Engineering topics

Richie, C. B.

Publications and source records attributed to Richie, C. B..

5 records

NASA NTRS1982

Hypersonic airbreathing missile concepts under study at NASA Langley Research Center

The design and performance of several tactical and strategic hypersonic airbreathing missile concepts under study at the NASA Langley Research Center are discussed from an evolutionary perspective. A mid- and chin inlet missile design, constrained to the Navy's vertical box launcher, was investigated; a performance comparison is presented that is favorable to the mid-inlet approach. Parasol wing, confined flow field, and spatula-like cruise missile configurations were examined with strategic applications in mind. The preliminary results are encouraging with respect to aerodynamic and volumetric efficiency and choice of engine integration schemes.

Hunt, J. L.↗

NASA NTRS1977

Aerodynamic characteristics at Mach 6 of a hypersonic research airplane concept having a 70 deg swept delta wing

The hypersonic aerodynamic characteristics of an air-launched, delta-wing research aircraft concept were investigated at Mach 6. The effect of various components such as nose shape, wing camber, wing location, center vertical tail, wing tip fins, forward delta wing, engine nacelle, and speed brakes was also studied. Tests were conducted with a 0.021 scale model at a Reynolds number, based on model length, of 10.5 million and over an angel of attack range from -4 deg to 20 deg. Results show that most configurations with a center vertical tail have static longitudinal stability at trim, static directional stability at angles of attack up to 12 deg, and static lateral stability throughout the angle of attack range. Configurations with wing tip fins generally have static longitudinal stability at trim, have lateral stability at angles of attack above 8 deg, and are directionally unstable over the angle of attack range.

Clark, L. E.↗

NASA NTRS1975

Fail-safe system for activity cooled supersonic and hypersonic aircraft

A fail-safe-system concept was studied as an alternative to a redundant active cooling system for supersonic and hypersonic aircraft which use the heat sink of liquid-hydrogen fuel for cooling the aircraft structure. This concept consists of an abort maneuver by the aircraft and a passive thermal protection system (TPS) for the aircraft skin. The abort manuever provides a low-heat-load descent from normal cruise speed to a lower speed at which cooling is unnecessary, and the passive TPS allows the aircraft skin to absorb the abort heat load without exceeding critical skin temperature. On the basis of results obtained, it appears that this fail-safe-system concept warrants further consideration, inasmuch as a fail-safe system could possibly replace a redundant active cooling system with no increase in weight and would offer other potential advantages.

Jones, R. A.↗

NASA NTRS1974

Effect of wall temperature spatial variations on aerodynamic heating to an actively cooled aircraft

The design of cooled skin panels, which requires consideration of many parameters such as heat load, skin thickness, panel size, tube spacing, tube size, temperature limits, coolant flow, etc., must minimize weight and thermal stresses. A design procedure for such panels has been developed by Bell Aerospace Company (Helenbrook and Anthony, 1971). This design procedure assumes a uniform heating rate to the panel, i.e., the variation of aerodynamic heating rate due to the cyclic wall temperature distribution, which results from the discrete coolant tube locations in such a panel, has been neglected. This study was made to determine if these variations were sufficiently large to cause variations in aerodynamic heating rate which would require coupling of boundary layer and conductive thermal analysis solutions.

Jones, R. A.↗

NASA NTRS1973

Application of phase-change technique to thin sections with heating on both surfaces

A numerical digital computer program was developed to calculate the heat-transfer coefficients for both sides of a one-dimensional finite slab subject to the boundary conditions ascribed to the phase-change coating heat-transfer technique. In a typical tunnel test situation where a thin wing was exposed to heating on both sides, the data reduction procedures for a semi-infinite slab gave heat-transfer coefficients as much as 375 percent too high on the side with the lowest heating. The results from the one-dimensional finite-slab procedure are presented in the form of correction factors to the solution for a semi-infinite slab in terms of parameters normally used with the phase-change heat-transfer technique. These correlations are not restricted to slab thickness or thermophysical properties and are easily used to obtain accurate data on thin model sections.

Hunt, J. L.↗