINDA/FLUINT providing a simplified approach for merging spacecraft and instrument models. In addition to the spacecraft thermal model, there is the Advanced Land Imager (ALI) instrument model by MIT/LL, the Hyperion instrument by TRW, the Atmospheric Corrector (AC) instrument by GSFC, and the New Millenium Program (NMP) experiments. ...The thermal design and analysis of the Earth Observing-1 (EO-1) spacecraft, built by Swales Aerospace for NASA's Goddard Space Flight Center (GSFC), consisted of a Thermal Synthesis System1 (TSS) geometric math model (GMM) and a SINDA/FLUINT2 thermal math model (TMM). ...EO-1 will fly in a sun-synchronous, 705 km circular orbit, one minute behind the Landsat-7 Spacecraft. Using data from Landsat researchers can validate the ALI, Hyperion, AC and NMP technologies performance over EO-1's one-year primary mission duration.

The progress in the development of a sensor for the detection of trace air constituents to monitor spacecraft air quality is reported. A continuous wave, external-cavity tunable diode laser centered at 1.55 μm is used to pump an optical cavity absorption cell in cw-Cavity Ringdown Spectroscopy (cw-CRDS).

The European Space Agency is presently funding the development of a new facility which could be used for the thermal control of : automated space platforms such as EURECA the modules of the future space station other manned space vehicule such as HERMES. The reasons for the development of such a facility are presented and the design requirements are given, mainly the pressure head-flow rate characteristics and the noise levels. The potential near future users are mentionned. A description of the Freon Pump Package is given including a lay-out, a mass breakdown, a system block diagramme, together with the design objectives and the design philosophy to achieve a very low level of noise. A special emphasis is given to the flow rate regulation modes which have been incorporated in the FPP design. The paper concentrates on the tradeoffs performed to select the main components of the pump unit which is the central active part of the FPP.

Future long term missions in space require the development of a thermal management system capable of collecting, transporting and rejecting waste heat from various components in the nominal temperature range of 0-100°C. A summary of four, two-phase heat transport systems currently under consideration at NASA-JSC is presented. The comparative merits of four systems, which include a mechanically pumped two-phase loop, a capillary pumped loop, an osmotic pumped loop and a loop using a biomorph pump, are presented based on a comprehensive literature review and contractor reports supplied by NASA-JSC. A recommended conceptual design for a thermal utility using a capillary pumped loop is proposed along with an evaluation of its strengths and weaknesses. System start up and control are discussed and areas of continued investigation are identified.

It is proposed the approach to the analysis of SC's thermal condition in orbital flight, based on study of its signal characteristics and thermal optical portraits, which are gotten with various resolutions. The given approach consists of analysis and comparison of measurements in various bands of spectrum of radiation and reflection characteristics of investigated SC, with calculated and experimental data. The calculated estimations of SC's infra-red radiation were carried out for the occurrence on it the typical non-standard situation. In this paper it is proposed the criteria, which can be used during the fulfillment of distant analysis of SC's functional condition on the base of the data of observation of its signal optical characteristics.

This part of the manual presents methods for arriving at a solution to the problem of spacecraft inflight equipment environmental control. The temperature aspect of this problem may be defined as the maintenance of a proper balance and integration of the following thermal loads: equipment-generated, personnel-generated, and transmission through external boundary.

The Mars Pathfinder spacecraft which will be launched in December 1996 features an active cooling system for controlling the temperature of the spacecraft. ...The Mars Pathfinder spacecraft which will be launched in December 1996 features an active cooling system for controlling the temperature of the spacecraft. This will be the first time that such a mechanical pump cooling system is used on an interplanetary or long duration flight (over two weeks) in space. ...It uses centrifugal pumps to circulate liquid freon to transfer heat from spacecraft electronics to an external radiator. The IPA consists of redundant pumps, motor control electronics, thermal control valves, check valves, and an accumulator.

Mechanically pumped single-phase fluid loops are well suited for transporting and rejecting large amounts of waste heat from spacecraft electronics and power supplies. While past implementations of such loops on spacecraft have used moderate operating temperatures (less than 60ºC), higher operating temperatures would allow equivalent heat loads to be rejected by smaller and less massive radiators.

Electrical disturbances caused by charging of cables in spacecraft can impair electrical systems for long periods of time. The charging originates primarily from electrons trapped in the radiation belts of the earth. ...The model Space Electrons Electromagnetic Effects (SEEE) is applied in computing the transient charge and electric fields in cables on spacecraft at low to middle earth altitudes. The analysis indicated that fields exceeding dielectric breakdown strengths of common dielectric materials are possible in intense magnetic storms for systems with inadequate shielding.

While an extensive database of testing for arc track resistant wire insulations has been developed for aircraft applications, the counterpart requirements for spacecraft are very limited. This paper presents the electrical, thermal, mechanical, chemical, and operational requirements for specification and testing of candidate wiring systems for spacecraft applications. ...This paper presents the electrical, thermal, mechanical, chemical, and operational requirements for specification and testing of candidate wiring systems for spacecraft applications.