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This paper describes a computational technique for prediction of the flow and thermal environment in the Space Shuttle Solid Rocket Motor field joint cavities. The SRM field joint hardware has been tested with a defect in the insulation. Due to this defect, the O-ring gland cavities are pressurized during the early part of the ignition. A computer model has been developed to predict the flow and thermal environment through the simulated flaw, during the pressurization of the field joint. The transient mass, momentum, and energy conservation equations in the flow passage in conjunction with the thermodynamic equation of state are solved by a fully implicit iterative numerical procedure. Since this is a conjugate flow and heat transfer problem, wall temperatures are calculated by solving the one-dimensional transient heat conduction equation in the solid along with the other governing equations. The pressure and temperature predictions have been compared with the test data.

The design results for a 250 kW switched reluctance aircraft engine starter/generator system power inverter are presented. The starter/generator employs a single switched reluctance machine and a generating system architecture that produces two separate 270 Vdc buses from that single switched reluctance machine. The machine has six phases with three of the phases connected to one inverter supplying 125 kW to one 270 Vdc bus while the other three phases are connected to a second inverter supplying 125kW to the other 270 Vdc bus. Each bus has its own EM1 filter and control in addition to its own inverter. Two types of inverters have been developed, one type employs MOS Controlled Thyristors (MCTs) for the controlled switches and the other type employs Insulated Gate Bipolar Transistors (IGBTs). High-current 500 A peak turn-off MCT modules were specifically developed for the MCT inverters. Two of these modules are placed in parallel to form the required 1000 A switches.

Thermo-mechanical fatigue and natural aging due to environmental conditions are difficult to simulate in an actual test with the advanced fiber-reinforced composites, where their fatigue and aging behavior is little understood. Predictive modeling of these processes is challenging. Thermal cyclic tests take a prohibitively long time, although the strain rate effect can be scaled well for accelerating the mechanical stress cycles. Glass fabric composites have important applications in aircraft and spacecraft structures including microwave transparent structures, impact-resistant parts of wing, fuselage deck and many other load bearing structures. Often additional additively manufactured features and coating on glass fabric composites are employed for thermal and anti-corrosion insulations. In this paper we employ a thermo-mechanical fatigue model based accelerated fatigue test and life prediction under hot to cold cycles.

A vapor-compression refrigerator/freezer with solar power and grid-backup power was tested at Johnson Space Center. The refrigerator was used to test three main energy saving devices: 1) solar power with grid-backup, 2) vacuum panel insulation, and 3) phase change material. Refrigerator power consumption and temperature were monitored while various configuration changes were made to the refrigerator. The testing of the refrigerator showed that the concepts had the potential to save energy although several drawbacks were discovered as a result of the tests.

This technical overview of the Passive Thermal Control System (PTCS) for the Space Station Freedom (SSF) identifies the need for a Passive Thermal Control System, presents PTCS design concepts, highlights significant design requirements, addresses important design factors and how design goals are accomplished and summarizes development testing results as well as future development test plans. A Passive Thermal Control System is an essential component in ensuring the survivability of Space Station Freedom in the hostile space environment for a design life of 30 years. Design requirements specify a thermal environmental range of -250° F (-156.7°C) to +300° F (148.9°C), which accounts for the range of surface temperatures that can be encountered during orbit, and a maximum acceptable thermal leak from the pressurized modules.

This specification covers a polychloroprene (CR) rubber, resin-modified, solvent-type contact adhesive in the form of a liquid.

This specification covers an adhesive in the form of a liquid or paste.

This specification covers an adhesive in the form of a liquid or paste.

This specification covers a high-temperature, electrical-grade, polyimide resin adhesive in the form of film or paste.

No scope available.

No scope available.

No scope available.

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No scope available.

This document presents information concerning environmental control requirements peculiar to cargo and high density transport airplanes which is not covered in ARP 85 (Reference 1).

This document does not dictate a specific design approach or technology, but rather it provides design consideration to assist the specification writer in establishing a fail/safe design.