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NASA NTRS1990

The effect of small streamwise velocity distortion on the boundary layer flow over a thin flat plate with application to boundary layer stability theory

Researchers show how an initially linear spanwise disturbance in the free stream velocity field is amplified by leading edge bluntness effects and ultimately leads to a small amplitude but linear spanwise motion far downstream from the edge. This spanwise motion is imposed on the boundary layer flow and ultimately causes an order-one change in its profile shape. The modified profiles are highly unstable and can support Tollmein-Schlichting wave growth well upstream of the theoretical lower branch of the neutral stability curve for a Blasius boundary layer.

Goldstein, M. E.↗

NASA NTRS1992

Distortion of a flat-plate boundary layer by free-stream vorticity normal to the plate

A nominally uniform flow over a semiinfinite flat plate is considered. The analysis shows how a small streamwise disturbance in the otherwise uniform flow ahead of the plate is amplified by leading-edge bluntness effects and eventually leads to a small-amplitude but nonlinear spanwise motion far downstream from the leading edge of the plate. This spanwise motion is then imposed on the viscous boundary-layer flow at the surface of the plate - causing an order-one change in its profile shape. This ultimately reduces the wall shear stress to zero, causing the boundary layer to undergo a localized separation, which may be characterized as a kind of bursting phenomenon that could be related to the turbulent bursts observed in some flat-plate boundary-layer experiments.

Goldstein, M. E.↗

NASA NTRS1991

Leading-edge receptivity for blunt-nose bodies

This research program investigates boundary-layer receptivity in the leading-edge region for bodies with blunt leading edges. Receptivity theory provides the link between the unsteady distrubance environment in the free stream and the initial amplitudes of the instability waves in the boundary layer. This is a critical problem which must be addressed in order to develop more accurate prediction methods for boundary-layer transition. The first phase of this project examines the effects of leading-edge bluntness and aerodynamic loading for low Mach number flows. In the second phase of the project, the investigation is extended to supersonic Mach numbers. Singular perturbation techniques are utilized to develop an asymptotic theory for high Reynolds numbers. In the first year, the asymptotic theory was developed for leading-edge receptivity in low Mach number flows. The case of a parabolic nose is considered. Substantial progress was made on the Navier-Sotkes computations. Analytical solutions for the steady and unsteady potential flow fields were incorporated into the code, greatly expanding the types of free-stream disturbances that can be considered while also significantly reducing the the computational requirements. The time-stepping algorithm was modified so that the potential flow perturbations induced by the unsteady pressure field are directly introduced throughout the computational domain, avoiding an artificial 'numerical diffusion' of these from the outer boundary. In addition, the start-up process was modified by introducing the transient Stokes wave solution into the downstream boundary conditions.

Kerschen, Edward J.↗

NASA NTRS1992

Leading-edge receptivity for blunt-nose bodies

Boundary-layer receptivity in the leading edge region for bodies with blunt leading edges is investigated in this research program. Receptivity theory provides the link between the unsteady disturbance environment in the freestream and the initial amplitudes of instability waves in the boundary layer. This is a critical problem which must be addressed in order to develop more accurate prediction methods for boundary-layer transition.

Hammerton, P. W.↗

NASA NTRS2010

Receptivity and Transition of Supersonic Boundary Layers Over Swept Wings

The receptivity, stability, and transition of three-dimensional supersonic boundary layers over (1) a swept cylinder, (2) a swept wing with a sharp leading edge, and (3) a swept wing with a blunt leading edge are numerically investigated for a free-stream Mach number of 3. These computations are compared to an earlier experimental and computational study performed by Archambaud et al.1 The steady flow fields with and without roughness elements are obtained by solving the full Navier-Stokes equations. The N-factors computed in this study at the transition onset locations reported in Ref. 1 for flow over the swept cylinder are approximately 16.5 for traveling crossflow disturbances and 9 for stationary disturbances. The N-factors for the traveling crossflow are high based on our past experiences. However, they are comparatively smaller than those reported by Archambaud et al., who found N-factor values in the range of 20 to 25 for traveling disturbances and 13 to 20 for stationary disturbances. Similarly, the N-factors computed in this study for the traveling and stationary disturbances for the flow over the sharp wing are approximately 7 and 2.5, respectively, and for the flow over the blunt wing are 6.5 and 4.8, respectively. Using the envelope method, Archambaud et al. obtained values of approximately 8.0 and 4.0 for the sharp wing case and 16.0 and 12.0 for the blunt wing case.

