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In the Present paper effect of angle of incidence on Stiffness derivative of a delta wing with Curved leading edges 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 stiffness derivatives at low supersonic to high supersonic Mach numbers. From the results it is seen that with the increase in the Mach number, there is a continuous decrease in the magnitude of stiffness derivatives for all the Mach number tested, however, the magnitude of decrement for different inertia level will differ. It is seen that with the increase in the angle of attack the stiffness 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 was observed that with the shift of the pivot position from the leading edge to the trailing edge, the magnitude of stiffness and the damping derivatives continue to decrease progressively. 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.

Effect of Angle Of Attack on Stiffness Derivative of an Oscillating Supersonic Delta Wing with Curved Leading Edges