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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 883 records · Page 49
All sourcesNASA NTRSDOE OSTI
NASA NTRS1981

Application of holography to the study of helicopter rotor flow fields

The feasibility of an experiment which is intended to measure the density field about a model helicopter rotor using holographic interferometry is considered. The numerical simulation used to study the experiment is described as well as the measurement technique itself. Data generated by the simulation are presented and prospects for both determining the density field from these data, and for actually obtaining such data in practice are assessed. A few significant problems which may be expected to arise are indicated and discussed.

Baganoff, D.↗
NASA NTRS1983

Application of a full potential method for predicting supersonic flow fields and aerodynamic characteristics

A nonlinear aerodynamic prediction technique which solves the conservative full potential equation has been applied to the analysis of three waverider configurations. This technique was selected based on its capability to analyze the off-design characteristics of the waveriders. Very good correlations were achieved with surface pressure data for both the Mach 4 elliptic cone waverider and the Mach 6 caret-wing derivative. Off-design Mach number and angle-of-attack pressure correlations were very good for the elliptic cone waverider. The range of correlation with data exceeded that expected based on the theory limitations. A surface pressure integration routine was demonstrated and agreement between predicted aerodynamic forces and experimental force data for the Mach 4 waverider was excellent. Analysis of a nonconical waverider configuration was initiated where a discrete input option is used to achieve the computational gridding. Preliminary analysis of this configuration indicates the correct shock location will be predicted.

Jones, K. M.↗
NASA NTRS1983

The efficient solution of transonic wing flow fields

An evaluation of the transonic-wing-analysis computer code TWING is presented. TWING utilizes a fully implicit, approximate-factorization iteration scheme to solve the full-potential equation in conservative form. A numerical elliptic-solver grid-generation scheme is used to generate the required finite-difference mesh. Several wing configurations have been analyzed, and comparisons of computed results have been made with available experimental data. Results indicate that the code is robust, accurate (when significant viscous effects are not present), and efficient. TWING generally produces solutions an order of magnitude faster than other conservative, full-potential codes using successive-line overrelaxation. The present method is applicable to a wide range of isolated wing configurations, including high-aspect-ratio transport wings and low-aspect-ratio, high-sweep, fighter configurations.

Holst, T. L.↗
NASA NTRS1982

Experimental trim drag values and flow-field measurements for a wide-body transport model with conventional and supercritical wings

The purpose of this study was to determine if advanced supercritical wings incur higher trim drag values at cruise conditions than current wide body technology wings. Relative trim drag increments were measured in an experimental wind tunnel investigation conducted in the Langley 8 Foot Transonic Pressure Tunnel. The tests utilized a high aspect ratio supercritical wing and a wide body aircraft wing, in conjunction with five different horizontal tail configurations, mounted on a representative wide body fuselage. The three low tail and two T-tail configurations were designed to measure the effects of horizontal tail size, location, and camber on the trim drag increments for the two wings. Longitudinal force and moment data were taken at a Mach number of 0.82 and design cruise lift coefficients for the wide body and supercritical wings of 0.45 and 0.55, respectively. The data indicate that the supercritical wing does not have significantly higher trim drag than the wide body wing. A reduction in tail size, combined with relaxed static stability, produced trim drag reductions for both wings. The cambered tails had higher trim drag increments than the symmetrical tails for both wings, and the T-tail configurations had lower trim drag increments than the low tail configurations.

Jacobs, P. F.↗
NASA NTRS1985

Inviscid analysis of advanced turboprop propeller flow fields

A three-dimensional Euler code, previously developed for high speed propellers, was rewritten and vectorized for the Cray computer. Calculations were compared with previous results and measurements. The new code requires less memory and is more than two times faster than its predecessor. The new boundary conditions produced more accurate results and improved stability. The accuracy and stability characteristics also resulted in reduced manual intervention required to obtain solutions. The code is a useful tool for examining the fluid dynamics and performance of propellers.

Barton, J. M.↗
NASA NTRS1987

Automated reduction of instantaneous flow field images

An automated data reduction system for the analysis of interference fringe patterns obtained using the particle image velocimetry technique is described. This system is based on digital image processing techniques that have provided the flexibility and speed needed to obtain more complete automation of the data reduction process. As approached here, this process includes scanning/searching for data on the photographic record, recognition of fringe patterns of sufficient quality, and, finally, analysis of these fringes to determine a local measure of the velocity magnitude and direction. The fringe analysis as well as the fringe image recognition are based on full frame autocorrelation techniques using parallel processing capabilities.