Balakumar, Ponnampalam↗

NASA NTRS1972

Computation of transonic flow about finite lifting wings.

The results of Murman and Krupp (1971) are used to develop a procedure for computing a transonic flow about a finite lifting wing. A small disturbance equation describing the velocity potential of three-dimensional wings is solved in the process. The procedure is applied to a wing with a symmetric biconvex airfoil section in a zero-incidence subsonic flow, to a supersonic rectangular lifting wing with a sharp leading edge and to a subcritical nonlifting rectangular wing with a blunt leading edge.

Newman, P. A.↗

NASA NTRS1990

Effects of nose bluntness and shock-shock interactions on blunt bodies in viscous hypersonic flows

A numerical study was conducted to investigate the effects of blunt leading edges on the viscous flow field around a hypersonic vehicle such as the proposed National Aero-Space Plane. Attention is focused on two specific regions of the flow field. In the first region, effects of nose bluntness on the forebody flow field are investigated. The second region of the flow considered is around the leading edges of the scramjet inlet. In this region, the interaction of the forebody shock with the shock produced by the blunt leading edges of the inlet compression surfaces is analyzed. Analysis of these flow regions is required to accurately predict the overall flow field as well as to get necessary information on localized zones of high pressure and intense heating. The results for the forebody flow field are discussed first, followed by the results for the shock interaction in the inlet leading edge region.

Singh, D. J.↗

NASA NTRS1961

Surface Pressures and Heat Transfer on Unswept Blunt Plates in Helium at High Mach Numbers

Pressure distributions and local convective heat-transfer coefficients on a flat plate at zero angle of attack were measured in helium. Data were obtained with various amounts of leading-edge bluntness at Mach numbers of 12.5 and 14.7. The pressures on a sharp leading-edged plate were not influenced by the leading edge and were predicted by the first-order, hypersonic, weak-interaction theory. Pressures on blunt plates were correlated by introducing the leading-edge Reynolds number as a parameter. Measured heat-transfer coefficients on the sharp plate agreed with predictions obtained form existing exact solutions for hear transfer across the laminar boundary layer. For the blunt plates a comparison of theory with experiment indicated that more knowledge of the flow field between the sock wave and plate surface is necessary before an adequate prediction of convective heat transfer can be made. Shock-wave shapes for the blun plates at a Mach number 12.5 and zero angle of attack were measured. At distances between 2 and 60 leading-edge thicknesses from the shock vertex, the shock-wave shapes were found to be represented by a modified form of the blast-wave analogy.

Marvin, Joseph G.↗

NASA NTRS1982

Experimental study of delta wing leading-edge devices for drag reduction at high lift

The drag reduction devices selected for evaluation were the fence, slot, pylon-type vortex generator, and sharp leading-edge extension. These devices were tested on a 60 degree flatplate delta (with blunt leading edges) in the Langley Research Center 7- by 10-foot high-speed tunnel at low speed and to angles of attack of 28 degrees. Balance and static pressure measurements were taken. The results indicate that all the devices had significant drag reduction capability and improved longitudinal stability while a slight loss of lift and increased cruise drag occurred.

