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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.
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A review of NASA research programs related to developing and improving space flight technology is presented. Technical report topics summarized include: space flight feeding; aerospace medicine; reusable spacecraft; satellite soil, vegetation, and climate studies; microwave landing systems; anthropometric studies; satellite antennas; and space shuttle fuel cells.
NASA negotiations and contractual arrangements involving European space research organizations' participation in manned space operations and efforts in building Spacelab for the U.S. Reusable Space Shuttle are discussed. Some of the diplomatic and technical collaboration involved in the international effort is reviewed.
A spray coating apparatus is provided for rotating a workpiece relative to a spray station to obtain a uniform coating of the workpiece. In a typical example, the workpiece comprises a ceramic tile which is to be coated with a ceramic coating and the tile is to be used as a reusable component of the thermal protection system for a space shuttle. The apparatus for rotating the workpiece includes a base support having a first rotatable stage for rotation in the horizontal plane and a second rotatable stage for rotation in a second plane inclined at an angle, such as 45 degrees, to the horizontal plane and the workpiece is supported on this second stage. Thus the workpiece is rotatable in both of two planes of rotation.
A space based orbital transfer vehicle (SBOTV) and ground based OTV's (GBOTV) are compared for debris protection, space based OTV maintenance provisions, flight performance, onorbit refueling, and launch and return operations. Debris protection has a severe impact on the SBOTV, while the penalty for the GBOTV is much less severe. A key technology issue is the protection capability of composite materials. Onorbit maintenance is critical for SBOTV. Reduction of losses during the various transfers is th maine problem with refueling a SBOTV. Zero-g propellant storage and transfer is an important technology area for SBOTV. A reusable shroud must be developed to return GBOTV's if a Shuttle derivative vehicle is used. The advantage of space basing lies in more efficient use of the launch vehicle. Since most of the mass going to LEO is OTV propellant, and the launches to deliver the SBOTV propellant are generally mass limited, substantially fewer launches are required to support the SBOTV.
Flexible fin radiators are being developed to meet spacecraft requirements of light weight, compact launch volume, modular growth, reusability, and the capability of an on-orbit deployment and retraction. Two prototype versions of deployable/retractable, flexible finspace radiators (soft and hard tube) were tested simultaneously under thermal vacuum conditions. A map of the heat rejection performance of each radiator panel under various environmental conditions is given. The soft tube radiator deployment system demonstrated minimal gas pressure and was capable of maintaining partial deployment; temperature control was achieved by varying the deployment length. The radiator could be recovered after an inadvertant freezing of the coolant in the Teflon tube. The hard tube radiator deployment mechanism performed adequately throughout the testing, but did exhibit some undesirable performance characteristics. The hard tube radiator pressure drop was higher than expected, but was not affected by the length of deployment. The radiator operated at less than full deployment, and the manifolds were found to have a significant effect on heat rejection. At low-load conditions, the hard tube radiator appeared to bypass its deployed area.
A status report on the development of the Space Shuttle system is presented, with emphasis on the system capability, the major objectives accomplished during the development period, the Orbital Flight Test program, and the STS-1 mission profile. The system incorporates virtually all of the characteristics and requirements initially established as design goals over a decade ago: both the system and the solid rocket booster are reusable; the orbiter, a delta wing vehicle, is capable of aerodynamic cross-range in excess of 1,100 nautical miles and has an approximate 4,000 nautical mile down-range capability; the flight control system and software provide various abort modes, and the orbiter crew compartment is designed for a shirt-sleeve environment; the main liquid fuel engines provide 470,000 pounds of thrust at vacuum, and the payload lift capacity is 65,000 pounds to a 150 nautical mile orbit.
The evolution of reusable surface insulation is described, with attention to the mullite and other aluminosilicates, zirconia, and silicon carbide-coated carbon compositions experimented with in the late 1960s and early 1970s. Evaluation and development concentrated in this period on such aspects of design and material properties as cold soak performance, attachment, shock impingement, surface coating cracks, and tile gap design and heating. In addressing the central problem of heat-shield thermal conductivity, it was found that for a given density, silica fibers had a lower conductivity than those of mullite. This was due to the one-micron, as opposed to 4.7-micron, diameter of the silica fibers, which resulted in smaller pores and therefore less convective and radiative heat transfer. Attention is also given to tile coating materials and the high-temperature processes by which they were applied.
Examples of the contribution of microstructural analysis to the development of the Space Shuttle tile insulation system are presented, with photographic examples of the scanning electron microscope (SEM) investigations. After the basic thermal protection system materials had been selected, it was neccessary to analyze the mechanical responses of the combined materials; which included: (1) the polymer strain isolation pad (SIP), (2) the room temperature-vulcanizing silicone rubber bond, (RTV), and (3) rigid ceramic fiber reusable surface insulation (RSI). Microstructural analysis was used to provide information on deformation and fracture mechanisms, load transfer mechanisms, and structural alterations occurring before final failure. Both quantitative and qualitative information was obtained in the open, three-dimensional fibrous structures of the ceramic tiles by means of novel techniques of encapsulation and dissolution.
Spacelab is a manned, reusable laboratory which is being developed for the European Space Agency (ESA). In its working mode it will fly in low earth orbit in the cargo bay of the Shuttle Transportation System (STS) Orbiter. A description is presented of the structural development of the various features of Spacelab. System requirements are considered along with structural requirements, quasi-static loads, acoustic loads, pressure loads, crash loads, ground loads, and the fatigue profile. Aspects of thermal environment generation are discussed, and questions regarding the design evolution of the pallet structure are examined. Details of pallet structure testing are reported, taking into account static strength tests, acoustic tests, the modal survey test, crash tests, and fatigue/fracture mechanics testing.
