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At least 523 records · Page 29

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NASA NTRS1981

Computer program documentation for a subcritical wing design code using higher order far-field drag minimization

A subsonic, linearized aerodynamic theory, wing design program for one or two planforms was developed which uses a vortex lattice near field model and a higher order panel method in the far field. The theoretical development of the wake model and its implementation in the vortex lattice design code are summarized and sample results are given. Detailed program usage instructions, sample input and output data, and a program listing are presented in the Appendixes. The far field wake model assumes a wake vortex sheet whose strength varies piecewise linearly in the spanwise direction. From this model analytical expressions for lift coefficient, induced drag coefficient, pitching moment coefficient, and bending moment coefficient were developed. From these relationships a direct optimization scheme is used to determine the optimum wake vorticity distribution for minimum induced drag, subject to constraints on lift, and pitching or bending moment. Integration spanwise yields the bound circulation, which is interpolated in the near field vortex lattice to obtain the design camber surface(s).

Kuhlman, J. M.↗

NASA NTRS1987

Distortion of the vortex core during blade-vortex interaction

The effect of vortex distortion on the flow arising from the interaction of a convecting vortex and an airfoil is analyzed for two-dimensional, incompressible, inviscid flow. Significant distortion of the vortex structure, owing to interaction with the airfoil, can occur, especially during close interactions. The vorticity field is represented by multiple vortices. Vortex instability and diffusion, resulting from discretization in space and time, are discussed. An integral boundary element method is used to represent the airfoil surface. The body panels are re-distributed on the surface as the vortices are convected past to calculate the effect of the vorticity field more accurately. The method is applied to several cases to examine effects of vortex strength, distribution, initial separation position, and airfoil angle of attack. Calculations of vortex trajectory and distortion, and airfoil surface-pressure distribution, lift, drag, and circulation as functions of time are presented.

Lee, G. J.↗

NASA NTRS1987

Aerodynamic interaction between vortical wakes and lifting two-dimensional bodies

Unsteady rotor wake interactions with the empenage, tail boom, and other aerodynamic surfaces of a helicopter have a significant influence on its aerodynamic performance, the ride quality, and amount of vibration. A numerical method for computing the aerodynamic interaction between an interacting vortex wake and the viscous flow about arbitrary two-dimensional bodies has been developed to address this helicopter problem. The method solves for the flow field velocities on a body-fitted computational mesh using finite-difference techniques. The interaction of a rotor wake with the flow about a 4:1 elliptic cylinder at 45-deg incidence was calculated for a Reynolds number of 3000.

Stremel, Paul M.↗

NASA NTRS1997

Toward Understanding Wake Vortices and Atmospheric Turbulence Interactions Using Large-Eddy Simulation

The vortices produced by an aircraft in flight are a complex phenomena created from a 'sheet of vorticity' leaving the trailing edge of the aircraft surfaces. This sheet tends to roll-up into two counter-rotating vortices. After a few spans downstream of the aircraft, the roll-up process is complete and the vortex pair may be characterized in a simple manner for modeling purposes. Our research will focus on what happens to these post roll-up vortices in the vicinity of an airport terminal. As the aircraft wake vortices descend, they are transported by the air mass which they are embedded and are decayed by both internal and external processes. In the vicinity of the airport, these external influences are usually due to planetary boundary layer (PBL) turbulence. Using large-eddy simulation (LES), one may simulate a variety of PBL conditions. In the LES method, turbulence is generated in the PBL as a response to surface heat flux, horizontal pressure gradient, wind shear, and/or stratification, and may produce convective or unstably stratified, neutral, or stably stratified PBL's. Each of these PBL types can occur during a typical diurnal cycle of the PBL. Thus it is important to be able to characterize these conditions with the LES method. Once this turbulent environment has been generated, a vortex pair will be introduced and the interactions are observed. The objective is to be able to quantify the PBL turbulence vortex interaction and be able to draw some conclusions of vortex behavior from the various scale interactions. This research is ongoing, and we will focus on what has been accomplished to date and the future direction of this research. We will discuss the model being used, show results that validate its use in the PBL, and present a nested-grid method proposed to analyze the entire PBL and vortex pair simultaneously.

