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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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At least 1,261 records · Page 70

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NASA NTRS1983

A comparison with theory of peak to peak sound level for a model helicopter rotor generating blade slap at low tip speeds

Mini-tuft and smoke flow visualization techniques have been developed for the investigation of model helicopter rotor blade vortex interaction noise at low tip speeds. These techniques allow the parameters required for calculation of the blade vortex interaction noise using the Widnall/Wolf model to be determined. The measured acoustics are compared with the predicted acoustics for each test condition. Under the conditions tested it is determined that the dominating acoustic pulse results from the interaction of the blade with a vortex 1-1/4 revolutions old at an interaction angle of less than 8 deg. The Widnall/Wolf model predicts the peak sound pressure level within 3 dB for blade vortex separation distances greater than 1 semichord, but it generally over predicts the peak S.P.L. by over 10 dB for blade vortex separation distances of less than 1/4 semichord.

Fontana, R. R.↗

NASA NTRS1984

The influence of leading-edge load alleviation on supersonic wing design

A theoretical and experimental program to assess the effect of leading-edge load constraints on wing design and performance was conducted. For a planform characterized by a highly swept leading edge on the inboard region, linear theory was used to design camber surfaces which produced minimum drag-due-to-lift at the design lift coefficient of 0.08 and a design Mach number of 2.4. In an effort to delay the formation of leading edge vortices which often occur on highly swept wings, two approaches were used in the design criteria to limit the loadings on the leading edge. One wing was constrained to have the normal Mach number less than one everywhere along the leading edge and the second wing was constrained to have a pressure coefficient of zero on the leading edge. Force tests were run on the two constrained wings, on a flat reference wing and on an optimized wing with no leading edge constraints. All wings had identical planforms and thicknesses and were tested over a range of Mach numbers from 1.8 to 2.8 and a range in angles of attack from -5 deg to 8 deg. A comparison of the experimental performance of these four models is shown. Correlations of these results with theoretical predictions and flow visualization photographs are also included.

Darden, C. M.↗

NASA NTRS1984

Real-gas simulation for the Shuttle Orbiter and planetary entry configurations including flight results

By testing configurations in a gas (like CF4) which can produce high normal-shock density ratios, such as those encountered during hypersonic entry, certain aspects of real-gas effects can be simulated. Results from force-moment, shock-shape and oil flow visualization tests are presented for both the Shuttle Orbiter and a 45 deg sphere-cone in CF4 and air at M = 6, and comparisons are made with flight results. Pitching-moment coefficients measured on a Shuttle Orbiter model in CF4 showed a nose-up increment, compared with air results, that was almost identical to the difference between preflight predictions and flight in the high hypersonic regime. The drag coefficient measured in CF4 on the 45 deg sphere-cone, which is the same configuration used on the forebody of the Pioneer Venus entry vehicles, showed excellent agreement with flight data at M = 6.

Calloway, R. L.↗

NASA NTRS1984

Flow structure within a heated rectangular jet ejector

Preliminary results of an experimental study of the flow field structure within a rectangular jet ejector are presented. The tests determined the effect of primary jet heating and ejector duct width on the flow field at a fixed jet exit Mach number of 0.3. For comparison the rectangular jet without the presence of the ejector shroud was also studied. Measurements of the mean velocity field, total temperature field, and ejector shroud surface pressure distribution, along with schlieren flow visualization, were made. It was found that: the ejector did not enhance centerline temperature decay in comparison to the free jet; the presence of the ejector shroud significantly reduced jet growth; and, the ejector centerline velocity/temperature decay was not sensitive to duct width variations.

