Effect of sweep angle and a half sine wave on roll damping derivative of a Delta Wing
This paper presents the effect of sweep angle on a roll of damping derivative of a delta wing with half-sine wave for an attached shock case in supersonic/hypersonic flow has been studied analytically. The Ghosh Strip theory is replicated. By combining this with the similitude at high-speed flows l...
| Main Authors: | , , , |
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| Format: | Article |
| Language: | English English |
| Published: |
Blue Eyes Intelligence Engineering & Sciences
2019
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| Subjects: | |
| Online Access: | http://irep.iium.edu.my/73973/ http://irep.iium.edu.my/73973/ http://irep.iium.edu.my/73973/ http://irep.iium.edu.my/73973/1/Asha.pdf http://irep.iium.edu.my/73973/7/73973_Effect%20of%20sweep%20angle%20and%20a%20half%20sine%20wave%20on%20roll%20damping%20derivative%20of%20a%20delta%20wing_Scopus.pdf |
| Summary: | This paper presents the effect of sweep angle on a roll of damping derivative of a delta wing with half-sine wave for an attached shock case in supersonic/hypersonic flow has been studied analytically. The Ghosh Strip theory is
replicated. By combining this with the similitude at high-speed flows lead to giving a piston theory. The initial conditions for the applicability of the theory are that the attached wave must be attached with the leading edge of the wing. The results of the present study reveal that with the increments in the sweep angles; it results in continuous decrease in the roll damping derivative, it is also seen that the magnitude of the decrement for lower sweep angle is considerable as compared to the higher values of the sweep angles due to the drastic change in the surface area of the wing. Roll damping derivative progressively increases with the angle of attack; however, with the increase in the inertia level of the flow, it results in the decrement in the damping derivative and later conforms to the Mach number
independence principle. Effect of the leading edge bluntness and viscous effects are neglected. Results have been obtained for the supersonic/hypersonic flow of perfect gases over a wide range of angle of attack, planform area for different Mach numbers. In the present study, attention is on the effect of sweep angle of the wing on roll damping derivative at a different angle of attack and inertia level has been studied. In contemporary theory, Leeward surface is taken along with shock waves attached with the leading edge. |
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