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.

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.

Lift Force of Delta Wings

1990 ◽  
Vol 43 (9) ◽  
pp. 209-221 ◽  
Author(s):  
Mario Lee ◽  
Chih-Ming Ho
Keyword(s):  
Lift Force ◽  
Delta Wing ◽  
Angle Of Attack ◽  
Leading Edge ◽  
Surface Flux ◽  
Experimental Results ◽  
Geometrical Parameters ◽  
Two Dimensional ◽  
Delta Wings ◽  
Vorticity Balance

On a delta wing, the separation vorticies can be stationary due to the balance of the vorticity surface flux and the axial convection along the swept leading edge. These stationary vortices keep the wing from losing lift. A highly swept delta wing reaches the maximum lift at an angle of attack of about 40°, which is more than twice as high as that of a two-dimensional airfoil. In this paper, the experimental results of lift forces for delta wings are reviewed from the perspective of fundamental vorticity balance. The effects of different operational and geometrical parameters on the performance of delta wings are surveyed.

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.

Effect of Shapes and Turbulent Inlet Flow to Vortices on Delta Wings

2019 ◽  
Vol 889 ◽  
pp. 434-439
Author(s):  
Ngoc Khanh Tran ◽  
Van Khang Nguyen ◽  
Phu Khanh Nguyen ◽  
Thi Kim Dung Hoang ◽  
Van Quang Dao
Keyword(s):  
Turbulence Intensity ◽  
Delta Wing ◽  
Lift Coefficient ◽  
Pressure Coefficient ◽  
Leading Edge ◽  
Aerodynamic Characteristics ◽  
Delta Wings ◽  
Ansys Fluent ◽  
Inlet Flow ◽  
Stall Angle

This paper aims to estimate the effect of turbulent inlet flow to vortices on Delta wing with four different turbulence intensity from 0.5% to 15% and the effect of taper ratios to aerodynamic characteristics of Delta wings with four taper ratios: 0, 0.1, 0.2, 0.3, 0.4, 0.5, 0.6, 0.7. The main purpose of this paper is to find out the formation, development, and breakdown of vortices on Delta wings when changing taper ratios and turbulence intensity thence determining the center of vortices with the range of attack angles from 5o to 40o in low velocities about 2.5 m/s. This research uses Delta wing models with a 40o swept-back leading edge, the root chord length 150 mm, and a thickness 5 mm. The problem is simulated by using ANSYS fluent and experiment in the subsonic wind tunnel to compare and validate results. The Delta wing models are meshed by using ICEM to improve the mesh quality and using the turbulence model for low Reynolds number flows Transition SST (4 equations) to calculate aerodynamic characteristics such as lift coefficient, drag coefficient, pressure coefficient... find the paths which connect centers of the vortices, and show the contours of pressures and velocities to evaluate the change of centers of the vortices. The results showed that the two vortices grow up and tend to move inward when the attack angle increase, the vortices are broken strongly in high attack angles, the aerodynamic quality of Delta wings change insignificantly when changing turbulent intensity at inlet. This research also carried out that the stall angle increase when increasing the taper ratio.

Possible types of flow on lee-surface of delta wings at supersonic speeds

1988 ◽  
Vol 92 (915) ◽  
pp. 185-199 ◽  
Author(s):  
S. N. Seshadri ◽  
K. Y. Narayan
Keyword(s):  
Shock Wave ◽  
Static Pressure ◽  
Delta Wing ◽  
Separation Bubble ◽  
Separated Flow ◽  
Leading Edge ◽  
Delta Wings ◽  
Oil Flow ◽  
Normal Angle ◽  
Secondary Separation

SummaryExperiments were conducted to study the types of flow that occur on the lee surface of delta wings at supersonic speeds. Two sets of flat topped delta wings of different thickness (wedges with 10° and 25° normal angle respectively), each with leading edge sweep angles of 45°, 50°, 60° and 70°, were tested. The measurements, carried out at Mach numbers of 1·4, 1·6, 1·8, 2·0, 2·5 and 3·0, included oil flow visualisations (on both sets of wings) and static pressure distributions (on the thicker wing only). In addition, a 60° sweptback delta wing with a normal angle of 40° was also tested. The tests on this wing included both oil flow visualisations and static pressure measurements. From these and other existing measurements, the leeside flows have been classified into nine distinct types, namely (i) leading edge separated flow with secondary separation, (ii) leading edge separated flow with secondary and tertiary separation, (iii) leading edge separated flow with a shock wave beneath the primary vortex, (iv) leading edge separated flow with shock-induced secondary separation, (v) fully attached flow, (vi) flow attached at the leading edge with inboard shock-induced separation, (vii) mixed type of flow, (viii) flow with a leading edge separation bubble and (ix) leading edge separated flow with a shock wave lying on the lee surface in between the leading edge vortices. These types of flow have been displayed in a plane of Mach number and angle of attack normal to the leading edge. The experimental results indicate that increasing wing thickness has no qualitative effect on the types of flow observed but does shift the boundaries between some of the types of flow.

