Engineering topics

Rose, W. C.

Publications and source records attributed to Rose, W. C..

At least 19 records

NASA NTRS1989

Image quality on the Kuiper Airborne Observatory. I - Results of the first flight series

The NASA Kuiper Airborne Observatory (KAO) was flown three times during June and July, 1984 in order to study the causes of the poor seeing obtained with the 0.9-m telescope. High-speed pressure and temperature sensors were placed in the telescope cavity. Several thousand stellar images were recorded under various flight and optical configurations. It is found that the long-exposure image size is affected by telescope tracking errors, imperfect optics, poor optical alignment, telescope and instrument vibration, thermal fluctuations in the telescope cavity, and density fluctuations in the shear layer that forms the boundary between the cavity air and outside air. Possible ways to improve the quality of the images are discussed.

Elliot, J. L.↗

NASA NTRS1986

Aerodynamics of seeing on large transport aircraft

Data were obtained in the full scale flight environment of the Kuiper Airborne Observatory (KAO) on the nature of turbulent shear layer over the open cavity. These data were used to verify proposed aerodynamic scaling relationships to describe the behavior of the turbulent layers and to estimate the optical performance of systems of various wavelengths operating within the KAO environment. These data and wind tunnel data are used to scale the expected optical effects for a potential stratospheric observatory for infrared astronomy (SOFIA) in which a telescope approximately 3.5 times larger than that on the KAO is envisioned. It appears that the use of combinations of active and passive aeromechanical flow control techniques can improve the optical behavior of systems in the SOFIA environment. Experiments to verify these potential improvements can be performed on the KAO with sufficient modifications to the cavity and aero-mechanical technique installations.

Rose, W. C.↗

NASA NTRS1985

Computations and turbulent flow modeling in support of helicopter rotor technology

The angle of attack (AOA) tandem cavity wind tunnel experiment was prepared. Actual wind tunnel testing started shortly after the beginning of 1985. A multi-probe aerodynamic rake was designed and installed for use in surveying the shear layers present over the open cavity on the Kuiper Airborne Observatory (KAO). The nature of the behavior of the thermal environment within the KAO cavity at operational altitudes was determined. Assistance was given in the design of the cavity for the University of Denver radiometer. Attempts to distinguish between the optical terms of blur circle size (or image size) and the term due to jitter were discussed.

Rose, W. C.↗

NASA NTRS1982

Computations and turbulent flow modeling in support of helicopter rotor technology

Efforts continue to investigate the applicability of the Deiwert's time dependent numerical airfoil calculation code to the simulation of two dimensional airfoil flows with large amounts of separation present. A modified logarithmic region mixing length term produced substantial differences in the calculated flow fields for the large separation of the M=0.4, high angle of attack cases. Reasons are given for interest in applying that model to previously investigated cases of the 64A010 airfoil section at Mach number 0.8 for angles of attack of 4 deg and 6.2 deg. The effects of these turbulence models at alterations and the time dependent behavior from impulsively started conditions are discussed.

Rose, W. C.↗

NASA NTRS1982

The effect of ejector augmentation on test-section flow quality in the Calspan 8-ft transonic wind tunnel

Tests to determine the flow disturbances effects of four ejectors located in the corners just downstream of the diffuser in the Calspan wind tunnel are described. The flow quality in the nonejector mode is employed as the base-line configuration, and operating parameters are compared with data from other wind tunnels. During tests with the ejectors working, fluctuation levels increased between Mach 0.4-0.6, while temperature and vorticity levels remained constant. The ejector exhibited broad spectrum noise typical of free jet noise, yet static pressure measurements revealed only a slight increase in the broadband rms levels with the ejectors on, indicating negligible disturbances upstream caused by the ejectors. Choking the diffuser eliminated the jet noise, and the use of ejectors in the Mach range considered is concluded to cause no significant degradation in the Calspan tunnel flow quality.

Rose, W. C.↗

NASA NTRS1980

Rugged fast-response temperature probe

Very-sensitive probe uses tungsten sensor wire wrapped around to notched electrodes. Design combines ruggedness of earlier but less sensitive probe using very-short sensor wire with very-sensitive but fragile version with sensor wires wrapped around non-conducting frames.