Mclachlan, B. G.↗

NASA NTRS1983

Acoustic power dissipation on radiation through duct terminations - Experiments

This paper describes the acoustic transmission characteristics of ducts, nozzles, orifices, and perforated plates, studied under an experimental program using an acoustic impulse technique. In this technique high intensity pulses, generated by discharging a capacitor across a spark gap, were used as the sound source. The test conditions include heated and unheated flows, with and without simulated flight. Results for a straight round duct, three convergent nozzles, a suppressor nozzle, 12 orifice plates, and 10 perforated plates are presented. A low frequency acoustic power loss phenomenon was observed for all configurations at all test conditions including the no flow condition. It was suspected that the power loss phenomenon at the no flow condition could be due to the conversion of acoustic energy into vortical energy due to non-linear propagation of high intensity pulses. However, a small amount of low frequency power loss was noticed even when tests were repeated with a low intensity sound. Detailed flow visualization results were also obtained to complement the acoustic results.

Salikuddin, M.↗

NASA NTRS1985

Power Supply For 25-Watt Arc Lamp

Dual-voltage circuitry both strikes and maintains arc. New power supply designed (and several units already in use) that replaces relay/choke combination with solid-state starter. New power supply consists of two main sections. First section (low voltage power supply section) is 84-volt directcurrent supply. Second section (high-voltage starter circuit) is CockroftWalton voltage multiplier. Used as light sources for schlieren, shadowgraph, and other flow-visualization techniques.

Leighty, B. D.↗

NASA NTRS1984

Experimental investigation of shock-cell noise reduction for dual-stream nozzles in simulated flight

Six scale-model nozzles were tested in an anechoic facility to evauate the effectiveness of convergent-divergent (C-D) terminations in reducing shock-cell noise of unsuppressed and mechanically suppressed coannular plug nozzles. One hundred fifty-three acoustic test points with inverted velocity profiles were conducted under static and simulated flight conditions. Diagnostic flow visualization with a shadowgraph and velocity measurements with a laser velocimeter were performed on selected plumes. Shock-cells were identified on the plug and downstream of the plug of the unsuppressed convergent coannular nozzle with truncated plug. Broadband peak frequencies predicted with the two shock-cell structures were correlated with the observed spectra using the measured shock-cell spacings. Relative to a convergent circular nozzle, the perceived noise level (PNL) data at an observer angle of 60 deg relative to inlet, indicated a reduction of (1) 6.5 dB and 9.2 dB with unsuppressed C-D coannular nozzle with truncated plug and (2) 7.7 dB and 8.3 dB with suppressed C-D coannular nozzle under static and simulated flight conditions, espectively. The unsuppressed C-D coannular nozzle with truncated plug, operating at the C-D design condition, had shock-cells downstream of the plug with no shock-cells on the plug. The downstream shock-cells were eliminated by replacing the truncated plug with a smooth extension to obtain an additional 2.4 dB and 3 dB front quadrant PNL reduction, under static and simulated flight conditions, respectively. Other results are discussed.

Janardan, B. A.↗

NASA NTRS1984

Experimental Investigation of Shock-Cell Noise Reduction for Single Stream Nozzles in Simulated Flight

Seven single stream model nozzles were tested in the Anechoic Free-Jet Acoustic Test Facility to evaluate the effectiveness of convergent divergent (C-D) flowpaths in the reduction of shock-cell noise under both static and mulated flight conditions. The test nozzles included a baseline convergent circular nozzle, a C-D circular nozzle, a convergent annular plug nozzle, a C-D annular plug nozzle, a convergent multi-element suppressor plug nozzle, and a C-D multi-element suppressor plug nozzle. Diagnostic flow visualization with a shadowgraph and aerodynamic plume measurements with a laser velocimeter were performed with the test nozzles. A theory of shock-cell noise for annular plug nozzles with shock-cells in the vicinity of the plug was developed. The benefit of these C-D nozzles was observed over a broad range of pressure ratiosin the vicinity of their design conditions. At the C-D design condition, the C-D annual nozzle was found to be free of shock-cells on the plug.

Yamamoto, K.↗

NASA NTRS1984

Solar Pond Fluid Dynamics and Heat Transfer

The primary objective of the solar pond research was to obtain an indepth understanding of solar pond fluid dynamics and heat transfer. The key product was the development of a validated one-dimensional computer model with the capability to accurately predict time-dependent solar pond temperature, salinities, and interface motions. Laboratory scale flow visualization experiments were conducted to better understand layer motion. Two laboratory small-scale ponds and a large-scale outdoor solar pond were designed and built to provide quantitative data. This data provided a basis for validating the model and enhancing the understanding of pond dynamic behavior.

