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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 253 records · Page 14

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NASA NTRS1987

High speed propeller performance and noise predictions at takeoff/landing conditions

The performance and noise of a high speed SR-7A model propeller under takeoff/landing conditions are considered. The blade loading distributions are obtained by solving the three-dimensional Euler equations and the sound pressure levels are computed using a time domain approach. At the nominal takeoff operating point, the blade sections near the hub are lightly or negatively loaded. The chordwise loading distributions are distinctly different from those of cruise conditions. The noise of the SR-7A model propeller at takeoff is dominated by the loading noise, similar to that at cruise conditions. The waveforms of the acoustic pressure signature are nearly sinusoidal in the plane of the propeller. The computed directivity of the blade passing frequency tone agrees fairly well with the data at nominal takeoff blade angle.

Nallasamy, M.↗

NASA NTRS1988

High speed propeller performance and noise predictions at takeoff/landing conditions

Nallasamy, M.↗

NASA NTRS1987

Program VSAERO theory document: A computer program for calculating nonlinear aerodynamic characteristics of arbitrary configurations

The VSAERO low order panel method formulation is described for the calculation of subsonic aerodynamic characteristics of general configurations. The method is based on piecewise constant doublet and source singularities. Two forms of the internal Dirichlet boundary condition are discussed and the source distribution is determined by the external Neumann boundary condition. A number of basic test cases are examined. Calculations are compared with higher order solutions for a number of cases. It is demonstrated that for comparable density of control points where the boundary conditions are satisfied, the low order method gives comparable accuracy to the higher order solutions. It is also shown that problems associated with some earlier low order panel methods, e.g., leakage in internal flows and junctions and also poor trailing edge solutions, do not appear for the present method. Further, the application of the Kutta conditions is extremely simple; no extra equation or trailing edge velocity point is required. The method has very low computing costs and this has made it practical for application to nonlinear problems requiring iterative solutions for wake shape and surface boundary layer effects.

Maskew, Brian↗

NASA NTRS1994

Skewness in large-scale structure and non-Gaussian initial conditions

We compute the skewness of the galaxy distribution arising from the nonlinear evolution of arbitrary non-Gaussian intial conditions to second order in perturbation theory including the effects of nonlinear biasing. The result contains a term identical to that for a Gaussian initial distribution plus terms which depend on the skewness and kurtosis of the initial conditions. The results are model dependent; we present calculations for several toy models. At late times, the leading contribution from the initial skewness decays away relative to the other terms and becomes increasingly unimportant, but the contribution from initial kurtosis, previously overlooked, has the same time dependence as the Gaussian terms. Observations of a linear dependence of the normalized skewness on the rms density fluctuation therefore do not necessarily rule out initially non-Gaussian models. We also show that with non-Gaussian initial conditions the first correction to linear theory for the mean square density fluctuation is larger than for Gaussian models.

Fry, J. N.↗

NASA NTRS1984

Comparison of viscous-shock-layer heating analysis with Shuttle flight data in slip flow regime

Comparison of STS-2 Shuttle flight heating data along the windward centerline has been made with two-dimensional nonequilibrium viscous shock-layer solutions obtained with shock and wall-slip conditions at an altitude range of 90 to 110 km. The shock slip condition used is the modified Rankine-Hugoniot relations of Cheng as used by Davis, and the wall-slip conditions are based on the first order consideration derived from kinetic theory as given by Scott and Hendricks. The results indicate that the calculated heating distributions with slip boundary conditions agree better with the flight data than those without slip conditions. The agreement improves when the accommodation coefficient or freestream density is decreased to one-half, suggesting the possibility of less than full accommodation for the tile surface and (or) an overestimate of freestream density using the Jacchia-Roberts model. Heating reduction due to the slip effect becomes very pronounced as the flow becomes more rarefied, and the effect is more significant for the stagnation region than the aft region of the vehicle.