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.

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.

Effects of leading-edge bevel angle on the aerodynamic forces of a non-slender 50° delta wing

2005 ◽  
Vol 109 (1098) ◽  
pp. 403-407 ◽  
Author(s):  
J. J. Wang ◽  
S. F. Lu
Keyword(s):  
Delta Wing ◽  
Lift Coefficient ◽  
Leading Edge ◽  
Relative Thickness ◽  
Aerodynamic Performances ◽  
Stall Angle ◽  
Drag Ratio ◽  
Low Speed Wind Tunnel ◽  
Lift To Drag Ratio ◽  
Blunt Leading Edge

Abstract The aerodynamic performances of a non-slender 50° delta wing with various leading-edge bevels were measured in a low speed wind tunnel. It is found that the delta wing with leading-edge bevelled leeward can improve the maximum lift coefficient and maximum lift to drag ratio, and the stall angle of the wing is also delayed. In comparison with the blunt leading-edge wing, the increment of maximum lift to drag ratio is 200%, 98% and 100% for the wings with relative thickness t/c = 2%, t/c = 6.7% and t/c = 10%, respectively.

Optimization of Aerodynamic Parameters of Cropped Delta Wing with Fence at Sonic Mach Number

2019 ◽  
Vol 16 (2) ◽  
pp. 403-409
Author(s):  
M. P. Arun ◽  
M. Satheesh ◽  
Edwin Raja J. Dhas
Keyword(s):  
Shock Wave ◽  
Boundary Condition ◽  
Mach Number ◽  
Sound Speed ◽  
Delta Wing ◽  
Leading Edge ◽  
Supersonic Speed ◽  
Sweep Angle ◽  
Ansys Fluent ◽  
Aerodynamic Parameters

Manufacturing and maintaining different aircraft fleet leads to various purposes, which consumes more money as well as man power. Solution to this, nations that are leading in the field of aeronautics are performing much research and development works on new aircraft designs that could do the operations those were done by varied aircrafts. The foremost benefit of this delta wing is, along the huge rearward sweep angle, the wing’s leading edge would not contact the boundary of shock wave. Further, the boundary is produced at the fuselage nose due to the speed of aircraft approaches and also goes beyond the transonic to supersonic speed. Further, rearward sweep angle greatly worse the airspeed: wings under normal condition to leading edge, so permits the aircraft to fly at great transonic, subsonic, or supersonic speed, whereas the over wing speed is kept to minimal range than that of the sound speed. The cropped delta wing with fence has analysed in three cases: Fences at 3/4th distance from the centre, with fences at half distance from the centre and with fences at the centre. Further, the delta wing that cropped is exported to ANSYS FLUENT V14.0 software and analysed by making the boundary condition settings like sonic Mach number of flow over wing along with the angle of attack.

Function Airfoil and its Pressure Distribution and Lift Coefficient Calculation

2013 ◽  
Vol 328 ◽  
pp. 351-356 ◽  
Author(s):  
Hai Bo Jiang ◽  
Zhong Qing Cheng ◽  
Yun Peng Zhao
Keyword(s):  
Pressure Distribution ◽  
Distribution Curve ◽  
Integration Method ◽  
Lift Coefficient ◽  
Analytical Formula ◽  
Leading Edge ◽  
Independent Variable ◽  
The One ◽  
The Impact ◽  
The Relationship

To study the impact of an airfoil shape on performance, a curve expression of airfoil shape was proposed, the analytical formula for the pressure distribution of flow around the airfoil was derived, and the pressure distribution view around airfoil with azimuth as the independent variable was put forward, which can clearly express the details of the pressure distribution curve on airfoil leading edge. Used both the pressure distribution integration method and Blasius theorem, the lift coefficient calculation formulas of ideal fluid flow around the airfoil were derived respectively, and the same results were obtained. Studies have shown that the shape of an airfoil can be expressed by a function, and various types of shapes can be easily obtained by adjusting the constant value in the expression; The pressure distribution and lift coefficient can be calculated by analytical method; For function airfoil, lift coefficient formula could be derived by two methods, and could be verified with each other. The one-to-one relationship exists between the constant values in the airfoil function, airfoil shapes and airfoil performances, and the relationship expression was given in this paper.

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