Jones, G. F.↗

NASA NTRS1985

Unsteady pressure measurements on a biconvex airfoil in a transonic oscillating cascade

Flush-mounted dynamic pressure transducers were installed on the center airfoil of a transonic oscillating cascade to measure the unsteady aerodynamic response as nine airfroils were simultaneously driven to provide 1.2 deg of pitching motion about the midchord. Initial tests were performed at an incidence and angle of 0 deg and A Mach number of 0.65 in order to obtain results in a shock-free compressible flowfield. Subsequent tests were performed at an incidence angle of 7 deg and Mach number of 0.8 in order to observe the surface pressures with an oscillating shock near the leading edge of the airfoil. Results are presented for interblade phase angles of 90 and -90 deg and at blade oscillatory frequencies of 200 and 500 Hz (semi-chord reduced frequencies up to about 0.5 at a Mach number of 0.8). Results from the zero-incidence cascade are compared with a classical unsteady flat-plate analysis. Flow visualization results depicting the shock motion on the airfoils in the high-incidence cascade are discussed. The airfoil pressure data are tabulated.

Shaw, L. M.↗

NASA NTRS1985

Investigation of flowfields found in typical combustor geometries

Activities undertaken during the entire course of research are summarized. Studies were concerned with experimental and theoretical research on 2-D axisymmetric geometries under low speed nonreacting, turbulent, swirling flow conditions typical of gas turbine and ramjet combustion chambers. They included recirculation zone characterization, time-mean and turbulence simulation in swirling recirculating flow, sudden and gradual expansion flowfields, and furher complexities and parameter influences. The study included the investigation of: a complete range of swirl strengths; swirler performance; downstream contraction nozzle sizes and locations; expansion ratios; and inlet side-wall angles. Their individual and combined effects on the test section flowfield were observed, measured and characterized. Experimental methods included flow visualization (with smoke and neutrally-buoyant helium-filled soap bubbles), five-hole pitot probe time-mean velocity field measurements, and single-, double-, and triple-wire hot-wire anemometry measurements of time-mean velocities, normal and shear Reynolds sresses. Computational methods included development of the STARPIC code from the primitive-variable TEACH computer code, and its use in flowfield prediction and turbulence model development.

Lilley, D. G.↗

NASA NTRS1985

Continuous laminar smoke generator

A smoke generator capable of emitting a very thin, laminar stream of smoke for use in high detail flow visualization was invented. The generator is capable of emitting a larger but less stable rope of smoke. The invention consists of a pressure supply and fluid supply which supply smoke generating fluid to feed. The feed tube is directly heated by electrical resistance from current supplied by power supply and regulated by a constant temperature controller. A smoke exit hole is drilled in the wall of feed tube. Because feed tube is heated both before and past exit hole, no condensation of smoke generating occurs at the smoke exit hole, enabling the production of a very stable smoke filament. The generator is small in size which avoids wind turbulence in front of the test model.

Weinstein, L. M.↗

NASA NTRS1985

Convective storm downdraft outflows detected by NASA/MSFC's Airborne 10.6 micron pulsed Doppler Lidar System

The capability of a unique Airborne Doppler Lidar System to measure the horizontal winds in the vicinity of severe storm activity is demonstrated. The Airborne Doppler Lidar System (ADLS), developed at NASA/MSFC, was flown on a CV990 research aircraft during the CCOPE (Cooperative Convective Precipitation Experiments, Montana, Summer 1981). Flown between 400 and 600 m AGL, the lidar probed the subcloud regions of several deep convection storms. ADLS data collected near storms on 21 and 23 July 1981 are presented along with satellite imagery, radar echo maps and surface station measurements. These case studies are evidence of the successful performance of an airborne remote wind sensing system and the advantages of two dimension flow visualization of storm outflow structures and interactions.

