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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 37 records · Page 2
All sourcesNASA NTRSDOE OSTI
NASA NTRS1986

High Reynolds number tests of a Douglas DLBA 032 airfoil in the Langley 0.3-meter transonic cryogenic tunnel

A wind-tunnel investigation of a Douglas advanced-technology airfoil was conducted in the Langley 0.3-Meter Transonic Cryogenic Tunnel (0.3-m TCT). The temperature was varied from 227 K (409 R) to 100 K (180 R) at pressures ranging from about 159 kPa (1.57 atm) to about 514 kPa (5.07 atm). Mach number was varied from 0.50 to 0.78. These variables provided a Reynolds number range (based on airfoil chord) from 6.0 to 30.0 x 10 to the 6th power. This investigation was specifically designed to: (1) test a Douglas airfoil from moderately low to flight-equivalent Reynolds numbers, and (2) evaluate sidewall-boundary-layer effects on transonic airfoil performance characteristics by a systematic variation of Mach number, Reynolds number, and sidewall-boundary-layer removal. Data are included which demonstrate the effects of fixing transition, Mach number, Reynolds number, and sidewall-boundary-layer removal on the aerodynamic characteristics of the airfoil. Also included are remarks on model design and model structural integrity.

Johnson, Charles B.↗
NASA NTRS1961

Investigation of an Axisymmetric Internal Compression Inlet at a Mach Number of About 3.8

An idealized axisymmetric, all-internal compression inlet was designed for a Mach number of 3.75. The objective of the design was to obtain a steady, one-dimensional transonic flow and a high over-all total pressure recovery. Boundary-layer removal was employed in the vicinity of the inflection point of the supersonic contour. Static and total-pressure fluctuations were measured in the transonic flow region. A total pressure recovery of about 90 percent was obtained with a boundary-layer-removal mass flow rate of 15 percent of the inlet mass flow rate. The accompanying root-mean-square total pressure fluctuation in the throat region was only 1 percent of the free stream total pressure. The test Mach number was 3.80 and the Reynolds number based on inlet diameter was 2.63 x 10(exp 6).

Lundell, John H.↗
NASA NTRS1952

Performance Characteristics of a Normal-shock Side Inlet Located Downstream of a Canard Control Surface at Mach Numbers of 1.5 and 1.8/

The performance characteristics of a downward canted normal-shock side (scoop) inlet located downstream of a triangular control surface are presented for free-stream Mach numbers of 1.5 and 1.8 in terms of total pressure recovery and mass flow ratio for various boundary-layer removal systems,angles of attack, control surface deflections and adverse yaw. An engine operating condition for a hypothetical turbojet engine is established, and the match point characteristics of the engine-inlet configuration are summarized. 520::It is shown that the diffuser performance increases with increased boundary-layer removal and decreases because of the presence of the wake from the forward control surface. At the higher angles of attack the wake passes over the inlet and does not affect the inlet performance. Adverse yaw reduces the total pressure recovery values below those for the unawed case. Magnitudes of the total pressure recovery were below the theoretical normal-shock recovery for the respective test Mach numbers.

INLETS, SIDE - DUCTED BODIES↗
NASA NTRS1979

Calculation of the three-dimensional flow field in supersonic inlets at angle of attack using a bicharacteristic method with discrete shock wave fitting

An analysis is presented for calculating the flow field in supersonic mixed-compression aircraft inlets operating at angle of attack. The flow field is computed by a steady three-dimensional bicharacteristic method. The bow shock wave and the reflected internal shock wave system are computed by a three-dimensional discrete shock wave fitting procedure. Viscous and thermal diffusion may be included as source terms in the bicharacteristic method. A production type computer program capable of determining the flow field in a variety of axisymmetric mixed-compression supersonic inlets is available. The results of the present analysis agree well with those produced by the two-dimensional method of characteristics when axisymmetric flow fields are computed. For three-dimensional flow fields, the results of the present analysis agree well with experimental data except in regions of high viscous interaction and boundary layer removal. The present analysis does not compute the boundary layer, nor does it account for boundary layer bleed.

Vadyak, J.↗
NASA NTRS1976

Comparison of solutions to boundary-layer and Navier-Stokes equations for the case of mass removal from a turbulent boundary layer

Comparisons are made over a range of Mach numbers from 1.5 to 7.5 at a fixed Reynolds number for various amounts of mass removal (boundary-layer bleed). Results indicate that boundary-layer theory can be used for design purposes in the range of bleed rates required to eliminate separation in an isolated, strong, adverse pressure gradient. Only about 5% to 10% of the boundary-layer mass flow must be removed to prevent such separations. Direct comparisons of bleed rates required to eliminate separation predicted by boundary-layer and Navier-Stokes equations for an M = 3.5-engine inlet centerbody show comparable results.

