Engineering topics

Curry, Donald M.

Publications and source records attributed to Curry, Donald M..

27 records · Page 2

NASA NTRS1994

Analysis of the Shuttle Orbiter reinforced carbon-carbon oxidation protection system

Reusable, oxidation-protected reinforced carbon-carbon (RCC) has been successfully flown on all Shuttle Orbiter flights. Thermal testing of the silicon carbide-coated RCC to determine its oxidation characteristics has been performed in convective (plasma Arc-Jet) heating facilities. Surface sealant mass loss was characterized as a function of temperature and pressure. High-temperature testing was performed to develop coating recession correlations for predicting performance at the over-temperature flight conditions associated with abort trajectories. Methods for using these test data to establish multi-mission re-use (i.e., mission life) and single mission limits are presented.

Williams, S. D.↗

NASA NTRS1993

Space Shuttle Orbiter thermal protection system design and flight experience

The Space Shuttle Orbiter Thermal Protection System materials, design approaches associated with each material, and the operational performance experienced during fifty-five successful flights are described. The flights to date indicate that the thermal and structural design requirements were met and that the overall performance was outstanding.

Curry, Donald M.↗

NASA NTRS1993

Thermal Protection System design studies for lunar crew module

The results of a study to predict aeroheating and Thermal Protection System (TPS) requirements for manned entry vehicles returning to Earth from the moon are presented. The effects of vehicle size and lunar-return strategies on the aerothermodynamic environment and TPS design were assessed. Study guidelines were based on an Apollo Command Module (CM) configuration and lunar return strategies included direct entry and aerocapture followed by Low Earth Orbit entry (LEO). Convective heating was obtained by the boundary layer integral matrix procedure (BLIMP) code, and radiative heating was computed with the QRAD program. The AESOP-STAB code and the AESOP-THERM code were used for TPS analysis for ablating materials and nonablating materials respectively. Results indicated that there was an optimum size for minimum heating and that direct entry had higher heating rates than aerocapture. Aerocapture resulted in higher heat loads and TPS weight. The TPS weight factor was 6-8 percent for all lunar return strategies, with the TPS weight being about 50 percent less than that of the original Apollo CM vehicle.

Williams, S. D.↗

NASA NTRS1993

Space assembled entry systems certification

The approach taken to the issue, that 'How do you say you're 'good for go' if you space assemble an entry vehicle?', are: (1) shuttle orbiter thermal protection certification; (2) shuttle thermal protection system flight experience; and (3) space assembled entry system certification.

Curry, Donald M.↗

NASA NTRS1993

Prediction of rigid silica based insulation conductivity

A method is presented for predicting the thermal conductivity of low density, silica based fibrous insulators. It is shown that the method can be used to extend data values to the upper material temperature limits from those obtained from the test data. It is demonstrated that once the conductivity is accurately determined by the analytical model the conductivity for other atmospheres can be predicted. The method is similar to that presented by previous investigators, but differs significantly in the contribution due to gas and internal radiation.

Williams, Stanley D.↗

NASA NTRS1992

Thermal protection materials: Thermophysical property data

This publication presents a thermophysical property survey on materials that could potentially be used for future spacecraft thermal protection systems (TPS). This includes data that was reported in the 1960's as well as more current information reported through the 1980's. An attempt was made to cite the manufacturers as well as the data source in the bibliography. This volume represents an attempt to provide in a single source a complete set of thermophysical data on a large variety of materials used in spacecraft TPS analysis. The property data is divided into two categories: ablative and reusable. The ablative materials have been compiled into twelve categories that are descriptive of the material composition. An attempt was made to define the Arrhenius equation for each material although this data may not be available for some materials. In a similar manner, char data may not be available for some of the ablative materials. The reusable materials have been divided into three basic categories: thermal protection materials (such as insulators), adhesives, and structural materials.

Williams, S. D.↗

NASA NTRS1992

Thermal protection systems manned spacecraft flight experience

Since the first U.S. manned entry, Mercury (May 5, 1961), seventy-five manned entries have been made resulting in significant progress in the understanding and development of Thermal Protection Systems (TPS) for manned rated spacecraft. The TPS materials and systems installed on these spacecraft are compared. The first three vehicles (Mercury, Gemini, Apollo) used ablative (single-use) systems while the Space Shuttle Orbiter TPS is a multimission system. A TPS figure of merit, unit weight lb/sq ft, illustrates the advances in TPS material performance from Mercury (10.2 lb/sq ft) to the Space Shuttle (1.7 lb/sq ft). Significant advances have been made in the design, fabrication, and certification of TPS on manned entry vehicles (Mercury through Shuttle Orbiter). Shuttle experience has identified some key design and operational issues. State-of-the-art ceramic insulation materials developed in the 1970's for the Space Shuttle Orbiter have been used in the initial designs of aerobrakes. This TPS material experience has identified the need to develop a technology base from which a new class of higher temperature materials will emerge for advanced space transportation vehicles.

Curry, Donald M.↗

NASA NTRS1990

Thermal structural analysis of the RCC chin panel

This paper presents the design constraints, flight environments, conceptual design iterations and analytical methods used in the design and flight certification of the Reinforced Carbon-Carbon (RCC) chin panel for the Space Shuttle Orbiter. A discussion of the thermal/structural response of the chin panel to the imposed environments is presented. As with all critical airframe components, it was desirable to conduct a structural test to verify design adequacy. However, with the chin panel, a meaningful test to simulate the critical thermal stresses was not considered feasible.

Bohlmann, Frank J.↗

NASA NTRS1988

Prediction of aerodynamic heating and pressures on Shuttle Entry Air Data System (SEADS) nose cap and comparison with STS-61C flight data

Results are presented from predictions of aerothermodynamic heating rates, temperatures, and pressures on the surface of the Shuttle Entry Air Data System (SEADS) nosecap during Orbiter reentry. These results are compared with data obtained by the first actual flight of the SEADS system aboard STS-61C. The data also used to predict heating rates and surface temperatures for a hypothetical Transatlantic Abort Landing entry trajectory, whose analysis involved ascertaining the increases in heating rate as the airstream flowed across regions of the lower surface catalycity carbon/carbon composite to the higher surface catalycity columbium pressure ports.

Ting, Paul C.↗