Johnson, T. D., Jr.↗

NASA NTRS2015

Numerical Analysis of Incipient Separation on 53 Deg Swept Diamond Wing

A systematic analysis of incipient separation and subsequent vortex formation from moderately swept blunt leading edges is presented for a 53 deg swept diamond wing. This work contributes to a collective body of knowledge generated within the NATO/STO AVT-183 Task Group titled 'Reliable Prediction of Separated Flow Onset and Progression for Air and Sea Vehicles'. The objective is to extract insights from the experimentally measured and numerically computed flow fields that might enable turbulence experts to further improve their models for predicting swept blunt leading-edge flow separation. Details of vortex formation are inferred from numerical solutions after establishing a good correlation of the global flow field and surface pressure distributions between wind tunnel measurements and computed flow solutions. From this, significant and sometimes surprising insights into the nature of incipient separation and part-span vortex formation are derived from the wealth of information available in the computational solutions.

Frink, Neal T.↗

NASA NTRS2003

Vortical Flow Prediction Using an Adaptive Unstructured Grid Method

A computational fluid dynamics (CFD) method has been employed to compute vortical flows around slender wing/body configurations. The emphasis of the paper is on the effectiveness of an adaptive grid procedure in "capturing" concentrated vortices generated at sharp edges or flow separation lines of lifting surfaces flying at high angles of attack. The method is based on a tetrahedral unstructured grid technology developed at the NASA Langley Research Center. Two steady-state, subsonic, inviscid and Navier-Stokes flow test cases are presented to demonstrate the applicability of the method for solving practical vortical flow problems. The first test case concerns vortex flow over a simple 65 delta wing with different values of leading-edge radius. Although the geometry is quite simple, it poses a challenging problem for computing vortices originating from blunt leading edges. The second case is that of a more complex fighter configuration. The superiority of the adapted solutions in capturing the vortex flow structure over the conventional unadapted results is demonstrated by comparisons with the wind-tunnel experimental data. The study shows that numerical prediction of vortical flows is highly sensitive to the local grid resolution and that the implementation of grid adaptation is essential when applying CFD methods to such complicated flow problems.

Pirzadeh, Shahyar Z.↗

NASA NTRS2009

Vortical Flow Prediction using an Adaptive Unstructured Grid Method

A computational fluid dynamics (CFD) method has been employed to compute vortical flows around slender wing/body configurations. The emphasis of the paper is on the effectiveness of an adaptive grid procedure in "capturing" concentrated vortices generated at sharp edges or flow separation lines of lifting surfaces flying at high angles of attack. The method is based on a tetrahedral unstructured grid technology developed at the NASA Langley Research Center. Two steady-state, subsonic, inviscid and Navier-Stokes flow test cases are presented to demonstrate the applicability of the method for solving vortical flow problems. The first test case concerns vortex flow over a simple 65 delta wing with different values of leading-edge radius. Although the geometry is quite simple, it poses a challenging problem for computing vortices originating from blunt leading edges. The second case is that of a more complex fighter configuration. The superiority of the adapted solutions in capturing the vortex flow structure over the conventional unadapted results is demonstrated by comparisons with the wind-tunnel experimental data. The study shows that numerical prediction of vortical flows is highly sensitive to the local grid resolution and that the implementation of grid adaptation is essential when applying CFD methods to such complicated flow problems.

Pirzadeh, Shahyar Z.↗

NASA NTRS1962

Comparison of the Effectiveness of Flares With That of Fins for Stabilizing Low Fineness Ratio Bodies at Mach Numbers From 0.6 to 5.8

An experimental and analytical investigation has been made of the effectiveness of flares and of fins for providing aerodynamic stability of low-fineness-ratio bodies near zero angle of attack. Wind-tunnel tests were performed at Mach numbers from 0.6 to 3.8 to measure the normal force, center of pressure, and drag of bodies consisting of conical noses in combination with cylindrical midsections and finned or flared aftersections. This study included an investigation of the effects of nose bluntness, midbody length, flare angle (from 0° to 20°), and fin leading-edge bluntness. The results showed that for the same plan-form area, flares were more effective than fins, but for the same drag, fins were more effective than flares. Flow separation, which occurred ahead of a flare at certain Mach numbers, resulted in an increase in the stabilizing effectiveness of the flare and a decrease in the drag. However, flow separation led to large undesirable shifts in the center of pressure. Blunt fins were found to be more effective than sharp fins, with the vertical as well as the horizontal fins contributing to the effectiveness. Comparisons of the analytical with the experimental results indicated that the analytical method employed gave good estimates of the stabilizing effectiveness of both fins and flares throughout the supersonic Mach number range.

