Analytical prediction of vortex lift

Saeed

2008

In this study, an experimental and numerical investigation was carried out to obtain lift, drag, and pitching moment data on 65 degree delta and 65/40 degree double-delta wings. The experimental tests were conducted at the King Fahd University of Petroleum and Minerals low-speed wind-tunnel facility, whereas the numerical tests were performed using the commercial computational fluid dynamics software FLUENT. Results from both experiments and numerical predictions were compared to other experimental data found in literature as well as to the theory of Polhamus. The results of comparison of surface pressure coefficient distribution and vortex breakdown location show good agreement with experiments. Overall, the comparison of result shows good agreement between different experimental studies as well as good agreement with the computational fluid dynamics predictions and the theoretical calculations.

Section: Introductionmentioning

confidence: 99%

Experimental and Numerical Investigation of 65 Degree Delta and 65/40 Degree Double-Delta Wings

Saeed

2008

“…This new model consisted of a mixture of two different perfect gases, one for the external flow (perfect gas air), and one for the internal flow (a perfect gas with a reduced ratio of specific heats and molecular weight corresponding to that of a noncombusting simulant gas used in powered hypersonic model wind-tunnel testing). 4 The multizone discretization was used to independently begin the external and internal flows and allow them to mix downstream.…”

Section: General Aerodynamic Simulation Program 20 Attributes Employedmentioning

confidence: 99%

“…4) For the smallest gap, 0.5D, the nacelle inlet is partially inside the wing boundary layer. At a = -2 deg, some boundary layer enters the nacelle and distorts the nacelle exhaust flow symmetry.…”

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confidence: 99%

Analytic prediction of lift for delta wings with partial leading-edge thrust

Traub

1994

ENGINEERING NOTES 3) The wing boundary-layer thickness is equal to the wingnacelle gap of ID at the nacelle trailing edge when a = 0 deg. 4) For the smallest gap, 0.5D, the nacelle inlet is partially inside the wing boundary layer. At a = -2 deg, some boundary layer enters the nacelle and distorts the nacelle exhaust flow symmetry. The amount of wing boundary layer flowing through the nacelle and the associated distortion are lessened with the increase of a. Nomenclature = aspect ratio, b 2 IS = wing span = drag coefficient = lift dependent drag coefficientwing thrust coefficient C Tcff = overall leading-edge thrust coefficient c = local wing chord C R = wing root chord c, = sectional theoretical thrust coefficient kj = induced drag constant k p = potential constant r -leading-edge radius S = wing area a = angle of attack 17 = nondimensional span wise coordinate, 2y/b 7] x = spanwise position at which the leading-edge drag equals the leading-edge thrust r] T = spanwise position at which attainable thrust is equal to theoretical thrust A LE = leading-edge sweep angle A = taper ratio

“…Originally, only the overall characteristics of slender wings could be calculated by Polhamus' suction analogy. Purvis6 suggested an analytical method that could be used to calculate the distribution of the suction along the leading edge. Then the vortex-induced lift of the front part of the axial station, x h is given byThe meaning of the formula can be seen from Ref 6…”

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confidence: 99%

“…Moreover, the Kutta condition must be satisfied at each axial station,^ Eqs (3)(4)(5),. the new vortex strength and position can be determined.The bound vortex lines on the surface can be evaluated by(6) where the 7 is vortex vector, u and v velocity components in the x and y directions, and subscripts u and / represent the upper and lower surfaces, respectively. Downstream from the trailing edge, the trailing-edge vortex sheet must be considered along with the leading-edge vortex sheet.…”

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confidence: 99%

Roll up of strake leading/trailing-edge vortex sheets for double-delta wings

Yin

1985

ConclusionsThe Infuence Function Method previously has been demonstrated to be a revolutionary new tool for the prediction of store loads in aircraft flowfields. The major limitation of the method-the difficulty and expense involved in the calibration process-has been addressed for standard missile configurations by coupling the IFM with the Interference Distributed Loads code. Significant cost reductions have been realized with no compromise in the accuracy of the IFM predictions. an Aircraft," AIAA Paper 83-0266, Jan. 1983.tices and the complex naturfe of vortex sheets of double-delta wings (according to Hummers experiments, three concentrated vortices are probably formed), the author will not employ any artificial smoothing scheme. The main purpose of this Note is to examine the capability of a simple, twodimensional point vortex method to simulate roll up of complicated strake leading/trailing-edge vortex sheets rather than to investigate smoothing schemes. Further, the effect of the interaction of those vortex sheets on the downwash field will be considered. Theoretical ModelThis model is similar to the one used by Sacks et al. 5 According to the slender wing assumption, the original threedimensional steady flow around a wing can be analyzed approximately by a two-dimensional time-dependent flow analogy. The separated free shear layer emanating from the leading edge of the wing is replaced by a finite number of twodimensional point vortices. Each point vortex represents the vorticity shed from the leading edge during a time interval. Sacks' model does not rely on the assumption of conical flow.It is convenient to use complex variables. In terms of conformal mapping, the wing section in the physical plane, X-y^-iz, is mapped onto a circle in the transformed plane, f=£ + /r7, and the resulting problem is reduced to the flow around the circle with a finite number of point vortices outside the circle. For flat-plate wing the transformation is the Joukowski transformation, X= [f+ (a 2 /$) ], where a is the radius of the circle. Therefore, the complex velocity potential,