DeCroix, David↗

NASA NTRS1993

A time-accurate finite volume method valid at all flow velocities

A finite volume method to solve the Navier-Stokes equations at all flow velocities (e.g., incompressible, subsonic, transonic, supersonic and hypersonic flows) is presented. The numerical method is based on a finite volume method that incorporates a pressure-staggered mesh and an incremental pressure equation for the conservation of mass. Comparison of three generally accepted time-advancing schemes, i.e., Simplified Marker-and-Cell (SMAC), Pressure-Implicit-Splitting of Operators (PISO), and Iterative-Time-Advancing (ITA) scheme, are made by solving a lid-driven polar cavity flow and self-sustained oscillatory flows over circular and square cylinders. Calculated results show that the ITA is the most stable numerically and yields the most accurate results. The SMAC is the most efficient computationally and is as stable as the ITA. It is shown that the PISO is the most weakly convergent and it exhibits an undesirable strong dependence on the time-step size. The degenerated numerical results obtained using the PISO are attributed to its second corrector step that cause the numerical results to deviate further from a divergence free velocity field. The accurate numerical results obtained using the ITA is attributed to its capability to resolve the nonlinearity of the Navier-Stokes equations. The present numerical method that incorporates the ITA is used to solve an unsteady transitional flow over an oscillating airfoil and a chemically reacting flow of hydrogen in a vitiated supersonic airstream. The turbulence fields in these flow cases are described using multiple-time-scale turbulence equations. For the unsteady transitional over an oscillating airfoil, the fluid flow is described using ensemble-averaged Navier-Stokes equations defined on the Lagrangian-Eulerian coordinates. It is shown that the numerical method successfully predicts the large dynamic stall vortex (DSV) and the trailing edge vortex (TEV) that are periodically generated by the oscillating airfoil. The calculated streaklines are in very good comparison with the experimentally obtained smoke picture. The calculated turbulent viscosity contours show that the transition from laminar to turbulent state and the relaminarization occur widely in space as well as in time. The ensemble-averaged velocity profiles are also in good agreement with the measured data and the good comparison indicates that the numerical method as well as the multipletime-scale turbulence equations successfully predict the unsteady transitional turbulence field. The chemical reactions for the hydrogen in the vitiated supersonic airstream are described using 9 chemical species and 48 reaction-steps. Consider that a fast chemistry can not be used to describe the fine details (such as the instability) of chemically reacting flows while a reduced chemical kinetics can not be used confidently due to the uncertainty contained in the reaction mechanisms. However, the use of a detailed finite rate chemistry may make it difficult to obtain a fully converged solution due to the coupling between the large number of flow, turbulence, and chemical equations. The numerical results obtained in the present study are in good agreement with the measured data. The good comparison is attributed to the numerical method that can yield strongly converged results for the reacting flow and to the use of the multiple-time-scale turbulence equations that can accurately describe the mixing of the fuel and the oxidant.

Kim, S.-W.↗

NASA NTRS1987

Measurement of velocity and vorticity fields in the wake of an airfoil in periodic pitching motion

The velocity field created by the wake of an airfoil undergoing a prescribed pitching motion was sampled using hot wire anemometry. Data analysis methods concerning resolution of velocity components from cross wire data, computation of vorticity from velocity time history data, and calculation of vortex circulation from vorticity field data are discussed. These data analysis methods are applied to a flow field relevant to a two dimensional blade-vortex interaction study. Velocity time history data were differentiated to yield vorticity field data which are used to characterize the wake of the pitching airfoil. Measurement of vortex strength in sinusoidal and nonsinusoidal wakes show vortices in the sinusoidal wake have stronger circulation and more concentrated vorticity distributions than the tailored nonsinusoidal wake.