The Long Duration Exposure Facility (LDEF) is a large (14 feet diameter by 30 feet long) Shuttle transported, reusable spacecraft. The LDEF can accommodate up to 13,000 pounds of experiments mounted in 86 standard trays. This paper describes the philosophy and approach used for the passive thermal design of the LDEF. Also discussed are the standardized guidelines and techniques used in the thermal design and integration of approximately 50 different thermally passive experiments. A technique for reducing multinode thermal models of experiments to 2 nodes is also presented. This approach allows the efficient thermal integration of large numbers of experiments with the LDEF spacecraft.
Thermal performance verification of Reusable Surface Insulation (RSI) has been accomplished by comparisons of STS-2 Orbiter Flight Test (OFT) data with Thermal Math Model (TMM) predictions. The OFT data was obtained from Development Flight Instrumentation RSI plug and gap thermocouples. Quartertile RSI TMMs were developed using measured flight data for surface temperature and pressure environments. Reference surface heating rates, derived from surface temperature data, were multiplied by gap heating ratios to obtain tile sidewall heating rates. This TMM analysis resulted in good agreement of predicted temperatures with flight data for thermocouples located in the RSI, Strain Isolation Pad, filler bar and structure.
Technical aspects of the alternate thermal protection system (TPS) study for the Shuttle Orbiter are reviewed, and a status report on alternate TPS technology developments is presented. Mission impact, life cycle costs and risks, and selected candidate concepts are identified. The best system would consist of mechanically attached metallic and carbon-carbon TPS concepts employing a titanium multiwall prepackaged concept at temperatures below 1000 F, a superalloy honeycomb prepackaged concept at temperatures between 1000-1800 F, and an advanced carbon-carbon multipost standoff concept above 1800 F. Alternative concepts offer significant improvements in durability and are mass competitive with current ceramic tile reusable surface insulation.
Data are presented from Space Shuttle flight measurements of the temperature distribution within the gaps of the reusable surface insulation tiles. This is the first of a series of flight measurements in which the gap width and tile surface edge radius will be systematically varied. The data show several interesting features including the time and effect of boundary-layer transition. The local heating rates into the gap surfaces are calculated from the measured temperature distribution down the gap at two locations.
The current development status of components to control the flow of propellants (liquid fluorine and hydrazine) in a demonstration space storable propulsion system is discussed. The criteria which determined the designs for the pressure regulator, explosive-actuated valves, propellant shutoff valve, latching solenoid-actuated valve and propellant filter are presented. The test philosophy that was followed during component development is outlined. The results from compatibility demonstrations for reusable connectors, flange seals, and CRES/Ti-6Al4V transition tubes and the evaluations of processes for welding (hand-held TIG, automated TIG, and EB), cleaning for fluorine service, and decontamination after fluorine exposure are described.
The Space Shuttle stellar-inertial reference system is a velocity and attitude data source during flight operations. An overview of the reference system is presented as well as specifics discussing design concepts, functional operation, and performance capabilities. Techniques for star sighting and inertial measurement unit alignment and calibration are described, and alignment accuracy, star tracker capability, and gyro and accelerometer accuracy are discussed, with emphasis on flight test results. Test programs have confirmed that the system meets performance requirements such as being accurate to within 0.26 degree at the 400,000 foot altitude entry interface in order to execute an accurate touchdown, as well as demonstrating reusability, payload capability, and operational flexibility. Growth possibilities, such as the implementation of rendezvous target tracking, are discussed.
The Orbital Transfer Vehicle (OTV) is an advanced upper stage concept which will deliver spacecraft from operating systems at Low Earth Orbit (LEO) such as Space Shuttle, Earth-To-Orbit (ETO) vehicles, and Space Operations Center (SOC), to High Earth Orbit (HEO) and planetary excursions. The OTV will be driven by the need to achieve significant reductions in the operational costs for delivering payloads to Geostationary Equatorial Orbit (GEO). Aeroassist is a technological capability that has a potential for OTV's ranging from mission enhancing (reusable OTV for payload delivery) to mission enabling (manned GEO and some DOD). It is shown that the use of aeroassist for OTV's is a high leverage technology which can potentially reduce space transportation costs and enable a number of highly desirable missions.
Information from the Survey of Spaceflight Safety Systems considered most applicable to present and future spaceflight activities is summarized. Mission reports for the first three Shuttle Orbiter flights discuss all anomalies which occurred and their resolution, such as the solid rocket booster ignition overpressure spike, and thermal tile damage. Two emergency rescue systems for orbiting space stations are proposed: a dish type rescue capsule and a reusable orbit vehicle. Prebreathing and stage decompression are considered as means of protecting astronauts against decompression effects. The Spacelab Space Sled experiments are described, as is a test program on the assembly of large space structures completed in mid-1981. Finally, collision hazards for space missions are estimated, and operational procedures for minimizing these hazards are discussed.
A concept has been developed for providing for a permanent manned presence in space for science, applications, and technology plus growth potential to support major future operational missions. An initial crew of two to four is accommodated along with installations for interior, exterior and tended payloads. Tended payloads support includes servicing of remote free-flying spacecraft, basing of reusable orbital transfer vehicles and basing of teleoperator-mounted payloads for near-station excursions. Candidate payloads, plus configuration, flight operation, subsystem and ground logistics aspects are described, as are Shuttle interactions.