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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. SEEE also computes the minimal shielding needed to keep the electric fields below that for dielectric breakdown.

The uncontrolled flow of cabin air into the aircraft envelope caused by indoor-outdoor temperature gradients and associated stack pressure differentials causes cabin air to circulate between the cabin and the envelope, producing condensation on the cold fuselage behind the insulation with a number of adverse consequences. These include an inability to practically maintain cabin humidity at normal levels, a reduction in ventilation effectiveness, microbial growth, metal corrosion and structural failures, insulation performance degradation, thermal discomfort, increased engine noise transmission, additional fuel consumption and electrical failures and fires.

Manned tests in Chamber B at NASA JSC were conducted in May and June of 1990 to better quantify the Space Shuttle Extravehicular Mobility Unit's (EMU) thermal performance in the cold environmental extremes of space. Use of an infrared imaging camera with real-time video monitoring of the output significantly added to the scope, quality and interpretation of the test conduct and data acquisition. Results of this test program have been effective in the thermal certification of a new insulation configuration and the “5000 Series” glove. In addition, the acceptable thermal performance of flight garments with visually deteriorated insulation was successfully demonstrated, thereby saving significant inspection and garment replacement cost. This test program also established a new method for collecting data vital to improving crew thermal comfort in a cold environment.

The aim of this paper is to present the Hermes Thermal Control Concept which constitutes the reference baseline for the phase 2 of the Hermes program. The technical solutions based on active and passive means, are dictated by a variety of requirements including temperature and heat flux limits, mass & power minimization, availability of space for accommodation of TCS devices. Furthermore thecomplex mission profile (including atmospheric flight and several modes of operation o002d;orbit) and the configuration of the space-vehicle, require the adoption of a particularly flexible thermal control. The ATCS (Active Thermal Control Section) is based on a dual loop concept, using both water and Freon R114 cooling loops for heat collection from the various sources, heat transportation and heat rejection through dedicated devices.

The two primary wire construction types being used in military aircraft today are cross-linked Ethylene Tetrafluoroethylene (XL-ETFE) and composite fluoropolymer / polyimide tape wrap with an outer Polytetrafluoroethylene (PTFE) tape wrap. These insulations offer significant improvements over earlier polyimide (MIL-DTL-81381) and polyvinyl chloride (PVC) constructions but are not without drawbacks. XL-ETFE provides a low smoke, high fluid resistant, non-arc tracking insulation that is durable during installation and ground repair operations. However, durability and abrasion resistance are reduced at elevated temperatures, and maximum operating temperature peaks at 200 degrees Celsius. Composite insulation provides a more abrasion resistant solution with the inclusion of polyimide tape for hard surface chafe conditions and a PTFE outer layer that improves wear life during wire to wire contact.

Piezoelectric pumps offer great potential as an alternative electro-mechanical actuator and as a hydraulic power source. As an actuator, this pump may provide solutions to control system problems in robotics, process control, bioengineering, advanced remote control (telepresence), and automation. As a hydraulic power source they may be useful for active thermal cooling, fluid management, and metering pumps in life support applications. The benefits of piezoelectric based pumps and actuators include increased efficiency, self-cooling, lightweight, compact size, high mechanical reliability, positive displacement, self-priming, no lubrication, no vibration, and rotational inertia. Oceaneering Space Systems (OSS) has produced two successful piezoelectric pump prototypes. The first one is a double-acting diaphragm pump driven by piezoelectric PolyVinylidine DiFluoride (PVDF) polymer. The second prototype is a Lead Zirconate Titanate (PZT) thermoplastic laminated pump.

COROT mission is managed by CNES (French National Space Agency) in association with three major French laboratories (LAM, LESIA, IAS) and several European countries, contributing to the payload and the ground segment. This astronomy mission objectives are astero-seismology as well as planet finding. The COROT spacecraft is based on a PROTEUS low Earth orbit recurrent platform, developed by CNES and Alcatel Alenia Space. It was injected on 27th December 2006 at a 898 km polar and circular orbit by a Soyuz launcher and is being operated from CNES-Toulouse. This paper focuses on the thermal control design and first in-orbit performances of the payload which mainly consists in an afocal telescope, a wide field camera with cooled CCDs, and an equipment bay. Largely using standard and well-proven technologies, this paper also points out some thermal control specificities and techniques used.

