Recent Hypersonic Studies of Wings and Bodies

Bertram, M. H.; Henderson, Arthur

doi:10.2514/8.5730

Cited by 25 publications

(6 citation statements)

References 3 publications

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“…At a = 90 deg the flow over the wing is essentially radial from a stagnation point on the centerline and is similar to the flow over a circular disk normal to the free stream (Ref. 7). At angles of attack above 90 deg the flow resembles the subsonic potential flow over a wedge with the wing trailing edge acting as a leading edge for the wedge.…”

Section: Pressure Distributionmentioning

confidence: 81%

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The Aerodynamic Characteristics of a 75-Deg Swept Delta Wing at High Angles of Attack and Mach Numbers of 2 to 8 (U)

Clark¹

1962

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Section: Pressure Distributionmentioning

confidence: 81%

“…7). Therefore, the angle of attack at which shock detachment occurs on a delta wing should be closely approximated by the theoretical cone angle for shock detachment as given in Ref.…”

Section: Shock Wave Geometrymentioning

confidence: 99%

The Aerodynamic Characteristics of a 75-Deg Swept Delta Wing at High Angles of Attack and Mach Numbers of 2 to 8 (U)

Clark¹

1962

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“…4 and the present work. An examination of schlieren photographs taken of the planform of delta wings at zero angle of attack, 4 gives some indication of the nature of the phenomena. These photographs show that the shock, which envelopes the body, has a definite reflex in its profile when viewed in planform.…”

Section: Discussionmentioning

confidence: 99%

Some experiments on delta wings in hypersonic flow.

Barber

1966

AIAA Journal

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An exploratory experimental investigation was made of flat delta wings at angles of attack.The tests were made in the Princeton University Helium Hypersonic Wind Tunnel at Mach number 11.8 and at Reynolds numbers based on leading edge thickness ranging from 4200 to 28,000. The leading edges of the models used in the investigation were swept back at angles of 60° and 70° and had thicknesses of 0.005 in. and 0.030 in. The angle of attack of the models was varied from 0° to 50°. Static pressures and flow directions were measured on the surface of the model. The resulting data were analyzed, and the pressure measurements correlated with the viscous interaction parameter were evaluated using conical flow properties. The effect of leading edge and apex bluntness on static pressures was examined and shown to be present in various degrees throughout the spectrum of angle of attack tested. Blunting the apex and the leading edge are shown to cause unexpected effects on the pressure distribution along the leading edge. The flow in the boundary layer was shown to have strong spanwise tendencies in both the low and high angle-of-attack regimes, and to depend on bluntness. The concepts of spanwise and chordwise strip theory are discussed and applied to the analysis of the experimental data. C r M P RN S T t a A /* 6 X 4L Kd KM x y z (L/Z)) U Nomencl ature root chord Mach number pressure Reynolds number distance from centerline taken normal to the leading edge distance from centerline to the leading edge taken normal to the leading edge temperature thickness of leading edge angle of attack sweep angle coefficient of viscosity local inclination to freestream velocity viscous interaction parameter centerline 8 cota axial dimension lateral dimension in plane of wing lateral dimension normal to plane of wing maximum lift-to-drag ratio freestream velocity out flow angle in plane of the wing body inclination Subscripts LE = leading edge w = wedge 0 = total conditions t = leading edge thickness b = body conditions s = shock c = cone oo = freestream conditions ( ) e -edge of boundary layer N = normal

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“…By Eqs. (16) and (33) In conclusion, we may say that because of moderate angle of attack, the pressure field p eQ is strong in the external flow, and the interaction effects are relatively much smaller than those on a flat plate at zero angle of attack. 15…”

Section: Induced Pressurementioning

confidence: 91%

“…(16) parameters defined by Eq. (23) static pressure Reynolds number based on radial distance and freestream conditions Stanton number, defined by Eq.…”

mentioning

confidence: 99%

Three-dimensional boundary-layer flow over a flat delta wing at a moderate angle of attack.

Polak

1967

AIAA Journal

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This is a study of the three-dimensional boundary layer over the windward side of a flat delta wing in hypersonic flow at a moderate angle of attack. The interaction between the inviscid and the viscous flow regions is analyzed, and its effect is taken into account. The theory is developed for thin, flat delta wings with the shock detached from the leading edges and attached at the apex. The inviscid solution based on the Newtonian-conical flow theory of Messiter-Hida has been matched to the three-dimensional boundary-layer solution over the wing. The transverse velocity component normal to the conical ray is small, and a perturbation method leads to a system of ordinary differential equations. The solution of these equations together with the tangent-cone method is used to study the viscous-inviscid interaction. The results are applied to calculate the streamline pattern, induced pressure, skin drag, and heat transfer over the wing surface. The present theory is in general agreement with the available experimental data.

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Recent Hypersonic Studies of Wings and Bodies

Cited by 25 publications

References 3 publications

The Aerodynamic Characteristics of a 75-Deg Swept Delta Wing at High Angles of Attack and Mach Numbers of 2 to 8 (U)

The Aerodynamic Characteristics of a 75-Deg Swept Delta Wing at High Angles of Attack and Mach Numbers of 2 to 8 (U)

Some experiments on delta wings in hypersonic flow.

Three-dimensional boundary-layer flow over a flat delta wing at a moderate angle of attack.

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