Shinn, J. L.↗

NASA NTRS1973

A boundary layer approach to the analysis of atmospheric motion over a surface obstruction

A boundary layer approach for the solution of the flow field induced over a two-dimensional surface obstruction, such as a building or other man-made structure, is proposed. Adopting a specific geometry in the form of a semi-elliptical cylinder, the characteristics of atmospheric shear flow over a rough terrain are coupled with the well-known boundary layer equations. Two approaches are presented to incorporate the pressure field and boundary conditions which exist within the large viscous region over the obstruction. The first considers a region in the immediate vicinity of the body in which the pressure distribution and outer boundary condition on the velocity are computed from potential theory for flow over the elliptical cylinder. The second approach considers a much larger region of influence, extending from the surface to the undisturbed flow at large heights above the obstruction. Methods which appear to provide an improved theoretical model of the flow over the ellipse, such as a technique for simulating the effect of the seperation regions upstream and downstream of the body, are also presented.

Frost, W.↗

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 NTRS1992

Preliminary results from the White Sands Missile Range sonic boom propagation experiment

Sonic boom bow shock amplitude and rise time statistics from a recent sonic boom propagation experiment are presented. Distributions of bow shock overpressure and rise time measured under different atmospheric turbulence conditions for the same test aircraft are quite different. The peak overpressure distributions are skewed positively, indicating a tendency for positive deviations from the mean to be larger than negative deviations. Standard deviations of overpressure distributions measured under moderate turbulence were 40 percent larger than those measured under low turbulence. As turbulence increased, the difference between the median and the mean increased, indicating increased positive overpressure deviations. The effect of turbulence was more readily seen in the rise time distributions. Under moderate turbulence conditions, the rise time distribution means were larger by a factor of 4 and the standard deviations were larger by a factor of 3 from the low turbulence values. These distribution changes resulted in a transition from a peaked appearance of the rise time distribution for the morning to a flattened appearance for the afternoon rise time distributions. The sonic boom propagation experiment consisted of flying three types of aircraft supersonically over a ground-based microphone array with concurrent measurements of turbulence and other meteorological data. The test aircraft were a T-38, an F-15, and an F-111, and they were flown at speeds of Mach 1.2 to 1.3, 30,000 feet above a 16 element, linear microphone array with an inter-element spacing of 200 ft. In two weeks of testing, 57 supersonic passes of the test aircraft were flown from early morning to late afternoon.

Willshire, William L., Jr.↗

NASA NTRS1974

Aerodynamic influence coefficient method using singularity splines

A numerical lifting surface formulation, including computed results for planar wing cases is presented. This formulation, referred to as the vortex spline scheme, combines the adaptability to complex shapes offered by paneling schemes with the smoothness and accuracy of loading function methods. The formulation employes a continuous distribution of singularity strength over a set of panels on a paneled wing. The basic distributions are independent, and each satisfied all the continuity conditions required of the final solution. These distributions are overlapped both spanwise and chordwise. Boundary conditions are satisfied in a least square error sense over the surface using a finite summing technique to approximate the integral. The current formulation uses the elementary horseshoe vortex as the basic singularity and is therefore restricted to linearized potential flow. As part of the study, a non planar development was considered, but the numerical evaluation of the lifting surface concept was restricted to planar configurations. Also, a second order sideslip analysis based on an asymptotic expansion was investigated using the singularity spline formulation.

Mercer, J. E.↗

NASA NTRS1993

Thermoviscoplastic analysis of fibrous periodic composites using triangular subvolumes

The nonlinear viscoplastic behavior of fibrous periodic composites is analyzed by discretizing the unit cell into triangular subvolumes. A set of these subvolumes can be configured by the analyst to construct a representation for the unit cell of a periodic composite. In each step of the loading history, the total strain increment at any point is governed by an integral equation which applies to the entire composite. A Fourier series approximation allows the incremental stresses and strains to be determined within a unit cell of the periodic lattice. The nonlinearity arising from the viscoplastic behavior of the constituent materials comprising the composite is treated as fictitious body force in the governing integral equation. Specific numerical examples showing the stress distributions in the unit cell of a fibrous tungsten/copper metal matrix composite under viscoplastic loading conditions are given. The stress distribution resulting in the unit cell when the composite material is subjected to an overall transverse stress loading history perpendicular to the fibers is found to be highly heterogeneous, and typical homogenization techniques based on treating the stress and strain distributions within the constituent phases as homogeneous result in large errors under inelastic loading conditions.