The requirements related to the thermal control of HERMES space-plane suggest the utilization of technical solutions based on active and passive techniques. The configuration of the space-vehicle with pressurized and unpressurized compartments presents a variety of different problems, which generally require dedicated solutions. The complex mission profile include atmospheric and orbital conditions, with several phases and modes of operation (free-flying and docked). The extremely wide range of environmental conditions, together with the timeline of internal power dissipations requires the adoption of a particularly flexible thermal control. The ATCS (Active Thermal Control Section) mainly relies on the capability of water and Freon 114 fluid cooling loops. They collect waste heat from the sources located inside the compartments and transport this power to the available rejection devices.

This paper summarizes some of the efforts to advance hypersonic structures technology through the National Aero-Space Plane (NASP) Structures Technology Maturation Program. The ranges of expected structural loads and results from analysis and test activities are described. Topics briefly covered include shock impingement effects on aerothermal loads, actively-cooled structures concepts and test results to date for leading edges and panels, design and fabrication of a carbon-carbon control surface, a program for predicting performance of metal matrix composites, an analysis and sizing procedure for thermal structures, and some recently-developed test fixtures for seals, thermal insulation, and actively-cooled panels.

Fire barrier coatings are being used to protect nacelle and cowling components by insulating the surfaces from excessive heat exposure during normal flight operations and to provide a fire barrier in the event of an engine fire. Design criteria, formulating, considerations, application methods, and typical performance characteristics for such systems will be discussed.

This study investigates the effect of Teflon and adhesive coatings on the torque-tension relationship and the self-loosening performance of threaded fasteners. Two Teflon insulation coatings and one locking adhesive are considered. The torque-tension relationship is established for coated and uncoated fasteners for both tightening and loosening. Finally, the fasteners are tested to determine their self-loosening performance under cyclic transverse loads. A computer controlled fastener tightening system is used to establish the torque-tension relationship during tightening. The coefficients of thread and bearing friction, and the overall nut factor are measured. The breakaway loosening torque of tightened bolts, along with the coefficients of thread and underhead friction and nut factor are investigated. A modified Junker machine is used to evaluate the self-loosening performance of fasteners with various coatings.

Maintaining thermal balance in a space-based plant chamber has proven to be difficult to achieve, particularly if the air temperature in the plant chamber is desired to be below that of the atmosphere of the space vehicle. Analysis of the thermal condition of a plant chamber has identified three heat sources as major contributions to this serious problem. The first is the input of radiant energy into the chamber required to support plant growth. The second is via thermal conduction through the chamber walls. The last major thermal input is from the fans and other electronic components embedded inside the chamber. Design solutions to achieve thermal balance are further exacerbated by virtue of the limited power availability, volume and mass restrictions, and safety considerations.

Through the development of 35 spaceflight payloads during the last ten years, BioServe Space Technologies has gained valuable practical experience in developing thermal control systems for the microgravity environment. Design constraints imposed by NASA, such as limited power availability, limited material selections, and limited acoustic emissions, coupled with the design constraints imposed by the functional requirements of each payload, impact spaceflight designs in a manner that requires a high degree of optimization. BioServe payloads typically employ thermoelectric coolers (TEC's), air and liquid heat exchangers, a variety of insulation materials, several types of fans and blowers, and various control strategies in order to achieve the desired thermal environment. In the present work methods of selecting thermal system components are discussed.