Rose, W. C.↗
NASA NTRS2004

Surface-Satellite Measurements for Atmospheric Radiative Transfer (SMART)and Chemical, Optical and Microphysical Measurements of In-Situ Troposphere (COMMIT) Research Activities

The GSFC SMART consists of a suite of remote sensing instruments, including many commercially available radiometers, spectrometer, interferometer, and three in-house developed inskuments: micro-pulse lidar (MPL), scanning microwave radiometer (SMiR), and sun-sky-surface photometer (S(sup 3)). SMART cover the spectral range from UV to microwave, and passive to active remote sensing. This is to enrich the EOS-like research activities (i.e., EOS validation, innovative investigations, and long-term local environmental observations). During past years, SMART has been deployed in many NASA supported field campaigns to collocate with satellite nadir overpass for intercomparisons, and for initializing model simulations. Built on the successful experience of SMART, we are currently developing a new ground-based in-situ sampling package, COMMIT, including measurements of trace gases (CO, SO2, NOx, and O3,) concentrations, fine/coarse particle sizers and chemical composition, single- and three-wavelength nephelometers, and surface meteorological probes. COMMIT is built for seeking a relationship between surface in-situ measurements and boundary layer characteristics. This is to enrich EOS-like research on removing boundary layer signal from the entire column from space observation - to deduce the stable (less variability) free-troposphere observations. The COMMIT will try its best to link the chemical, microphysical, and optical properties of the boundary layer with radiation. The next major activities for SMART-COMMIT are scheduled for 2004-2006 in BASE-ASIA and EAST-AIRE. The expected close collaboration of BASE-ASIA with various research projects in Asia (i.e., ABC in South Asia, led by Ramanathan et al.; EAST-AIRE in East Asia, led by Li et al.; and APEX in Northeast Asia, led by Nakajima et al.) will definitely provide a better understanding of the impact of the biomass burning and air pollutants on regional-to-global climate, hydrological and carbon cycles, and tropospheric chemistry in Asia.

Tsay, Si-Chee↗
NASA NTRS1974

Experimental and theoretical investigation of a slotted noise shield model for wind tunnel walls

Tests at Mach 6 and 8 have been conducted on a model of a wind tunnel sound radiation shield which consists of interchangeable circular and wedge-shaped longitudinal rods with gaps between the rods for boundary layer removal. In view of the large effective sweep angle of the circular rods with respect to the local flow, the boundary layer behavior was surprisingly well represented by swept cylinder theory. Laminar flow was maintained up to a maximum local unit Reynolds number of 8 million per foot which would presumably be independent of model length for a given rod diameter by analogy with results on swept cylinders. Reduction in wind-tunnel sound levels within the shielded region of the circular rod model were about 90% of the maximum possible reduction for an ideal shield.

Harvey, W. D.↗
NASA NTRS1976

Fluctuating disturbances in a Mach 5 wind tunnel

An experimental investigation has been conducted to determine the source and nature of disturbances in the settling chamber and test section of a Mach 5 wind tunnel. Various changes in the air supply piping to the wind tunnel are shown to influence the disturbance levels in the settling chamber. These levels were reduced by the use of an acoustic muffler section in the settling chamber. Three nozzles were tested with the same settling chamber and hot-wire measurements indicated that the test section disturbances were entirely acoustic. Significant reductions in the test section noise levels were obtained with an electroplated nozzle utilizing boundary-layer removal upstream of the throat. The source of test section noise is shown to be different for laminar and turbulent nozzle-wall boundary layers.

Anders, J. B.↗
NASA NTRS1984

Nozzle optimization study for quiet supersonic wind tunnels

In the present study of the effects of operational factors and nozzle design parameters on the size of the quiet test section in several Mach 3.5 wind tunnel nozzles, the test results from a rapid expansion pilot nozzle with boundary layer removal slots upstream of the throat have shown that the low stream noise levels required for transition research are only obtainable with laminar nozzle wall boundary layers. Six different nozzles were evaluated, including a flat wall radial flow nozzle and a novel rod-wall nozzle. Attention is given to rms surface finish effects on quiet test region length.

Beckwith, I. E.↗
NASA NTRS1956

On the Contribution of Turbulent Boundary Layers to the Noise Inside a Fuselage

The following report deals in preliminary fashion with the transmission through a fuselage of random noise generated on the fuselage skin by a turbulent boundary layer. The concept of attenuation is abandoned and instead the problem is formulated as a sequence of two linear couplings: the turbulent boundary layer fluctuations excite the fuselage skin in lateral vibrations and the skin vibrations induce sound inside the fuselage. The techniques used are those required to determine the response of linear systems to random forcing functions of several variables. A certain degree of idealization has been resorted to. Thus the boundary layer is assumed locally homogeneous, the fuselage skin is assumed flat, unlined and free from axial loads and the 'cabin' air is bounded only by the vibrating plate so that only outgoing waves are considered. Some of the details of the statistical description have been simplified in order to reveal the basic features of the problem. The results, strictly applicable only to the limiting case of thin boundary layers, show that the sound pressure intensity is proportional to the square of the free stream density, the square of cabin air density and inversely proportional to the first power of the damping constant and to the second power of the plate density. The dependence on free stream velocity and boundary layer thickness cannot be given in general without a detailed knowledge of the characteristics of the pressure fluctuations in the boundary layer (in particular the frequency spectrum). For a flat spectrum the noise intensity depends on the fifth power of the velocity and the first power of the boundary layer thickness. This suggests that boundary layer removal is probably not an economical means for decreasing cabin noise. In general, the analysis presented here only reduces the determination of cabin noise intensity to the measurement of the effect of any one of our variables (free stream velocity, boundary layer thickness, plate thickness or the characteristic velocity of propagation in the plate). The plate generates noise by vibrating in resonance over a wide range of frequencies and increasing the damping constant is consequently an effective method of decreasing noise generation. One of the main features of the results is that the relevant quantities upon which noise intensity depends are non-dimensional numbers in which boundary layer and plate properties enter as ratios. This is taken as an indication that in testing models of structures for boundary layer noise it is not sufficient to duplicate in the model the structural characteristics of the fuselage. One must match properly the characteristics of the exicitng pressure fluctuations to that of the structure.

Corcos, G. M.↗