Flow Regimes over Delta Wings at Supersonic and Hypersonic Speeds

1976 ◽  
Vol 27 (1) ◽  
pp. 1-14 ◽  
Author(s):  
L C Squire
Keyword(s):  
Shock Layer ◽  
Delta Wing ◽  
Leading Edge ◽  
Flow Regimes ◽  
Lower Surface ◽  
Delta Wings ◽  
Thin Shock Layer ◽  
Wide Range ◽  
Wing Sweep ◽  
Edge Separation

SummaryThis paper concerns the boundaries between flow regimes for sharp-edged delta wings in supersonic flow and the relation of some predictions of thin-shock-layer theory to these boundaries. In particular, it is shown that the theory predicts that the attachment lines on the lower surface of a thin delta wing at supersonic speeds suddenly jump from just inboard of the leading edges to the centre line in certain flight conditions. In general there is close agreement between the conditions for this jump and the flight conditions corresponding to the change-over from attached flow to the leading-edge separation on the upper surface. Since the movement of the attachment lines on the lower surface must change the position of the sonic line and the nature of the expansion around the edge, it is suggested that the two phenomena are directly related. Thus thin-shock-layer theory can be used to establish the boundaries of the various flow regimes for a wide range of Mach number, incidence and wing sweep. The theory can also be used to predict the effects of wing thickness on leading-edge separation, but here the experimental data is very sparse and somewhat contradictory, so the value of the prediction in the case of thickness requires further investigation.

A Note on the Vorticity Distribution on the Surface of Slender Delta Wings with Leading Edge Separation

1961 ◽  
Vol 65 (603) ◽  
pp. 195-198 ◽  
Author(s):  
B. J. Elle ◽  
J. P. Jones
Keyword(s):  
Experimental Evidence ◽  
Delta Wing ◽  
Leading Edge ◽  
Water Tunnel ◽  
Delta Wings ◽  
Vorticity Distribution ◽  
Edge Separation

A description is given of the distribution of vorticity in the surface of thin wings with large leading edge sweep. Although the delta wing is chosen as the basic plan form the deductions are general and applicable to other types of wing. The conclusions are illustrated with experimental evidence from a water tunnel.

Improved calculations of leading-edge separation from slender, thin, delta wings

1968 ◽  
Vol 306 (1484) ◽  
pp. 67-90 ◽  
Keyword(s):  
Boundary Layer ◽  
Leading Edge ◽  
Vortex Sheet ◽  
Lower Surface ◽  
Lateral Position ◽  
Quantitative Agreement ◽  
Delta Wings ◽  
Real Flow ◽  
Edge Separation ◽  
Secondary Separation

In previous calculations (Mangler & Smith 1959) of the vortex-sheet model of leading-edge separation, only qualitative agreement was found with experimental observations. Because the numerical treatment of the model was then necessarily incomplete, it was uncertain how far the lack of quantitative agreement was to be attributed to the limitations of the model. The use of an automatic digital computer has now made it possible to reduce the uncertainties in the calculation to a negligible level. The features of interest in the real flow are more accurately predicted and the remaining discrepancies can be ascribed to the deficiencies in the model. The paper describes the method used to locate the vortex sheet and determine its strength in terms of the two boundary conditions on it; assesses the credibility of the results; and relates them to the observations. It is concluded that the model successfully predicts the observed height of the vortex above the wing, though the predicted lateral position is in error by up to 6% of the semi-span of the wing. This error falls as the incidence increases and is less when transition occurs in the boundary-layer upstream of secondary separation. Normal force is predicted accurately as is the distribution of pressure on the lower surface and the inboard part of the upper surface. The observed suction peak below the vortex changes its character when transition occurs in the boundary-layer upstream of secondary separation. The model predicts the suction peak in the turbulent case fairly well, but it is clear that detailed prediction of the suction peak is not possible by a model which is wholly inviscid.

