Investigation of cross flow shocks on delta wings in supersonic flow

1979 ◽  
Author(s):  
M. SICLARI
Keyword(s):  
Supersonic Flow ◽  
Cross Flow ◽  
Delta Wings

Investigation of Crossflow Shocks on Delta Wings in Supersonic Flow

AIAA Journal ◽  
1980 ◽  
Vol 18 (1) ◽  
pp. 85-93 ◽  
Author(s):  
Michael J. Siclari
Keyword(s):  
Supersonic Flow ◽  
Delta Wings

The Caret Wing at Certain Off-Design Conditions

1972 ◽  
Vol 23 (4) ◽  
pp. 263-275 ◽  
Author(s):  
W H Hui
Keyword(s):  
Supersonic Flow ◽  
Central Region ◽  
Cross Flow ◽  
Lower Surface ◽  
Uniform Flow ◽  
Pressure Curve ◽  
Angle Of Incidence ◽  
Internal Shock ◽  
The Cross ◽  
Leading Edges

SummaryA unified theory is given of hypersonic and supersonic flow over the lower surface of a caret wing at certain off-design conditions when the bow shock is attached to the leading edges of the wing and when there exists no internal shock. The flow field on the lower surface of a caret wing consists of uniform flow regions near the leading edges, where the cross-flow is supersonic, and a non-uniform flow in the central region, where the cross-flow is subsonic. The basic assumption is that the flow in the central region differs slightly from the two-dimensional supersonic flow over a flat plate at the same angle of incidence as that of the lower ridge of the wing. Based on this assumption, a first-order perturbation flow is first calculated and then strained and corrected so that it matches the uniform flow which is obtained exactly. Slope discontinuities of the pressure curve are found at the cross-flow sonic line. Numerical examples and comparisons with previous theories and experiments are included.

Pressure fields over hypersonic wing-bodies at moderate incidence

1973 ◽  
Vol 59 (4) ◽  
pp. 673-691 ◽  
Author(s):  
N. Malmuth
Keyword(s):  
Cross Sections ◽  
Three Dimensional ◽  
Series Representation ◽  
Cross Flow ◽  
Leading Edge ◽  
Delta Wings ◽  
Flow Stagnation ◽  
Arbitrary Body ◽  
Area Rule ◽  
Mach Cones

Delta wings with conically subsonic cones-bodies mounted on their compressive side are analysed in the hypersonic small disturbance regime. The weakly three-dimensional conditions associated with slender parabolic Mach cones are used to validate a linearized rotational approximation of the flow field. A combined integral–series representation is obtained for the pressure distribution between the wing-body and shock wave for arbitrary body cross-sections, and is specialized to give analytical formulae for arbitrary-order polynomial transversal contours. Numerical results are presented for wedge, parabolic and higher order cross-sections illustrating the dominant character of the cross-flow stagnation singularity associated with sharp wing-body intersections having a finite slope discontinuity. It is shown that the pressure has a logarithmic infinity at this secondary leading edge, as in corresponding Prandtl–Glauert irrotational flows. The relation of this finding to Lighthill's theorem on cross-stream vorticity is discussed. Other features of the pressure field are considered with particular emphasis on their relationship to a recently derived area rule for such configurations, and possibilities for favourable interference.

Comparison of Euler and Navier-Stokes solutions for vortex flow over a delta wing

1988 ◽  
Vol 92 (914) ◽  
pp. 145-153 ◽  
Author(s):  
A. Rizzi ◽  
B. Müller
Keyword(s):  
Large Scale ◽  
Stokes Equations ◽  
Delta Wing ◽  
Cross Flow ◽  
Leading Edge ◽  
Navier Stokes ◽  
Delta Wings ◽  
Navier Stokes Equations ◽  
Starting Location ◽  
Secondary Vortices

Summary A numerical method has been developed recently to solve the Navier-Stokes equations for laminar compressible flow around delta wings. A large-scale Navier-Stokes solution on a mesh of 129 × 49 × 65 points for transonic flow Mx = 0·85, α = 10° and Rex = 2·38 × 106 around a 65° swept delta wing with round leading edge is presented and compared with a correspondingly large-scale Euler solution. The viscous results reveal the presence of primary, secondary, and even tertiary vortices. The starting location of the primary vortex is seen to be quite different in the two solutions. In the viscous solution it starts at the wing apex, but in the Euler results it starts about one quarter chord downstream. The secondary reparations are also different, due to the up-lifting of the boundary layer in the viscous results, but to a cross-flow shock in the Euler computation. Comparison with experiment shows that the interaction between the primary and secondary vortices in the Navier-Stokes computation is obtained correctly and that these results are a more realistic simulation than the one given by the Euler equations.

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