Bailey, P. L.↗

NASA NTRS1980

Unsteady density and velocity measurements in the 6 foot x 6 foot wind tunnel

The methods used and the results obtained in four aero-optic tests are summarized. It is concluded that the rather large values of density fluctuation appear to be the result of much higher Mach number than freestream and the violent turbulence in the flow as it separates from the turret. A representative comparison of fairing on-fairing off rms density fluctuation indicates essentially no effect at M = 0.62 and a small effect at M = 0.95. These data indicate that some slight improvement in optical quality can be expected with the addition of a fairing, although at M = 0.62 its effect would be nil. Fairings are very useful in controlling pressure loads on turrets, but will not have first order effects on optical quality. Scale sizes increase dramatically with increasing azimuth angle for a reprensentative condition. Since both scale sizes and fluctuation levels increase (total turbulence path length also increases) with azimuth angle, substantial optical degradation might be expected. For shorter wave lengths, large degradations occur.

Rose, W. C.↗

NASA NTRS1980

Summary of all cycle II.5 shear and boundary layer measurements, aerodynamics

The two measurement systems were used to measure mean velocity and velocity, mass flux, and total temperature fluctuations in the turbulent boundary on the fuselage of a KC-135 aircraft. The boundary layer thickness ranged between about 20 and 30 cm for the range of flight Mach numbers from about 0.25 to 0.85 and Reynolds numbers between 3 and 6 x 10 to the 6th power/m. The adaptation of each system for use in airborne applications is discussed. The data obtained from each system are given and compared with each other and they indicate that the two systems represent viable ones for use in future airborne turbulence experiments.

Rose, W. C.↗

NASA NTRS1980

Airborne measurement of atmospheric turbulence

A system capable of making measurements of fluctuating atmospheric density is described. Spatial scales required in assessing the quality of coherent radiation propagation are discussed. The special sensors, aircraft installation, data reduction procedures, and other special requirements necessary to obtain meaningful atmospheric turbulence data are also described. The spectral distribution of density fluctuation are presented.

Rose, W. C.↗

NASA NTRS1978

Buried wire gage for wall shear stress measurements

A buried wire gage for measuring wall shear stress in fluid flow was studied and further developed. Several methods of making this relatively new type of gage were examined to arrive at a successful technique that is well-suited for wind-tunnel testing. A series of measurements was made to demonstrate the adequacy of a two-point calibration procedure for these gages. The buried wire gage is also demonstrated to be ideally suited for quantitative measurement of wall shear stress in wind-tunnel testing.

Murthy, V. S.↗

NASA NTRS1977

Form drag, skin friction, and vortex shedding frequencies for subsonic and transonic crossflows on circular cylinder

A series of wind-tunnel tests covering a range of Mach numbers and Reynolds numbers in subsonic and transonic flows was conducted on a circular cylinder placed normal to the flow. Form drag coefficients were determined from surface-pressure measurements and displayed as a function of Mach number to show the drag rise phenomenon. Buried wire gages arranged on the model surface were used to measure skin-friction distributions and vortex-shedding frequencies at different flow conditions. It was found that detectable periodic shedding ceases above M = 0.9. The measured skin-friction distributions indicate the positions of mean separation points clearly; these values are documented for the different flow conditions.

Murthy, V. S.↗

NASA NTRS1977

Direct measurements of wall shear stress by buried wire gages in a shock-wave boundary-layer interaction region

Detailed measurements of wall shear stress (skin friction) were made with specially developed buried wire gages in the interaction regions of a Mach 2.9 turbulent boundary layer with externally generated shocks. Separation and reattachment points inferred by these measurements support the findings of earlier experiments which used a surface oil flow technique and pitot profile measurements. The measurements further indicate that the boundary layer tends to attain significantly higher skin-friction values downstream of the interaction region as compared to upstream. Comparisons between measured wall shear stress and published results of some theoretical calculation schemes show that the general, but not detailed, behavior is predicted well by such schemes.