Walker, Kevin P.↗

NASA NTRS2013

Interaction of a Rectangular Jet with a Flat-Plate Placed Parallel to the Flow

An experimental study is carried out addressing the flowfield and radiated noise from the interaction of a large aspect ratio rectangular jet with a flat plate placed parallel to but away from the direct path of the jet. Sound pressure level spectra exhibit an increase in the noise levels for both the 'reflected' and 'shielded' sides of the plate relative to the free-jet case. Detailed cross-sectional distributions of flowfield properties obtained by hot-wire anemometry are documented for a low subsonic condition. Corresponding mean Mach number distributions obtained by Pitot-probe surveys are presented for high subsonic conditions. In the latter flow regime and for certain relative locations of the plate, a flow resonance accompanied by audible tones is encountered. Under the resonant condition the jet cross-section experiences an 'axis-switching' and flow visualization indicates the presence of an organized 'vortex street'. The trends of the resonant frequency variation with flow parameters exhibit some similarities to, but also marked differences with, corresponding trends of the well-known edgetone phenomenon.

Zaman, K. B. M. Q.↗

NASA NTRS1948

Preliminary Investigation to Determine Propeller Section Characteristics by Measuring the Pressure Distribution on an NACA 10-(3)(08)-03 Propeller Under Operating Conditions

An investigation has been made in the Langley 16-foot high-speed tunnel to determine the propeller-section characteristics by measuring the pressure distribution on the airfoil sections of a rotating propeller. The pressures were measured at nine radial stations on an NACA 10-(3)(08)-03 design two-blade propeller. This paper presents the results of the pressure measurements in the form of normal-force and moment coefficients and covers a range of nominal angle of attack (simple blade element theory) from 0 degrees to 4 degrees for a section Mach number range of approximately 0.6 to 1.15 for the outboard stations and approximately 0.3 to 0.6 for inboard stations.

Evans, Albert J↗

NASA NTRS1994

Thermoviscoplastic analysis of fibrous periodic composites by the use of triangular subvolumes

The non-linear viscoplastic behavior of fibrous periodic composites is analyzed by discretizing the unit cell into triangular subvolumes. A set of these subvolumes can be configured by the analyst to construct a representation for the unit cell of a periodic composite. In each step of the loading history the total strain increment at any point is governed by an integral equation which applies to the entire composite. A Fourier series approximation allows the incremental stresses and strains to be determined within a unit cell of the periodic lattice. The non-linearity arising from the viscoplastic behavior of the constituent materials comprising the composite is treated as a fictitious body force in the governing integral equation. Specific numerical examples showing the stress distributions in the unit cell of a fibrous tungsten/copper metal-matrix composite under viscoplastic loading conditions are given. The stress distribution resulting in the unit cell when the composite material is subjected to an overall transverse stress loading history perpendicular to the fibers is found to be highly heterogeneous, and typical homogenization techniques based on treating the stress and strain distributions within the constituent phases as homogeneous result in large errors under inelastic loading conditions.

Walker, Kevin P.↗

NASA NTRS1979

Sonic-boom minimization with nose-bluntness relaxation

A procedure which provides sonic-boom-minimizing equivalent area distributions for supersonic cruise conditions is described. This work extends previous analyses to permit relaxation of the extreme bluntness required by conventional low-boom shapes and includes propagation in a real atmosphere. The procedure provides area distributions which minimize either shock strength or overpressure.

