Effect of Angle of Attack on Stability Derivatives of a Delta Wing with Straight Leading Edge in Supersonic Flow

In the Present paper effect of angle of incidence on pitching derivatives of a delta wing with Straight 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. This combines with similitude to give a piston theory. From the results it is seen that with the increase in the Mach number, there is a decrement of stiffness as well as the damping derivatives in pitch 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 both stiffness and damping derivatives increase 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 h = 0.0, 0.2, 0.4, 0.6, 0.8, 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 both the stability derivatives were decreasing progressively with the pivot position. 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, and secondary wave reflections are neglected.