The numerical/experimental method for, and results of, evaluating insulation solar radiation transmission effects on in-flight temperature field of a space vehicle are presented. The experiment subsection describes a new experimental method for evaluating materials solar-radiation transmittance, reflectance, and absorptance, as well as demonstrates test results for silica samples with various thicknesses. The calculation subsection includes (1) interpolation of the experimental data into a function of the insulation layer thickness and (2) the numerical solution of the transient heat conduction equation for a thermal protection fragment consisting of an insulation layer cemented to a metal plate. Fragment temperature calculation results for various near-Earth orbits and fibrous silica insulation layer thicknesses are presented. The results obtained show the essential influence of insulation solar-radiation transmission on space vehicle structure temperature in space flight.

The classical fencing problem occurs in radiant heat-transfer computations when a surface extends from one compartment to another, with the two compartments otherwise exchanging little heat. The surface that separates the two compartments is called a “fence.” If the gap between the bottom of the fence and the surface that extends under the fence is small, the potential for a large fencing error is evident from an examination of the drawings. In large models, with many surfaces forming many compartments, the fencing error is less evident. In this paper we examine the fencing errors in two prototype geometries. If the fenced surface is adiabatic, the error is found to be significant for surprisingly large gaps. A surface can be adiabatic due either to a high reflectance or a layer of insulation. The error is found to become insignificant when there is no reflectance.

Mathematical models of the combined thermal protection systems of the radiation-evaporation type are proposed. The systems considered consist of an outer heat-resistant screen sensitive to force loading, a high-temperature porous insulation, and a layer of a material saturated with a coolant and arranged on the external surface of the protected structure. Appropriate coupled boundary-value problems describing heat and mass transfer in these systems are formulated and solved. Combined and passive radiation thermal protection systems are compared. It is shown that both the thermal insulation thickness and the total weight of the combined heat protection systems may be considerably less then the passive equivalents.

The numerical/experimental method for evaluation of radiant heat transfer and shields temperature as well as investigation results obtained for two variants of multi-layer insulation systems, irradiated by solar radiation, are presented. The experiment subsection describes the test method and examination of a test model, consisting of some (3-5) shields by means of irradiating one by known radiant flux. The test method permits to detect and evaluate the transparency effects of shields relatively external, in particular, solar radiation. On the base of the test results the temperatures and the effective radiant characteristics of shields are determined. The numerical alghorithm provides (with taking into account the above experimental data) evaluating the temperature of shields and radiant heat fluxes through multi-layer insulation systems with arbitrary number of shields.

NIRSpec is a near-infra-red spectrometer and one of the four instruments onboard the James Webb Space Telescope (JWST). The JWST observatory will be placed at the second Lagrange point (L2). The instrument will be operated at about 30 Kelvin. Temperature stability and controlled heat rejection to dedicated JWST radiators are important issues of the NIRSpec thermal design. Besides thermal insulation, the NIRSpec Optical Assembly Cover also has to provide light tightness and stray light suppression to prevent unwanted light entering the instrument. Air tightness is needed to allow a controlled purge gas flow for contamination prevention while allowing proper air venting during launch. Because of mass constraints a cover employing two-foil Kapton blankets supported by aluminum posts and a wire tent was chosen. Failure tolerance and cleanliness are other important design drivers. This paper describes the design solutions established to fulfil the contrary requirements

A unique combination of manikin measurement and mathematical simulation was used to evaluate the thermal performance of two prototype modular boot systems intended to provide protection across a range of environmental conditions. Measured thermal insulations ranged from 0.91 to 1.85 clo (1 clo = 0.155 m2°C/W), and predicted endurance times (e.g. time for the toe temperature to reach 5°C) ranged from 56 to 169 min at -20°C, from 40 to 102 min at -30°C, and from 27 to 62 min at -50°C. Both systems provided appropriate thermal protection in a wide range of environmental conditions. The approach is time-saving and cost-effective, and is useful to understand the physical characteristics and thermal protection afforded by complex modular footwear, and better meet the needs of the customer.

In the frame of contracts to ESTEC an advanced adaptive high temperature insulation and an innovative safety enhancing secondary protection have been developed and tested by HPS. Both solutions can be used together to create a new thermal protection concept in order to make it lighter, cheaper and safer.