Reynolds, G. A.↗
NASA NTRS1987

Mechanisms by which heat release affects the flow field in a chemically reacting, turbulent mixing layer

The mechanisms by which heat release affects the fluid dynamics in a turbulent reacting mixing layer are studied by direct numerical simulation. In agreement with previous laboratory experiments, the heat release is observed to lower the rate at which the mixing layer grows and to reduce the rate at which chemical products are formed. The baroclinic torque and thermal expansion in the mixing layer are shown to produce changes in the flame vortex structure that act to produce more diffuse vortices than in the constant density case, resulting in lower rotation rates of fluid elements. Previously unexplained anomalies observed in the mean velocity profiles of reacting jets and mixing layers are shown to result from vorticity generation by baroclinic torques. The density reductions also lower the generation rates of turbulent kinetic energy and the turbulent shear stresses, resulting in less turbulent mixing of fluid elements. Calculations of the energy in the various wave numbers show that the heat release has a stabilizing effect on the growth rates of individual modes. A linear stability analysis of a simlified model problem confirms this, showing that low density fluid in the mixing region will result in a shift of the frequency of the unstable modes to lower wave numbers (longer wavelengths). The growth rates of the unstable modes decrease, contributing to the slower growth of the mixing layer.

Riley, J. J.↗
NASA NTRS1987

Unsteady aerodynamics of an oscillating cascade in a compressible flow field

Fundamental experiments were performed in the NASA Lewis Transonic Oscillating Cascade Facility to investigate and quantify the unsteady aerodynamics of a cascade of biconvex airfoils executing torsion-mode oscillations at realistic reduced frequencies. Flush-mounted, high-response miniature pressure transducers were used to measure the unsteady airfoil surface pressures. The pressures were measured for three interblade phase angles at two inlet Mach numbers, 0.65 and 0.80, and two incidence angles, 0 and 7 deg. The time-variant pressures were analyzed by means of discrete Fourier transform techniques, and these unique data were then compared with predictions from a linearized unsteady cascade model. The experimental results indicate that the interblade phase angle had a major effect on the chordwise distributions of the airfoil surface unsteady pressure, and that reduced frequency, incidence angle, and Mach number had a somewhat less significant effect.

Buffum, Daniel H.↗
NASA NTRS1988

A numerical study of mixing enhancement in supersonic reacting flow fields

NASA Langley has intensively investigated the components of ramjet and scramjet systems for endoatmospheric, airbreathing hypersonic propulsion; attention is presently given to the optimization of scramjet combustor fuel-air mixing and reaction characteristics. A supersonic, spatially developing and reacting mixing layer has been found to serve as an excellent physical model for the mixing and reaction process. Attention is presently given to techniques that have been applied to the enhancement of the mixing processes and the overall combustion efficiency of the mixing layer. A fuel injector configuration has been computationally designed which significantly increases mixing and reaction rates.

Drummond, J. Philip↗
NASA NTRS1988

CFD prediction of the reacting flow field inside a subscale scramjet combustor

A three-dimensional, Reynolds-averaged Navier-Stokes CFD code has been used to calculate the reacting flowfield inside a hydrogen-fueled, subscale scramjet combustor. Pilot fuel was injected transversely upstream of the combustor and the primary fuel was injected transversely downstream of a backward facing step. A finite rate combustion model with two-step kinetics was used. The CFD code used the explicit MacCormack algorithm with point-implicit treatment of the chemistry source terms. Turbulent mixing of the jets with the airstream was simulated by a simple mixing length scheme, whereas near wall turbulence was accounted for by the Baldwin-Lomax model. Computed results were compared with experimental wall pressure measurements.

Chitsomboon, T.↗
NASA NTRS1989

Experimental and computational flow-field results for an all-body hypersonic aircraft

A comprehensive test program is defined which is being implemented in the NASA/Ames 3.5 foot Hypersonic Wind Tunnel for obtaining data on a generic all-body hypersonic vehicle for computational fluid dynamics (CFD) code validation. Computational methods (approximate inviscid methods and an upwind parabolized Navier-Stokes code) currently being applied to the all-body model are outlined. Experimental and computational results on surface pressure distributions and Pitot-pressure surveys for the basic sharp-nose model (without control surfaces) at a free-stream Mach number of 7 are presented.

Cleary, Joseph W.↗
NASA NTRS1988

Fundamental problem in computing radiating flow fields with thick shock waves

Possible reasons for the failure of the Navier-Stokes equations to yield realistic profiles of the temperature and density through the structure of a hypersonic shock wave are investigated. Models for bulk viscosity in a monatomic gas are examined which yield a realistic thickness for the shock-wave density profile, but not the temperature profile, and hence are not satisfactory. A tentative computational model for nitrogen is explored which yields considerably more realistic results than the Navier-Stokes equations. This model involves a nonlinear stress-strain tensor, nonlinear heat flux vector, and nonequilibrium rotational energy.

Chapman, Dean R.↗
NASA NTRS1989

Influence of shock-shock interactions on the blunt body flow field at hypersonic flight speeds

The effect of shock impingement on the blunt leading edge of the inlet cowl of a scramjet engine is investigated numerically. The impinging shock is caused by the vehicle forebody. The interaction of this forebody shock with the cowl leading edge shock results in a very complex flowfield containing local regions of high pressure and intense heating. In the present investigation, this complex flowfield is calculated by solving the full Navier-Stokes equations using a finite-volume flux splitting technique due to Van Leer. Results are also obtained for the downstream effects into the inlet of the leading edge shock-shock interactions. Results of the present numerical investigation are compared with available experimental results.

Singh, D. J.↗