Investigation of the expansion side of a delta wing at supersonic speed.

Bannink, W.J.; Nebbeling, C.

doi:10.2514/3.6886

Cited by 7 publications

(1 citation statement)

References 4 publications

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“…However, on the thick wing ( Fig. la), this type of flow occurs at much higher MN (> 1- 6) and for a N between about 10° and 40°. The type of flow in region 6 of Fig.…”

Section: ) C E N T R E Line Shock (Not Identfied By Mller and Wood) £mentioning

confidence: 87%

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

Seshadri¹,

Narayan²

1988

Aeronaut. j.

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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.

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“…However, on the thick wing ( Fig. la), this type of flow occurs at much higher MN (> 1- 6) and for a N between about 10° and 40°. The type of flow in region 6 of Fig.…”

Section: ) C E N T R E Line Shock (Not Identfied By Mller and Wood) £mentioning

confidence: 87%