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At least 325 records · Page 18

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NASA NTRS1972

An investigation of the unsteady flow associated with plume induced flow separation

A wind tunnel study of the basic nature of plume induced flow separation is reported with emphasis on the unsteady aspects of the flow. Testing was conducted in a 6 inch by 6 inch blow-down supersonic wind tunnel. A cone-cylinder model with a pluming jet was used as the test model. Tests were conducted with a systematic variation in Mach number and plume pressure. Results of the tests are presented in the form of root-mean-squared surface pressure levels, power spectral densities, photographs of the flow field from which shock angles and separation lengths were taken, and time-averaged surface pressure profiles.

Boggess, A. L., Jr.↗

NASA NTRS1973

Research on the sonic boom problem. Part 1: Second-order solutions for the flow field around slender bodies in supersonic flow for sonic boom analysis

A second-order theory for supersonic flow past slender bodies is presented. Through the introduction of characteristic coordinates as independent variables and the expansion procedure proposed by Lin and Oswatitsch, a uniformly valid solution is obtained for the whole flow field in the axisymmetric case and for far field in the general three-dimensional case. For distances far from the body the theory is an extension of Whitham's first-order solution and for the domain close to the body it is a modification of Van Dyke's second-order solution in the axisymmetric case. From the theory useful formulas relating flow deflections to the Whitham F-function are derived, which permits one to determine the sonic boom strength from wind tunnel measurements fairly close to the body.

Landahl, M.↗

NASA NTRS1974

Estimation of deviation angle for axial-flow compressor blade sections using inviscid-flow solutions

Development of a method of estimating deviation angles by analytical procedures was begun. Solutions for inviscid, irrotational flow in the blade-to-blade plane were obtained with a finite-difference calculation method. Deviation angles for a plane cascade with a rounded trailing edge were estimated by using the inviscid-flow solutions and three trailing-edge hypotheses. The estimated deviation angles were compared with existing experimental data over a range of incidence angles at inlet flow angles of 30 deg and 60 deg. The results indicate that deviation angles can be estimated accurately (within 1 deg) by using one of the three trailing-edge hypotheses, but only when pressure losses are low. A new trailing-edge hypotheses is presented which is suitable (for the cascade considered) for both low- and high-loss operating points.

Miller, M. J.↗

NASA NTRS1974

Mean and fluctuating flow measurements in axisymmetric supersonic boundary layer flow subjected to distributed adverse pressure gradients

Measurements have been made of the mean flow properties and turbulent fluctuations in adiabatic turbulent boundary layer flows subjected to distributed adverse pressure gradients. In the freestream region upstream of the adverse pressure gradient the Mach number was 3.86, the unit Reynolds number 5.3 million per foot. The boundary layer developed on the wall of an axisymmetric nozzle and straight test section. In order to avoid the effects of streamwise surface curvature the adverse pressure gradients at the test section wall were induced by contoured centerbodies mounted on the wind tunnel centerline. The flow under study simulated that which might be found in an axially symmetric engine inlet of a supersonic aircraft.

Gootzait, E.↗

NASA NTRS1974

Deviation angle estimation for axial-flow compressors using inviscid flow solutions

Development of a more satisfactory method of estimating deviation angles was attempted using analytical procedures. Solutions for inviscid, irrotational flow in the blade-to-blade plane were obtained with a finite-difference calculation method. Deviation angles for three cascades with rounded trailing edges were estimated using the inviscid flow solutions and four existing trailing edge hypotheses. The estimated deviation angles were compared with experimental values for a range of incidence angle and inlet flow angle. None of the hypotheses were satisfactory at all incidence angles for all the cascades considered. An alternative hypothesis is presented which is suitable for one of the cascade profiles over a range of incidence and blade setting angle.

Miller, M. J.↗

NASA NTRS1974

Iterative solution of transonic flows over airfoils and wings, including flows at Mach 1

A new method of calculating transonic flows based on a 'rotated' difference scheme is described. It is suitable for the calculation of both two- and three-dimensional flows without restriction on the speed at infinity and is well adapted to computer use. The Murman procedure is modified to eliminate any assumptions about the direction of flow when constructing the difference scheme. The proper directional property is obtained by rotating the difference scheme to conform with the local stream direction. In the hyperbolic region retarded difference formulas are used for all contributions to the streamwise second derivative, producing a correctly oriented positive artificial viscosity. In the absence of a simple implicit scheme in the hyperbolic and elliptic regions, the concept of iterations as steps in artificial time is introduced. Computer testing of this procedure provides numerical confirmation of the existence and uniqueness of weak solutions of the potential equation when a suitable entropy inequality is enforced.

Jameson, A.↗

NASA NTRS1974

Calculation of flow distribution in large radius ratio stages of axial flow turbines and comparison of theory and experiment

A method of calculating stage parameters and flow distribution of axial turbines is described. The governing equations apply to space between the blade rows and are based on the assumption of rotationally symmetrical, compressible, adiabatic flow conditions. Results are presented for stage design and flow analysis calculations. Theoretical results from the calculation system are compared with experimental data from low pressure steam turbine tests.

Herzog, J.↗

NASA NTRS1976

An experimental and computational investigation of the flow field about a transonic airfoil in supercritical flow with turbulent boundary-layer separation

A combined experimental and computational research program is described for testing and guiding turbulence modeling within regions of separation induced by shock waves incident in turbulent boundary layers. Specifically, studies are made of the separated flow the rear portion of an 18%-thick circular-arc airfoil at zero angle of attack in high Reynolds number supercritical flow. The measurements include distributions of surface static pressure and local skin friction. The instruments employed include highfrequency response pressure cells and a large array of surface hot-wire skin-friction gages. Computations at the experimental flow conditions are made using time-dependent solutions of ensemble-averaged Navier-Stokes equations, plus additional equations for the turbulence modeling.

