Hostname: page-component-cd9895bd7-7cvxr Total loading time: 0 Render date: 2024-12-26T17:09:38.677Z Has data issue: false hasContentIssue false
  • English
  • Français

The Flow Over a “High” Aspect Ratio Gothic Wing at Supersonic Speeds

Published online by Cambridge University Press:  07 June 2016

K Yegna Narayan
K Yegna Narayan*
Affiliation:
Cambridge University Engineering Department
Get access

Summary

Results 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.

Type
Research Article
Information
Aeronautical Quarterly , Volume 26 , Issue 3 , August 1975 , pp. 189 - 201
Copyright
Copyright © Royal Aeronautical Society. 1975

Access options

Get access to the full version of this content by using one of the access options below. (Log in options will check for institutional or personal access. Content may require purchase if you do not have access.)

References

1 Yegna, K Narayan Pressure distribution on two wings with curved leading edges at supersonic speeds. ARC R&M 3741, 1974.Google Scholar
2 Yegna, K Narayan Effects of yaw on a gothic wing at supersonic speeds. Aeronautical Quarterly, Vol XXV, p 119, May 1974.Google Scholar
3 Stetson, K F Scaggs, N E Shock detachment from the leading edge of delta wings. Presented at the 36th Semi-Annual Meeting of the Supersonic Tunnel Association, NASA Marshall Space Flight Center, Huntsville, Alabama, 1971.Google Scholar
4 Courant, R Friedrichs, K 0 Supersonic Flow and Shock Waves. Interscience, New York, 1948.Google Scholar
5 Squire, L C Jones, J G A Stanbrook Experimental investigation of the characteristics of some plane and cambered 65 degree delta wings at Mach numbers from 0.7 to 2.0. ARC R&M 3305, 1963.Google Scholar
6 Klunker, E B South, J C Jr Davisy, R M Calculation of non-linear conical flows by the method of lines. NASA TR R-374, 1971.Google Scholar
7 Hui, W H Supersonic and hypersonic flow with attached shock waves over delta wings. Proc. Roy. Soc. London, Series A, Vol 325, p 251, 1971.Google Scholar
8 Squire, L C Calculated pressure distributions and shock shapes on conical wings with attached shock waves. Aeronautical Quarterly, Vol XIX, p 31, February 1968.CrossRefGoogle Scholar
9 Puckett, A E Supersonic wave drag of thin airfoils. Journal of the Aeronautical Sciences, Vol 13, p 475, 1946.CrossRefGoogle Scholar
10 Yegna, K Wings, Narayan with curved leading edges at supersonic speeds. Ph D thesis, Cambridge University Engineering Department, 1974.Google Scholar