Emmitt, G. D.↗

NASA NTRS1985

Preliminary results of a study of the relationship between free-stream turbulence and stagnation region heat transfer

Preliminary results of a study to investigate the relationship between free stream turbulence and heat transfer augmentation in the stagnation region is presented. The effects of free stream turbulence and surface roughness on spanwise averaged heat transfer were investigated. Turbulence was measured upstream of a cylinder placed in the wake of an array of parallel wires that were perpendicular to the cylinder axis. Finally, flow visualization and thermal visualization techniques were combined to show the relationship between vortices in the stagnation region and spanwise variations in heat transfer.

Vanfossen, G. J., Jr.↗

NASA NTRS1985

Local heat-transfer measurements on a large, scale-model turbine blade airfoil using a composite of a heater element and liquid crystals

Local heat transfer coefficients were experimentally mapped along the midchord of a five-time-size turbine blade airfoil in a static cascade operated at room temperature over a range of Reynolds numbers. The test surface consisted of a composite of commercially available materials: a mylar sheet with a layer of cholesteric liquid crystals, that change color with temperature, and a heater sheet made of a carbon-impregnated paper, that produces uniform heat flux. After the initial selection and calibration of the composite sheet, accurate, quantitative, and continuous heat transfer coefficients were mapped over the airfoil surface. The local heat transfer coefficients are presented for Reynolds numbers from 2.8 x 10 to the 5th power to 7.6 x 10 to the 5th power. Comparisons are made with analytical values of heat transfer coefficients obtained from the STAN5 boundary layer code. Also, a leading edge separation bubble was revealed by thermal and flow visualization.

Hippensteele, S. A.↗

NASA NTRS1985

Circulation control lift generation experiment: Hardware development

A circulation control airfoil and its accompanying hardware were developed to allow the investigation of lift generation that is independent of airfoil angle of attack and relative flow velocity. The test equipment, designed for use in a water tunnel, includes the blown airfoil, the support systems for both flow visualization and airfoil load measurement, and the fluid control system, which utilizes hydraulic technology. The primary design tasks, the selected solutions, and the unforseen problems involved in the development of these individual components of hardware are described.

Panontin, T. L.↗

NASA NTRS1985

A three-dimensional turbulent boundary layer undergoing transverse strain and streamwise pressure gradient

Results from an experimental investigation designed to provide data on both mean and turbulence quantities in the axisymmetric, swirling boundary layer (with and without pressure gradient) flowing over a stationary cylinder downstreams of a spinning cylindrical section are presented. The pressure gradient was introduced into the flow field by a 25.4 mm-high, forward-facing, circular step mounted on the stationary cylinder, the step height being nearly equal to the thickness of the approaching boundary layer. All the measurements were made at a nominal upstream reference Reynolds number of 2.4 x 10 to the 6th power/m (corresponding to an upstream reference velocity of 36 to 37 m/sec) with the rotation of the spinner set to make its peripheral speed equal the reference velocity. The data reported included measurements of surface pressure and the mean surface shear-stress vector taken with a miniature, directional, surface-fence gage. These measurements were supplemented by oil-flow visualization studies of the stationary cylinder. The data indicates that the streamwise pressure gradient controls the development of the streamwise component of wall shear, but leaves the peripheral component of wall shear practically unaffected.

Hebbar, S. K.↗

NASA NTRS1985

Some observations on the occurrence of striation heating

Various wind tunnel flow visualization techniques have revealed the presence of striations on the surface of both simple geometries and complex winged entry configurations over a wide range of test conditions. The striations are attributed to streamwise vortices embedded in the boundary layer which produce locally enhanced heating levels. Mechanisms for such vortex formation are suggested based on observation of their occurrence on the simple configurations. These mechanisms are then related to striations observed in the wind tunnel on windward surfaces, swept wing leading edges and on the side fuselage of winged entry bodies. Flight data obtained on the side fuselage of ASSET and Shuttle support the conclusions derived from wind tunnel results. Quantitative criteria for the appearance of striations resulting from some of these mechanisms are presented. It is shown that the presence of embedded vorticity may be a consideration in the design of thermal protection systems for advanced winged entry vehicles.

Kipp, H. W.↗