Darden, C. M.↗

NASA NTRS1990

Evaluation of voice codecs for the Australian mobile satellite system

The evaluation procedure to choose a low bit rate voice coding algorithm is described for the Australian land mobile satellite system. The procedure is designed to assess both the inherent quality of the codec under 'normal' conditions and its robustness under 'severe' conditions. For the assessment, normal conditions were chosen to be random bit error rate with added background acoustic noise and the severe condition is designed to represent burst error conditions when mobile satellite channel suffers from signal fading due to roadside vegetation. The assessment is divided into two phases. First, a reduced set of conditions is used to determine a short list of candidate codecs for more extensive testing in the second phase. The first phase conditions include quality and robustness and codecs are ranked with a 60:40 weighting on the two. Second, the short listed codecs are assessed over a range of input voice levels, BERs, background noise conditions, and burst error distributions. Assessment is by subjective rating on a five level opinion scale and all results are then used to derive a weighted Mean Opinion Score using appropriate weights for each of the test conditions.

Bundrock, Tony↗

NASA NTRS1995

Pressure Distributions from Subsonic Tests of a NACA 0012 Semispan Wing Model

An unswept, semispan wing model incorporating a NACA 0012 airfoil section was tested in the Langley 14- by 22-Foot Subsonic Tunnel. This report contains pressure data which document effects of wing configuration and free-stream conditions on wing pressure distributions. The untwisted wing incorporated a full-span, leading-edge Krueger flap and a full-span, single-slotted trailing-edge flap. The trailing-edge flap was tested at a deflection angle of 40 degrees and the Krueger flap at a deflection of 55 degrees. Three wing configurations were tested: cruise, trailing-edge flap only, and Knueger flap and trailing-edge flap deployed. Tests were conducted at free-stream dynamic pressures of 15, 30 and 60 psf, with corresponding chord Reynolds numbers of 1.22 to 2.11 million and Mach numbers of 0.12 to 0.20. Angles of attack presented range from 0 to 20 degrees, depending on wing configuration. The data are presented without analysis.

Applin, Zachary T.↗

NASA NTRS1995

Wing pressure distributions from subsonic tests of a high-wing transport model

A wind tunnel investigation was conducted on a generic, high-wing transport model in the Langley 14- by 22-Foot Subsonic Tunnel. This report contains pressure data that document effects of various model configurations and free-stream conditions on wing pressure distributions. The untwisted wing incorporated a full-span, leading-edge Krueger flap and a part-span, double-slotted trailing-edge flap system. The trailing-edge flap was tested at four different deflection angles (20 deg, 30 deg, 40 deg, and 60 deg). Four wing configurations were tested: cruise, flaps only, Krueger flap only, and high lift (Krueger flap and flaps deployed). Tests were conducted at free-stream dynamic pressures of 20 psf to 60 psf with corresponding chord Reynolds numbers of 1.22 x 10(exp 6) to 2.11 x 10(exp 6) and Mach numbers of 0.12 to 0.20. The angles of attack presented range from 0 deg to 20 deg and were determined by wing configuration. The angle of sideslip ranged from minus 20 deg to 20 deg. In general, pressure distributions were relatively insensitive to free-stream speed with exceptions primarily at high angles of attack or high flap deflections. Increasing trailing-edge Krueger flap significantly reduced peak suction pressures and steep gradients on the wing at high angles of attack. Installation of the empennage had no effect on wing pressure distributions. Unpowered engine nacelles reduced suction pressures on the wing and the flaps.

Applin, Zachary T.↗

NASA NTRS1971

Optimal startup control of a jacketed tubular reactor.

The optimal startup policy of a jacketed tubular reactor, in which a first-order, reversible, exothermic reaction takes place, is presented. A distributed maximum principle is presented for determining weak necessary conditions for optimality of a diffusional distributed parameter system. A numerical technique is developed for practical implementation of the distributed maximum principle. This involves the sequential solution of the state and adjoint equations, in conjunction with a functional gradient technique for iteratively improving the control function.

Hahn, D. R.↗