Booth, Earl R., Jr.↗

NASA NTRS1993

Reduction in size and unsteadiness of a VTOL ground vortex by ground fences

A ground vortex, produced when a jet impinges on the ground in the presence of cross flow, is encountered by V/STOL aircraft hovering near the ground and is known to be hazardous to the aircraft. The objective of this research was to identify a ground-based technique by which both the mean size and fluctuation in size of the ground vortex could be reduced. A simple passive method was identified and examined in the laboratory. Specifically, one or two fine wire mesh screens (ground fences) bent in a horseshoe shape and located on the ground in front of the jet impingement point proved to be very effective. The ground fences work by decreasing the momentum of the upstream-traveling wall jet, effectively causing a higher freestream-to-jet velocity ratio (V(sub infinity)/V(sub j)) and thus, a ground vortex smaller in size and unsteadiness. At(V(sub infinity)/V(sub j)) = 0.15, the addition of a single ground fence resulted in a 70 percent reduction in mean size of the ground vortex. With two ground fences, the mean size decreased by about 85 percent. Fluctuations in size decreased nearly in proportion to the mean size, for both the single and double fence configurations. These results were consistent over a wide range of jet Reynolds number (10(exp 4) less than Re(sub jet) less than 10(exp 5)); further development and full-scale Reynolds number testing are required, however, to determine if this technique can be made practical for the case of actual VTOL aircraft.

Cimbala, John M.↗

NASA NTRS1989

Tip vortex geometry of a hovering helicopter rotor in ground effect

The wide-field shadowgraph method has been used to photograph the tip vortices of a hovering helicopter rotor in ground effect. The shadowgraphs were used to obtain quantitative measurements of the rotor tip vortex geometry both in and out of ground effect. Many important phenomena are visible in the rotor wake using this method. These include the variation in descent and contraction rates of the tip vortices in ground effect, and the interaction between tip vortices in the far wake. The tip vortex geometry from the shadowgraphs is compared with the tip vortex geometry predicted using a free wake hover performance analysis. The free wake analysis accurately predicts the tip vortex geometry both in and out of ground effect. Performance data from the test is compared with the performance predicted using several methods, including the free wake analysis. All methods provided reasonable predictions for the helicopter performance in ground effect.

Light, Jeffrey S.↗

NASA NTRS1989

Goertler vortices in supersonic and hypersonic boundary layers

The problem of Goertler vortices in compressible boundary layers over concave walls is considered. At O(1) wavelengths, the instability is governed by parabolic partial differential equations that are solved numerically to determine the effect of various initial conditions on the development of Goertler vortex instability in compressible boundary layers. The results show that both the velocity and temperature fluctuations may lead to a Goertler vortex. The vortex growth rates determined from the present method are found to differ somewhat from those given by a normal mode solution. At both the supersonic and hypersonic Mach numbers, cooling has a small destabilizing effect. In addition, the most unstable disturbances shift toward lower wavelengths because of thinning of the boundary layer. The results also show that compressibility has a stabilizing effect on the Goertler instability, while the effect of an adverse pressure gradient is found to be destabilizing. The behavior of the Goertler vortex structure with Mach number is also examined. At hypersonic Mach numbers, vortices are located near the edge of the boundary layer for adiabatic wall conditions. However, the entire boundary layer is affected when the wall is cooled.

Spall, R. E.↗

NASA NTRS1984

An experimental investigation of the parallel blade-vortex interaction

A scheme for investigating the parallel blade vortex interaction (BVI) has been designed and tested. The scheme involves setting a vortex generator upstream of a nonlifting rotor so that the vortex interacts with the blade at the forward azimuth. The method has revealed two propagation mechanisms: a type C shock propagation from the leading edge induced by the vortex at high tip speeds, and a rapid but continuous pressure pulse associated with the proximity of the vortex to the leading edge. The latter is thought to be the more important source. The effects of Mach number and vortex proximity are discussed.

Caradonna, F. X.↗

NASA NTRS1984

An experimental investigation of the parallel blade-vortex interaction

Caradonna, F. X.↗

NASA NTRS1955

The Hydrodynamic Characteristics of Modified Rectangular Flat Plates Having Aspect Ratios of 1.00, 0.25, and 0.125 and Operating near a Free Water Surface

Report presents the results of an investigation conducted to determine the hydrodynamic forces and moments acting on modified rectangular flat plates with aspect ratios of 1.00, 0.25, and 0.125 mounted on a single strut and operating at several depths of submersion. A simple method has been developed by modification of Falkner's vortex-lattice theory which enables the prediction of the lift characteristics in unseparated flow at large depths. This method shows good agreement with experimental data from the present tests and with aerodynamic data at all angles investigated for aspect ratios of 1.00 and 0.25 and at angles up to 16 degrees for aspect ratio 0.125. Above 16 degrees for aspect ratio 0.125, the predicted lift proved too high.

