Effect of angle of attack on damping derivatives of delta wing with full sine wave curved leading edge
In the Present paper effect of angle of incidence on Damping derivative of a delta wing with Curved leading edges (for a full sine wave) for attached shock case in Supersonic Flow has been studied. A Strip theory is used in which strips at different span wise location are independent of each other....
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| Main Authors: | , |
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| Format: | Article |
| Language: | English |
| Published: |
RS Publication
2015
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| Subjects: | |
| Online Access: | http://irep.iium.edu.my/49928/1/paper21.pdf http://irep.iium.edu.my/49928/ http://rspublication.com/ijeted/ijeted_archive.htm |
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| Summary: | In the Present paper effect of angle of incidence on Damping derivative of a delta wing with Curved leading edges (for a full sine wave) for attached shock case in Supersonic Flow has been studied. A Strip theory is used in which strips at different span wise location are independent of each other. This combines with similitude to give a piston theory which gives closed form solutions for damping derivatives at low to high supersonic Mach numbers. From the results it is seen that with the increase in the Mach number, there is a progressive decrease in the magnitude of damping derivatives for all the Mach numbers of the present studies; however, the decrease in the magnitude is variable at different inertia level. It is seen that with the increase in the angle of attack the damping derivative increases linearly, nevertheless, this linear behavior limit themselves for different Mach numbers. For Mach number M = 2, this limiting value of validity is fifteen degrees, for Mach 2.5 & 3, it is twenty five degrees, whereas, for Mach 3.5 & 4 it becomes thirty five degrees, when these stability derivatives were considered at various pivot positions; namely at h = 0.0, 0.4, 0.6, and 1.0. After scanning the results it is observed that with the shift of the pivot position from the leading edge to the trailing edge, the magnitude of the damping derivatives continue to decrease throughout. Results have been obtained for supersonic flow of perfect gases over a wide range of angle of attack and Mach number. The effect of real gas, leading edge bluntness of the wing, shock motion, and secondary wave reflections are neglected. |
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