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At least 163 records · Page 9
All sourcesNASA NTRSDOE OSTI
NASA NTRS1977

Performance of high-area-ratio annular dump diffuser using suction-stabilized-vortex flow control

A short annular dump diffuser having a geometry conductive to formation of suction stabilized toroidal vortices in the region of abrupt area change was tested. The overall diffuser area ratio was 4.0 and the length to inlet height ratio was 2.0. Performance data were obtained at near ambient temperature and pressure for inlet Mach numbers of 0.18 and 0.30 with suction rates ranging from 0 to 18 percent of total inlet mass flowrate. Results show that the exit velocity profile could be readily biased toward either wall by adjustment of inner and outer wall suction rates. Symmetric exit velocity profiles were inherently unstable with a tendency to revert to a hub or tip bias. Diffuser effectiveness was increased from about 38 percent without suction to over 85 percent at a total suction rate of 10 to 12 percent. At the same time diffuser total pressure loss was reduced from 3.1 percent to 1.1 percent at an inlet Mach number of 0.3.

Juhasz, A. J.↗
NASA NTRS1978

Noise calculation on the basis of vortex flow models

Flow-modeling technique yields relatively simple method for calculating sound radiation involving planar, cylindrical, or spherical surfaces. Model employs potential flow theory with action of viscosity on flowfield described in terms of point vortices. Surface presence in flow is analyzed, using classical image method; sound is calculated through sound generation theory reformulation.

Hardin, J. C.↗
NASA NTRS1978

Performance characteristics of two annular dump diffusers using suction-stabilized vortex flow control

Test results are described for two abrupt area change annular diffusers with provisions for maintaining suction stabilized toroidal vortices at the area discontinuity. Both diffusers had an overall area ratio of 4.0 with the prediffuser area ratio being 1.18 for diffuser A and 1.4 for diffuser B. Performance was evaluated at near atmospheric pressure and temperature for a range of inlet Mach numbers from 0.18 to 0.41 and suction rates from 0 to 18%. Static pressure recovery improved significantly as the suction rate was increased to approximately 11%. Results obtained with diffuser A were superior to that obtained with diffuser B. Flat radial profiles of exit velocity were not obtained since the flow showed preferential hub or tip attachment at moderate suction rates. At high suction rates the diffuser exit flow became circumferentially nonuniform and unstable.

Juhasz, A. J.↗
NASA NTRS1978

Measurements of unsteady vortex flow fields

A combined surface hot film and laser velocimeter measurement technique, used to obtain new information on the mean, constant phase-averaged and turbulent structure of time-dependent flow fields, is described. Data obtained in a cylinder wake are presented, and its structure in both the Eulerian and Lagrangian frames is discussed. Turbulence data obtained by conventional and conditional averaging of the velocity fluctuations are also presented. These data provide details of the small- and large-scale contributions to the total turbulent field.

Owen, F. K.↗
NASA NTRS1985

Fundamental aerodynamic characteristics of delta wings with leading-edge vortex flows

An investigation of the aerodynamics of sharp leading-edge delta wings at supersonic speeds has been conducted. The supporting experimental data for this investigation were taken from published force, pressure, and flow-visualization data in which the Mach number normal to the wing leading edge is always less than 1.0. The individual upper- and lower-surface nonlinear characteristics for uncambered delta wings are determined and presented in three charts. The upper-surface data show that both the normal-force coefficient and minimum pressure coefficient increase nonlinearly with a decreasing slope with increasing angle of attack. The lower-surface normal-force coefficient was shown to be independent of Mach number and to increase nonlinearly, with an increasing slope, with increasing angle of attack. These charts are then used to define a wing-design space for sharp leading-edge delta wings.

Wood, R. M.↗
NASA NTRS1986

An Euler aerodynamic method for leading-edge vortex flow simulation

The current capabilities and the future plans for a three dimensional Euler Aerodynamic Method are described. The basic solution algorithm is based on the finite volume, Runge-Kutta pseudo-time-stepping scheme of FLO-57. Several modifications to improve accuracy and computational efficiency were incorporated and others are being investigated. The computer code is used to analyze a cropped delta wing at 0.6 Mach number and an arrow wing at 0.85 Mach number. Computed aerodynamic parameters are compared with experimental data. In all cases, the configuration is impulsively started and no Kutta condition is applied at sharp edges. The results indicate that with additional development and validation, the present method will be a useful tool for engineering analysis of high speed aircraft.