Rubesin, M. W.↗

NASA NTRS1976

Noise generated by impingement of turbulent flow on airfoils of varied chord, cylinders, and other flow obstructions

Noise spectra were measured in three dimensions for several surfaces immersed in turbulent flow from a jet and over a range of flow conditions. The data are free field and were corrected to remove the small contributions of jet noise, atmospheric attenuation and feedback tones. These broadband data were compared with the results of available theories which are only strictly applicable to simple geometries over a limited range of conditions. The available theories proved to be accurate over the range of flow, chord length, thickness, angle of attack, and surface geometries defined by the experiments. These results apply to the noise generated by fixed surfaces in engine passages, the lifting surfaces of aircraft and also to fan noise.

Olsen, W. 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

Noise generated by impingement of turbulent flow on airfoils of varied chord, cylinders, and other flow obstructions

Olsen, W. A.↗

NASA NTRS1976

An experimental and computational investigation of the flow field about a transonic airfoil in supercritical flow with turbulent boundary-layer separation

A combined experimental and computational research program for testing and guiding turbulence modeling within regions of separation induced by shock waves incident on turbulent boundary layers is described. Specifically, studies are made of the separated flow over the rear portion of an 18%-thick circular-arc airfoil at zero angle of attack in high Reynolds number supercritical flow. The measurements include distributions of surface static pressure and local skin friction. The instruments employed include high-frequency response pressure cells and a large array of surface hot-wire skin-friction gages. Computations at the experimental flow conditions are made using time-dependent solutions of ensemble-averaged Navier-Stokes equations, plus additional equations for the turbulence modeling.

Rubesin, M. W.↗

NASA NTRS1977

Measurements of surface-pressure and wake-flow fluctuations in the flow field of a whitcomb supercritical airfoil

Measurements of surface pressure and wake flow fluctuations were made as part of a transonic wind tunnel investigation into the nature of a supercritical airfoil flow field. Emphasis was on a range of high subsonic Mach numbers and moderate lift coefficients corresponding to the development of drag divergence and buffeting. Fluctuation data were analyzed statistically for intensity, frequency content, and spatial coherence. Variations in these parameters were correlated with changes in the mean airfoil flow field.

Roos, F. W.↗

NASA NTRS1976

The scarcity of mappable flow lobes on the lunar maria - Unique morphology of the Imbrium flows

Unique features of Imbrium lava flows are their thickness (10-30 + m) and lengths (up to 400 km for phase-III, and up to 600 km for phase-II) when compared along later Imbrium and Eratosthenian units. They are distinct by virtue of their inferred short-lived (on the order of days) and extremely rapid rates of effusion. It is shown that there are numerous other basalt eruptives within the young blue western maria. The emplacement of these flows was by complex multilayering and intertonguing of individual flow units with thicknesses less than 10 m. They are generally restricted to the near vicinity of multiple vent sources.

Schaber, G. G.↗

NASA NTRS1978

Measurements in the turbulent boundary layer at constant pressure in subsonic and supersonic flow. Part 1: Mean flow

Experiments were carried out to test the accuracy of laser Doppler instrumentation for measurement of Reynolds stresses in turbulent boundary layers in supersonic flow. Two facilities were used to study flow at constant pressure. In one facility, data were obtained on a flat plate at M sub e = 0.1, with Re theta up to 8,000. In the other, data were obtained on an adiabatic nozzle wall at M sub e = 0.6, 0.8, 1.0, 1.3, and 2.2, with Re theta = 23,000 and 40,000. The mean flow as observed using Pitot tube, Preston tube, and floating element instrumentation is described. Emphasis is on the use of similarity laws with Van Driest scaling and on the inference of the shearing stress profile and the normal velocity component from the equations of mean motion. The experimental data are tabulated.

Collins, D. J.↗

NASA NTRS1978

Computational flow development for unsteady viscous flows: Foundation of the numerical method

A procedure is presented for effective consideration of viscous effects in computational development of high Reynolds number flows. The procedure is based on the interpretation of the Navier-Stokes equations as vorticity transport equations. The physics of the flow was represented in a form suitable for numerical analysis. Lighthill's concept for flow development for computational purposes was adapted. The vorticity transport equations were cast in a form convenient for computation. A statement for these equations was written using the method of weighted residuals and applying the Galerkin criterion. An integral representation of the induced velocity was applied on the basis of the Biot-Savart law. Distribution of new vorticity, produced at wing surfaces over small computational time intervals, was assumed to be confined to a thin region around the wing surfaces.

Bratanow, T.↗

NASA NTRS1978

Measurements of inlet flow distortions in an axial flow fan (6 and 9 blade rotor)

A large quantity of experimental data on inlet flow distortions in an axial flow fan were obtained. The purpose of the study was to determine the effects of design and operating variables and the type of distortion on the response of an axial flow turbomachinery rotor. Included are background information and overall trends observed in distortion attenuation and unsteady total pressure losses.

Barr, L. C.↗

NASA NTRS1978

The effects of design and operating variables on the response of an axial flow fan to inlet flow distortions

The results of a study of total pressure and velocity circumferential distortions in an axial-flow fan are presented. Some of the fundamental experimental data needed to understand distorted flow phenomena as affected by design and operating variables are provided. The flow through an isolated rotor was examined at various operating conditions with six different distortions and three different blade stagger angles. Circumferential surveys were conducted upstream and downstream of the rotor using five-hole probes in the nonnulling mode. The total pressure and axial velocity distortion data were analyzed to determine the degree of distortion attenuation as a function of blade stagger angle, mean incidence angle, and reduced frequency. The results indicate that, for the rotors tested, the mean incidence or loading has very little effect on the distortion attenuation.

Yocum, A. M., II↗