Hypersonic Flow with Attached Shock Waves over Delta Wings

1970 ◽  
Vol 21 (4) ◽  
pp. 379-399 ◽  
Author(s):  
B. A. Woods
Keyword(s):  
Shock Waves ◽  
Hypersonic Flow ◽  
Shock Layer ◽  
Numerical Solutions ◽  
Delta Wing ◽  
Simple Wave ◽  
Delta Wings ◽  
Conical Flows ◽  
Thin Shock Layer ◽  
Flow Variables

SummaryHypersonic conical flows over delta wings are treated in the thin-shock-layer approximation due to Messiter. The equations are hyperbolic throughout, even in regions where the full equations are elliptic, and have not hitherto been solved for flows with attached shock waves. The concept of the simple wave has been used to construct a class of solutions for such flows; they contain discontinuities in flow variables and shock slope but, for the case of flow over a delta wing with lateral symmetry, agreement with results of numerical solutions of the full equations is good. The method is applied to plane delta wings at yaw, and to wings with anhedral and dihedral. For the flow at the tip of a rectangular wing, it is shown that two distinct solutions may be constructed.

Investigation on lift characteristics of a double-delta wing pitching in various reduced frequencies

2021 ◽  
pp. 095441002199220
Author(s):  
Tianxiang Hu ◽  
Yue Zhao ◽  
Peiqing Liu ◽  
Qiulin Qu ◽  
Hao Guo ◽  
...  
Keyword(s):  
Delta Wing ◽  
Lower Surface ◽  
Sweep Angle ◽  
Delta Wings ◽  
Surface Pressure Distribution ◽  
Reduced Frequency ◽  
Wing Sweep ◽  
The Difference ◽  
The Impact ◽  
Double Delta Wing

The unsteady lift characteristics of a double-delta wing were studied using both experimental and numerical approaches, which were also compared with a single-delta wing with the same main wing sweep angle. It was found that by increasing the reduced frequency of pitching, the hysteresis effect of lift was magnified. Moreover, in the high reduced frequency case k = 0.48, the difference between the lift coefficients of single- and double-delta wings became rather subtle. The wing surface pressure distribution results indicated the flow phenomenon of dramatic lift losses was due to the effect of lower surface suction during the wing being pitched downstroke. It was observed that, as the reduced frequency became sufficiently high, the virtual camber effect induced by pitching could dominate the flow field, which would mitigate the impact of wing geometry on the lift characteristics.

Aerodynamics of a Yawed Blade in Reverse Flow

2012 ◽  
Author(s):  
Rafael Lozano ◽  
Vrishank Raghav ◽  
Narayanan Komerath
Keyword(s):  
Delta Wing ◽  
Reverse Flow ◽  
Leading Edge ◽  
Rotor Blades ◽  
Limiting Factor ◽  
Delta Wings ◽  
Flow Area ◽  
Wide Range ◽  
Sharp Edges ◽  
Finite Wing

The retreating blades of rotorcraft operated at high advance ratios will experience reverse flow through a sector encompassing a wide range of blade azimuth angles. There is a great deal of uncertainty in the blade aerodynamic loads under these conditions. This is a limiting factor when trying to improve the flight speed envelope of helicopters. Previous studies and work have used two-dimensional aerodynamic approaches for the reverse flow area, making the assumption that aerodynamic forces behave similar in magnitude but opposite in direction. There have been no 3-dimensional considerations being taken into account nor was vortex induced lift considered. We hypothesize that the reverse blade flow field includes phenomena similar to the formation of a leading edge vortex on highly-swept, sharp-edges delta wings. An approach is being developed to understand aerodynamic contributions to blade loading beyond linear theory, where vortex-induced lift might be significant. Rotor blades at highly yawed angles relative to the wind can be thought of as very low aspect ratio wings. Since reverse airfoils are thought of as sharp edges, theoretically it should stand that a reverse finite wing at high yawed angles could be considered as a slender delta wing. The main aim of this work is to progress towards testing this above hypothesis. Experimental data is collected from a scaled version of a rotor blade exposed to the reverse flow at various azimuth positions representing the retreating side of the disc, in a 1.07m×1.07m low turbulence wind tunnel.

Hypersonic flow with attached shock waves over plane delta wings

1977 ◽  
Vol 79 (2) ◽  
pp. 361-377 ◽  
Author(s):  
B. A. Woods ◽  
C. B. G. Mcintosh
Keyword(s):  
General Solution ◽  
Shock Waves ◽  
Hypersonic Flow ◽  
Shock Layer ◽  
Delta Wing ◽  
Numerical Calculations ◽  
Delta Wings ◽  
Thin Shock Layer ◽  
New Form

A new form is given for the general solution to the thin-shock-layer equations for the flow over a nearly plane delta wing. Using this, the solution described conjecturally by Hayes & Probstein for symmetrical flow with attached shock waves over a plane delta wing is realized numerically. The flow construction devised for this purpose is applied also to yawed flows. The solutions obtained are found to agree moderately well with the results of numerical calculations from the full equations, but contain a number of anomalous features characteristic of the thin-shock-layer approximation.