Wadlin, Kenneth L↗

NASA NTRS1988

Surface flow visualization of separated flows on the forebody of an F-18 aircraft and wind-tunnel model

A method of in-flight surface flow visualization similar to wind-tunnel-model oil flows is described for cases where photo-chase planes or onboard photography are not practical. This method, used on an F-18 aircraft in flight at high angles of attack, clearly showed surface flow streamlines in the fuselage forebody. Vortex separation and reattachment lines were identified with this method and documented using postflight photography. Surface flow angles measured at the 90 and 270 degrees meridians show excellent agreement with the wind tunnel data for a pointed tangent ogive with an aspect ratio of 3.5. The separation and reattachment line locations were qualitatively similar to the F-18 wind-tunnel-model oil flows but neither the laminar separation bubble nor the boundary-layer transition on the wind tunnel model were evident in the flight surface flows. The separation and reattachment line locations were in fair agreement with the wind tunnel data for the 3.5 ogive. The elliptical forebody shape of the F-18 caused the primary separation lines to move toward the leeward meridian. Little effect of angle of attack on the separation locations was noted for the range reported.

Fisher, David F.↗

NASA NTRS1964

Theoretical and experimental investigation of an unsteady vortex flow <theoretische und experimentelle untersuchungen an einer instationaeren wirbelstroemung<.

Unsteady vortex flow about vertical wedge in shock tube studied by interferometric and theoretical method

SHOCK TUNNEL↗

NASA NTRS2014

DNS of Flows over Periodic Hills using a Discontinuous-Galerkin Spectral-Element Method

Direct numerical simulation (DNS) of turbulent compressible flows is performed using a higher-order space-time discontinuous-Galerkin finite-element method. The numerical scheme is validated by performing DNS of the evolution of the Taylor-Green vortex and turbulent flow in a channel. The higher-order method is shown to provide increased accuracy relative to low-order methods at a given number of degrees of freedom. The turbulent flow over a periodic array of hills in a channel is simulated at Reynolds number 10,595 using an 8th-order scheme in space and a 4th-order scheme in time. These results are validated against previous large eddy simulation (LES) results. A preliminary analysis provides insight into how these detailed simulations can be used to improve Reynoldsaveraged Navier-Stokes (RANS) modeling

DNS↗

NASA NTRS1982

Effects of vortex breakdown on longitudinal and lateral-directional aerodynamics of slender wings by the suction analogy

A semi-empirical method based on the suction analogy is developed to predict longitudinal aerodynamics and lateral-directional characteristics of slender wings at high angles of attack, including effects of vortex breakdown. The latter is based on a correlation parameter derived from the predicted leading-edge suction distribution in the attached flow. Empirical formulas, derived from a least-square analysis of data, for the vortex-breakdown angle of attack at the trailing edge, the progression rate of breakdown points and the vortex lift recovery factor in the breakdown region are given. Comparison of predicted results with data in longitudinal aerodynamics and lateral-directional characteristics for wings exhibiting strong vortex flow shows that the present method is reasonably accurate. Explanation for peculiar lateral-directional characteristics is given.

Lan, C. E.↗

NASA NTRS1986

Recent advances in aerodynamics

Papers are presented on unsteady transonic aerodynamics and aeroelasticity, the unsteady separation phenomenon, and a wind tunnel method for V/STOL testing. Also considered are vortex-edge interactions, jet instability theory, large-scale organized motions in jets and shear layers, and the evolution of adaptive-wall wind tunnels. Other topics include advances in ejector thrust augmentation, multiple jet impingement flowfields, and recent advances in prediction methods for jet-induced effects on V/STOL aircraft.

Krothapalli, A.↗

NASA NTRS1986

Extensions of the concept of suction analogy to prediction of vortex lift effect

Flow field data for a double delta wing at low speed were used to determine the location of a vortex action point. The result was found to be consistent with what was determined for a delta wing. In supersonic flow, the action point location was determined empirically. For a wing with rounded leading edges, an assumption for initial vortex separation was shown to be equivalent to initial leading edge bubble separation for airfoils. A theoretical formulation by the section analogy to determine the delayed vortex separation on a cambered wing with rounded leading edges was presented. The method of suction analogy was further shown to be applicable to predicting the body vortex lift.

Lan, C. E.↗