Raj, P.↗
NASA NTRS1986

An overview of the fundamental aerodynamics branch's research activities in wing leading-edge vortex flows at supersonic speeds

For the past 3 years, a research program pertaining to the study of wing leading edge vortices at supersonic speeds has been conducted in the Fundamental Aerodynamics Branch of the High-Speed Aerodynamics Division at the Langley Research Center. The purpose of the research is to provide an understanding of the factors governing the formation and the control of wing leading-edge vortices and to evaluate the use of these vortices for improving supersonic aerodynamic performance. The studies include both experimental and theoretical investigations and focus primarily on planform, thickness and camber effects for delta wings. An overview of this research activity is presented.

Miller, D. S.↗
NASA NTRS1986

Nonlinear lift control at high speed and high angle of attack using vortex flow technology

Nonlinear lift control at subsonic, transonic and low supersonic speeds owes its origin to the separated but organized vortical flows interacting with the wing upper surface. Since most of this flow originates near the wing or control-surface leading-edge, a variety of devices have been studied experimentally which interact with and/or control this flow in order to gain a beneficial effect. The benefits (effects) originally studied were only associated with lift enhancement. Whereas, now the studied benefits encompass performance increase, attention to changes in trimmed conditions and longitudinal stability, improvements in lateral stability, and the attendant variation with changing Mach number. For those devices that can be theoretically modeled, state-of-the-art computer codes have been used for device design and/or analysis. Comparisons at design and off-design conditions are presented for validation purposes.

Lamar, J. E.↗
NASA NTRS1985

High resolution solutions of the Euler equations for vortex flows

Solutions of the Euler equations are presented for M = 1.5 flow past a 70-degree-swept delta wing. At an angle of attack of 10 degrees, strong leading-edge vortices are produced. Two computational approaches are taken, based upon fully three-dimensional and conical flow theory. Both methods utilize a finite-volume discretization solved by a pseudounsteady multistage scheme. Results from the two approaches are in good agreement. Computations have been done on a 16-million-word CYBER 205 using 196 x 56 x 96 and 128 x 128 cells for the two methods. A sizable data base is generated, and some of the practical aspects of manipulating it are mentioned. The results reveal many interesting physical features of the compressible vortical flow field and also suggest new areas needing research.

Murman, E. M.↗
NASA NTRS1986

Unsteady hybrid vortex technique for transonic vortex flows and flutter applications

The well developed surface panel method is extended by adding a volume integral term allowing calculation of the full effect of compressibility. The full effect of compressibility is calculated by using mixed type finite difference scheme consistent with the mixed nature of transonic flow. The solution is obtained through successive iteration cycles for subcritical flows and for critical flows the solution is obtained through satisfying the Rankine-Hugoniot relations across the captured shock in addition to the captured shock. The method is applied to airfoils in subcritical and critical flows and the results are in good agreement with the experimental data and finite-difference solutions of the full-potential and Euler equations.

Kandil, Osama A.↗
NASA NTRS1987

Finite-volume Euler and Navier-Stokes solvers for three-dimensional and conical vortex flows over delta wings

A unified central-difference finite-volume Euler and Navier-Stokes solver with four-stage Runge-Kutta time stepping is presented. The computer code developed for this purpose is capable of solving the standard set and nonstandard sets (zero-total-pressure loss) of Euler equations and the thin-layer and full Navier-Stokes equations. Applications are presented for conical supersonic flows with weak shocks using the standard and nonstandard sets of Euler equations, and the thin-layer and full Navier-Stokes equations for sharp and round-edged delta wings. Applications are also presented for three-dimensional transonic and subsonic flows using the standard set of Euler equations for sharp-edged delta wings. The computational results of the different sets of equations are compared with each other and with the experimental results and conclusions on the validity of these sets to these applications, are presented.

Kandil, Osama A.↗