Leland H Jorgensen↗

NASA NTRS1998

Supersonic Leading Edge Receptivity

This paper describes experimental studies of leading edge boundary layer receptivity for imposed stream disturbances. Studies were conducted in the supersonic T-325 facility at ITAM and include data for both sharp and blunt leading edges. The data are in agreement with existing theory and should provide guidance for the development of more complete theories and numerical computations of this phenomena.

Maslov, Anatoly A.↗

NASA NTRS1985

Optimum configuration of high-lift aeromaneuvering orbital transfer vehicles in viscous flow

The results of an analysis to determine the geometrical configuration of an aeroassisted transfer vehicle with a high lift-to-drag ratio (L/D) are described and the constraints imposed on this type of entry vehicle are considered. The aerodynamic characteristics of three configurations, a flat-plate delta wing, a truncated straight cone, and a truncated bent biconic are compared. The effect of viscosity is included in the analysis which examines the rounding of the sharp leading edges. It is shown that, under the constraints of carrying a given volume in the dead air region, the values of L/D are similar for each configuration and that a small blunt leading edge only slightly affects each vehicle's aerodynamic performance, causing less than a 5 percent drop in L/D. The truncated bent biconic is found to be the only configuration that provides the necessary stabilizing moments.

Davies, C. B.↗

NASA NTRS1979

Subsonic wind-tunnel investigation of leading-edge devices on delta wings (data report)

The drag reduction potential of leading edge devices on delta wings at high lift is presented. Fences, pylon type vortex generators, chordwise slots, and sharp plate extensions were tested on wing models with 60 and 74 degree swept blunt leading edges at angles of attack to 32 degrees. Six component balance and leading edge surface pressure measurements were taken with various geometry and spanwise arrangements of the devices. Tables listing the data, without analysis, are presented.

Dhanvada, M. R.↗

NASA NTRS1983

Experimental investigation of leading-edge thrust at supersonic speeds

Wings, designed for leading edge thrust at supersonic speeds, were investigated in the Unitary Plan Wind Tunnel at Mach numbers of 1.60, 1.80, 2.00, 2.16, and 2.36. Experimental data were obtained on a uncambered wing which had three interchangeable leading edges that varied from sharp to blunt. The leading edge thrust concept was evaluated. Results from the investigation showed that leading edge flow separation characteristics of all wings tested agree well with theoretical predictions. The experimental data showed that significant changes in wing leading edge bluntness did not affect the zero lift drag of the uncambered wings.

Wood, R. M.↗

NASA NTRS2007

Lightweight Thermal Protection System for Atmospheric Entry

TUFROC (Toughened Uni-piece Fibrous Reinforced Oxidation-resistant Composite) has been developed as a new thermal protection system (TPS) material for wing leading edge and nose cap applications. The composite withstands temperatures up to 1,970 K, and consists of a toughened, high-temperature surface cap and a low-thermal-conductivity base, and is applicable to both sharp and blunt leading edge vehicles. This extends the possible application of fibrous insulation to the wing leading edge and/or nose cap on a hypersonic vehicle. The lightweight system comprises a treated carbonaceous cap composed of ROCCI (Refractory Oxidation-resistant Ceramic Carbon Insulation), which provides dimensional stability to the outer mold line, while the fibrous base material provides maximum thermal insulation for the vehicle structure.

Stewart, David↗