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.

Method for Providing Warning of the Onset of Buffeting, Stalling and Other Undesirable Effects of Flow Separation

1958 ◽  
Vol 62 (573) ◽  
pp. 674-676 ◽  
Author(s):  
D. W Holder ◽  
H. H. Pearcey
Keyword(s):  
Shock Wave ◽  
Recent Work ◽  
Flow Separation ◽  
Static Pressure ◽  
Separated Flow ◽  
Leading Edge ◽  
Alternative Methods ◽  
Pressure Measurements ◽  
Aircraft Wing ◽  
Trailing Edge

A Method is described for providing warning of the onset of undesirable effects produced by flow separation on an aircraft wing. It is based on static pressure measurements near the trailing edge and appears to have advantages over alternative methods.Recent work has suggested that certain effects of flow separation such as buffeting, aileron buzz and undesirable changes in loading, occur when a “ bubble ” of separated flow originating at the leading edge of a wing, or at a shock wave on its surface, first becomes sufficiently large to affect the flow at the trailing edge.

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.

Characteristics of a separated flow past a semicircular leading-edge airfoil model under different imposed pressure gradient

2021 ◽  
pp. 095441002110212
Author(s):  
K Anand ◽  
KT Ganesh
Keyword(s):  
Boundary Layer ◽  
Particle Image Velocimetry ◽  
Pressure Gradient ◽  
Particle Image ◽  
Separation Bubble ◽  
Separated Flow ◽  
Leading Edge ◽  
Field Measurements ◽  
Bench Mark ◽  
Image Velocimetry

The effect of pressure gradient on a separated boundary layer past the leading edge of an airfoil model is studied experimentally using electronically scanned pressure (ESP) and particle image velocimetry (PIV) for a Reynolds number ( Re) of 25,000, based on leading-edge diameter ( D). The features of the boundary layer in the region of separation and its development past the reattachment location are examined for three cases of β (−30°, 0°, and +30°). The bubble parameters such as the onset of separation and transition and the reattachment location are identified from the averaged data obtained from pressure and velocity measurements. Surface pressure measurements obtained from ESP show a surge in wall static pressure for β = −30° (flap deflected up), while it goes down for β = +30° (flap deflected down) compared to the fundamental case, β = 0°. Particle image velocimetry results show that the roll up of the shear layer past the onset of separation is early for β = +30°, owing to higher amplification of background disturbances compared to β = 0° and −30°. Downstream to transition location, the instantaneous field measurements reveal a stretched, disoriented, and at instances bigger vortices for β = +30°, whereas a regular, periodically shed vortices, keeping their identity past the reattachment location, is observed for β = 0° and −30°. Above all, this study presents a new insight on the features of a separation bubble receiving a disturbance from the downstream end of the model, and these results may serve as a bench mark for future studies over an airfoil under similar environment.

Visualization of separation and reattachment in an incident shock-induced interaction

2021 ◽  
pp. 095441002098349
Author(s):  
Yun Jiao ◽  
Chengpeng Wang
Keyword(s):  
Numerical Simulation ◽  
Shock Wave ◽  
Shear Stress ◽  
Separation Bubble ◽  
Incident Shock ◽  
Incident Shock Wave ◽  
Bottom Wall ◽  
Surface Shear Stress ◽  
Surface Shear ◽  
Oil Flow

An experimental study is conducted on the qualitative visualization of the flow field in separation and reattachment flows induced by an incident shock interaction by several techniques including shear-sensitive liquid crystal coating (SSLCC), oil flow, schlieren, and numerical simulation. The incident shock wave is generated by a wedge in a Mach 2.7 duct flow, where the strength of the interaction is varied from weak to moderate by changing the angle of attack α of the wedge from 8° and 10° to 12°. The stagnation pressure upstream was set to approximately 607.9 kPa. The SSLCC technique was used to visualize the surface flow characteristics and analyze the surface shear stress fields induced by the initial incident shock wave over the bottom wall and sidewall experimentally which resolution is 3500 × 200 pixels, and the numerical simulation was also performed as the supplement for a clearer understanding to the flow field. As a result, surface shear stress over the bottom wall was visualized qualitatively by SSLCC images, and flow features such as separation/reattachment and the variations of position/size of separation bubble with wedge angle were successfully distinguished. Furthermore, analysis of shear stress trend over the bottom wall by a hue value curve indicated that the relative magnitude of shear stress increased significantly downstream of the separation bubble compared with that upstream. The variation trend of shear stress was consistent with the numerical simulation results, and the error of separation position was less than 2 mm. Finally, the three-dimensional schematic of incident shock-induced interaction has been achieved by qualitative summary by multiple techniques, including SSLCC, oil flow, schlieren, and numerical simulation.

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.

scholarly journals Shock-induced separated flows on the lee surface of delta wings

1987 ◽  
Vol 91 (903) ◽  
pp. 128-141 ◽  
Author(s):  
S. N. Seshadri ◽  
K. Y. Narayan
Keyword(s):  
Pressure Rise ◽  
Leading Edge ◽  
Experimental Results ◽  
Separated Flows ◽  
Delta Wings ◽  
Number Range ◽  
Pressure Distributions ◽  
Boundary Layer Interaction ◽  
Oil Flow ◽  
Attached Flow

Experiments were conducted to study shock-induced separated flows on the lee surface of delta wings with sharp leading edge at supersonic speeds. Two sets of delta wings of different thickness (10° and 25° normal angle), each with leading edge sweep angles varying from 45° to 70°, were tested. The measurements, carried out in a Mach number range from 1.4 to 3.0, included oil flow visualisations (on both sets of wings) and static pressure distributions (on the thicker wings only). Using the test results, some features of shock-induced separated flows, including in particular the boundary between this type of flow and fully attached flow, have been determined. The experimental results indicate that this boundary does not seem to show any significant dependence on wing thickness within the limit of thicknesses tested. It is shown that this boundary can be predicted for thin delta wings using a well known criterion for incipient separation in a glancing shock wave boundary layer interaction, namely that a pressure rise of 1.5 is required across the shock. Comparison of the predicted boundary with experimental results (from oil flow visualisations) shows good agreement.

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.

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.

Predicting Transitional Separation Bubbles

2004 ◽  
Vol 127 (3) ◽  
pp. 497-501
Author(s):  
John A. Redford ◽  
Mark W. Johnson
Keyword(s):  
Boundary Layer ◽  
Separation Bubble ◽  
Separated Flow ◽  
Leading Edge ◽  
Transition Process ◽  
Test Cases ◽  
Transition Model ◽  
Flow Transition ◽  
Free Stream Turbulence ◽  
Attached Flow

This paper describes the modifications made to a successful attached flow transition model to produce a model capable of predicting both attached and separated flow transition. This transition model is used in combination with the Fluent CFD software, which is used to compute the flow around the blade assuming that it remains entirely laminar. The transition model then determines the start of transition location and the development of the intermittency. These intermittency values weight the laminar and turbulent boundary layer profiles to obtain the resulting transitional boundary layer parameters. The ERCOFTAC T3L test cases are used to validate the predictions. The T3L blade is a flat plate with a semi-circular leading edge, which results in the formation of a separation bubble the length of which is strongly dependent on the transition process. Predictions were performed for five T3L test cases for differing free-stream turbulence levels and Reynolds numbers. For the majority of these test cases the measurements were accurately predicted.

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