The Flow Over a “High” Aspect Ratio Gothic Wing at Supersonic Speeds

1975 ◽  
Vol 26 (3) ◽  
pp. 189-201 ◽  
Author(s):  
K Yegna Narayan
Keyword(s):  
Shock Wave ◽  
Delta Wing ◽  
Lift Coefficient ◽  
Leading Edge ◽  
Two Dimensional ◽  
Delta Wings ◽  
Pressure Drag ◽  
Wide Range ◽  
Detached Shock Wave ◽  
The One

SummaryResults are presented of an experimental investigation on a non-conical wing which supports an attached shock wave over a region of the leading edge near the vertex and a detached shock elsewhere. The shock detachment point is determined from planform schlieren photographs of the flow field and discrepancies are shown to exist between this and the one calculated by applying the oblique shock equations normal to the leading edge. On a physical basis, it is argued that shock detachment has to obey the two-dimensional law normal to the leading edges. From this, and from other measurements on conical wings, it is thought that the planform schlieren technique may not be particularly satisfactory for detecting shock detachment. Surface pressure distributions are presented and are explained in terms of the flow over related delta wings which are identified as a vertex delta wing and a local delta wing. The forces acting on the wing are calculated and are shown to be very close to the two-dimensional wedge values over a wide range of incidence. In particular, it is shown that this wing, compared to one which supports a fully detached shock wave, generates a higher lift/(pressure drag) ratio at a given lift coefficient.

Supersonic Flow Past a Slender Delta Wing. An Experimental Investigation Covering the Incidence Range –5° ≤ α ≤ 50°

1976 ◽  
Vol 27 (2) ◽  
pp. 143-153 ◽  
Author(s):  
I C Richards
Keyword(s):  
Experimental Investigation ◽  
Supersonic Flow ◽  
Surface Pressure ◽  
Shock Layer ◽  
Delta Wing ◽  
Lower Surface ◽  
Flow Visualisation ◽  
Pressure Distributions ◽  
Thin Shock Layer ◽  
Oil Flow

SummaryA detailed survey of a delta wing of 70° sweep has been performed at M = 2.5. The measurements include upper- and lower-surface pressure distributions, schlieren photographs, vapour-screen photographs and surface oil-flow visualisation. The results have been compared with thin-shock-layer theory and various other predictions.

Separated Flow Past a Slender Delta Wing at Incidence

1973 ◽  
Vol 24 (2) ◽  
pp. 120-128 ◽  
Author(s):  
J E Barsby
Keyword(s):  
Delta Wing ◽  
Separated Flow ◽  
Leading Edge ◽  
Vortex Sheet ◽  
Simultaneous Equations ◽  
Delta Wings ◽  
Nested Iteration ◽  
Finite Difference Technique ◽  
Flow Past ◽  
Vortex Systems

SummarySolutions to the problem of separated flow past slender delta wings for moderate values of a suitably defined incidence parameter have been calculated by Smith, using a vortex sheet model. By increasing the accuracy of the finite-difference technique, and by replacing Smith’s original nested iteration procedure, to solve the non-linear simultaneous equations that arise, by a Newton’s method, it is possible to extend the range of the incidence parameter over which solutions can be obtained. Furthermore for sufficiently small values of the incidence parameter, new and unexpected results in the form of vortex systems that originate inboard from the leading edge have been discovered. These new solutions are the only solutions, to the author’s knowledge, of a vortex sheet leaving a smooth surface.Interest has centred upon the shape of the finite vortex sheet, the position of the isolated vortex, and the lift, and variations of these quantities are shown as functions of the incidence parameter. Although no experimental evidence is available, comparisons are made with the simpler Brown and Michael model in which all the vorticity is assumed to be concentrated onto an isolated line vortex. Agreement between these two models becomes very close as the value of the incidence parameter is reduced.

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