Murthy, V. S.↗

NASA NTRS1977

Calculation of transonic flow over supercritical airfoil sections

Two numerical methods for calculating the transonic flow, including viscous effects over lifting airfoil sections and experimental data are compared for turbulent flow over a supercritical airfoil. In addition, results for a NACA 64A010 airfoil at nonzero angle of attack are compared to demonstrate the applicability of the numerical methods to classical, lifting airfoils. One numerical method is a solution to the time-averaged Navier-Stokes equations throughout the entire flow field. The other is a hybrid method that combines inviscid, boundary-layer, and Navier-Stokes equations in appropriate regions of the flow field. Both methods adequately predict the surface pressures and flow field about the 64A010 airfoil at M = 0.8 and alpha = 2 deg when an appropriate turbulence model is used. The methods are not as successful for the supercritical airfoil.

Rose, W. C.↗

NASA NTRS1977

Hot-wire anemometry in transonic flow

The use of hot-wire anemometry for obtaining fluctuating data in transonic flows has been evaluated. From hot-wire heat loss correlations based on previous transonic data, the sensitivity coefficients for velocity, density, and total temperature fluctuations have been calculated for a wide range of test conditions and sensor parameters. For sensor Reynolds number greater than 20 and high sensor overheat ratios, the velocity sensitivity remains independent of Mach number and equal to the density sensitivity. These conditions were verified by comparisons of predicted sensitivities with those from recent direct calibrations in transonic flows. Based on these results, techniques are presented to obtain meaningful measurements of fluctuating velocity, density, and Reynolds shear stress using hot-wire and hot-film anemometers. Example of these measurements are presented for two transonic boundary layers.

Horstman, C. C.↗

NASA NTRS1977

An approximate calculation of the strong interaction on a transonic airfoil

This paper describes an interactive computational scheme that calculates the transonic flow field over an airfoil. The inviscid flow, the turbulent boundary layer in the weak interaction region, and the elliptic flow in the strong interaction region are computed separately, and the solutions are matched iteratively in a direct, second-order interaction scheme. The pressure distribution, obtained from a transonic, small-perturbation solution, is prescribed on the boundary layer and the Navier-Stokes equations and the ensuing displacement thicknesses are used to modify the inviscid flow. The best overall results are obtained by using a turbulence model that eliminates the eddy viscosity in the strong interaction region. Computed pressure distributions are compared with experimental data and with Navier-Stokes solutions. In weak interactions, the agreement with experimental data is very good, as it is also in strong interactions except for the pressures in the separated-flow region. There the agreement is qualitative only, and the pressure recovery is too high. The results in both cases are superior or equal to those of the Navier-Stokes solutions, and the present method is significantly faster.

Seginer, A.↗

NASA NTRS1976

A numerical solution of the flow field over a transonic airfoil including strong-shock-induced flow separation

An approximate numerical method has been developed for computing the entire transonic flow field over an airfoil, including the boundary layer, the shock-wave/boundary-layer interaction, and the shock-induced-separation. The inviscid flow, the turbulent boundary layer in the weak interaction zone, and the strong interaction elliptic region are solved separately. The inviscid flow is represented by the solution of the transonic small-disturbance equation, while the boundary layer is solved by the generalized Galerkin method, and the strong interaction region (including the wake) is computed from the time-averaged Navier-Stokes equations. A second-order interaction is computed iteratively using the inviscid pressure distribution as a boundary condition for the boundary layer and Navier-Stokes codes, and the viscous displacement thickness to modify the airfoil geometry for the inviscid computation. The computed pressure distributions and shock-wave profiles are compared with experimental data. Agreement is very good except that the pressure recovery on the trailing edge is greater than the experimental data.

Seginer, A.↗

NASA NTRS1976

Turbulence measurements in a transonic boundary layer and free-shear flow using laser velocimetry and hot-wire anemometry techniques

Quantitative measurements of the turbulence fluctuations in velocity and mass flux have been obtained in Mach 0.6 and 0.8 turbulent boundary layer and free-shear layer flows by laser velocimetry and hot-wire anemometry techniques. To evaluate the effects of compressibility, these transonic data are compared to available incompressible and supersonic results. Based on some simplifying assumptions, estimates of the rms density fluctuations are made for which error bounds are given. In addition to these fluctuation data, the compressible mean velocity data obtained with the laser velocimeter are presented and compared to pitot tube results. The investigation was conducted in the Ames 6- by 6-Foot Supersonic Wind Tunnel at free-stream Mach numbers of 0.6 and 0.8 for a unit Reynolds number of about 10,000,000 per meter.

Johnson, D. A.↗

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.↗