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Rashis, Bernard

Publications and source records attributed to Rashis, Bernard.

15 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS1961

Measurement and Empirical Correlation of Transpiration-Cooling Parameters on a 25 degree Cone in a Turbulent Boundary Layer in Both Free Flight and a Hot-Gas Jet

Transpiration-cooling parameters are presented for a turbulent boundary layer on a cone configuration with a total angle of 250 which was tested in both free flight and in an ethylene-heated high-temperature jet at a Mach number of 2.0. The flight-tested cone was flown to a maximum Mach number of 4.08 and the jet tests were conducted at stagnation temperatures ranging from 937 R to 1,850 R. In general, the experimental heat transfer was in good agreement with the theoretical values. Inclusion of the ratio of local stream temperature to wall temperature in the nondimensional flow rate parameter enabled good correlation of both sets of transpiration data. The measured pressure at the forward station coincided with the theoretical pressure over a sharp cone; however, the measured pressure increased with distance from the nose tip.

Walton, Thomas E., Jr.↗
NASA NTRS1961

Exploratory Investigation of Transpiration Cooling of a 40 Degree Double Wedge Using Nitrogen and Helium as Coolants at Stagnation Temperatures from 1,295 Degrees Fahrenheit to 2,910 Degrees Fahrenheit

An investigation of transpiration cooling has been conducted in the preflight jet of the Langley Pilotless Aircraft Research Station at Wallops Island, Va. The model consisted of a double wedge of 40 deg included angle having a porous stainless-steel specimen inserted flush with the top surface of the wedge. The tests were conducted at a free-stream Mach number of 2.0 for stagnation temperatures ranging from 1,295 F to 2,910 F. Nitrogen and helium were used as coolants and tests were conducted for values ranging from approximately 0.03 to 0.30 percent of the local weight flow rate. The data for both the nitrogen and helium coolants indicated greater cooling effectiveness than that predicted by theory and were in good agreement with the results for an 8 deg cone tested at a stagnation temperature of 600 F. The results indicate that the helium coolant, for the same amount of heat-transfer reduction, requires only about one-fourth to one-fifth the coolant flow weight as the nitrogen coolant.

Rashis, Bernard↗
NASA NTRS1961

Free-Flight Skin-Temperature and Surface-Pressure Measurements on a Highly Polished Nose Having a 100 deg Total-Angle Cone and a 10 deg Half-Angle Conical Flare Section up to a Mach Number of 4.08

The skin temperature and surface pressure were measured on a large-scale, highly polished nose having a relatively sharp-tipped 100 deg total-angle cone followed by a conical flare section of 10 deg half-angle. The measurements were obtained in flight from a rocket-propelled model up to a peak Mach number of 4.08 and a peak Reynolds number of 22 x 10(exp 6) per foot. Temperature distributions indicated that the heating on the forward 3.5 inches of the 100 deg cone was lower than the heating on the rearward portion. Likewise, measured temperatures on the flare portion of the test nose were generally lower than the temperatures on the 100 deg cone portion. The data indicated that the local Reynolds numbers of transition, based on calculated boundary-layer momentum thicknesses, ranged from 530 to 940 for a Mach number range from 2.72 to 3.75. Comparison of measured cone pressures with theory for a sharp cone showed that theory overestimates the cone pressures. Pressure measurements on the flare portion of the nose showed that in the lower speed range the flow expands below atmospheric pressure in going from the cone to the flare; however, as the speed increased, the expansion diminished and for speeds greater than a Mach number of approximately 3.0 the flare pressure coefficients were at or near a value of zero.

Rashis, Bernard↗
NASA NTRS1960

An Experimental Investigation and Correlation of the Heat Reduction to Nonporous Surfaces Behind a Porous Leading Edge Through Which Coolant is Ejected

A configuration of a wing segment having constant chord thickness, 0 deg. sweep, a porous steel semicircular leading edge, and solid Inconel surfaces was tested in a Mach number 2.0 ethlyene-heated high-temperature air jet. Measurements were made of the wing surface temperatures at chordwise stations for several rates of helium flow through the porous leading edge. The investigation was conducted at stagnation temperatures ranging from 500 F to 2,400 F, at Reynolds numbers per foot ranging from 0.3 x 10(exp 7) to 1.2 x 10(exp 7), and at angles of attack of 0, +/- 5, and +/- 15 deg. The results indicated that the reduction of wing surface temperatures with respect to their values for no coolant flow, depended on the helium coolant flow rates and the distance behind the area of injection. The results were correlated in terms of the wall cooling parameter and the coolant flow-rate parameter, where the nondimensional flow rate was referenced to the cooled area up to the downstream position. For the same coolant flow rate, lower surface temperatures are achieved with a porous-wall cooling system. However, since flow-rate requirements decrease with increasing allowable surface temperatures, the higher allowable wall temperatures of the solid wall as compared to the structurally weaker porous wall- sharply reduce the flow-rate requirements of a downstream cooling system. Thus, for certain flight conditions it is possible to compensate for the lower efficiency of the downstream or solid-wall cooling system. For example, a downstream cooling system using solid walls that must be maintained at 1,800 F would require less coolant for Mach numbers up to 5.5 than would a porous-wall cooling system for which the walls must be maintained at temperatures less than or equal to 9000 F.

Witte, William G.↗
NASA NTRS1958

Experimental Ablation Cooling

The film shows ablation tests on Teflon, nylon, a 27 percent phenolic resin, Haveg Rocketon, and graphite. Teflon hemisphere-shaped and flat face noses were tested with laboratory-scale ceramic-heated, pilot-model ceramic-heated, and electric-arc-powered air jets. Nylon hemisphere-shaped noses were tested with laboratory-scale ceramic-heated and electric-arc-powered air jets. Phenolic resin hemisphere-shaped noses were tested with laboratory-scaled ceramic-heated air jets. Haveg Rocketon and graphite hemisphere-shaped noses were tested with electric-arc-powered air jets.

Cooling↗