International OPEN ACCESS Journal Of Modern Engineering Research (IJMER) | IJMER | ISSN: 2249–6645 | www.ijmer.com | Vol. 4 | Iss.10| Oct. 2014 | 15| Effect of Sweep Angle on Rolling Moment Derivative of an Oscillating Supersonic/Hypersonic Delta Wing Asha Crasta 1 , S. A. Khan 2 , Antony A. J 3 1 Sr. Assistant Professor, Mangalore Institute of Technology and Engineering, Moodabidri, Karnataka, India, 2 Principal, Z. H. College of Engineering & Technology, AMU, Aligarh, UP, India, 3 Vice Principal, Sahyadri College of Engineering and Management, Mangalore, India. I. Introduction The analysis of hypersonic and supersonic flow over flat deltas with straight leading edge over a wide incidence range is of current interest since the desire for high speed, maneuverability and efficiency has been dominating the evolution of high performance military aircrafts. The knowledge of aerodynamic load and stability for such types is a need for calculating simple but reasonably accurate methods for parametric calculations facilitating the design process. The computation of dynamic stability for these shapes at high incidence which is likely to occur during the course of reentry or maneuver is of current interest. Usually the shock waves are very strong when descending and they can either be detached or attached. The theories for steady delta wings in supersonic/hypersonic flow with shock wave attached were given by Pike [1] and Hui [2]. Carrier [3] and Hui [4] gave exact solutions for 2-D flow in the case of an oscillating wedge and for an oscillating flat plate were given by Hui [5], which is valid uniformly for all supersonic Mach numbers and wedge angles or angles of attack with attached shock wave. Hui [5] also calculated pressure on the compression side of a flat delta. The importance of dynamic stability at large incidence during re-entry or maneuver has been pointed out by Orlik-Ruckemann [6]. The shock attached relatively high aspect ratio delta is often preferred for its high lift to drag ratio. Hui and Hemdan [7] have studied the unsteady shock detached case in the context of thin shock layer theory. Liu and Hui [8] have extended Hui‟s [5] theory to a shock attached delta wing in pitch. Light hill [9] has developed a “Piston Theory” for oscillating airfoils at high Mach numbers. A parameter δ is introduced, which is a measure of maximum inclination angle of Mach wave in the flow field. It is assumed that M ∞ δ is less than or equal to unity (i.e. M ∞ δ ≤ 1) and is of the order of maximum deflection of a streamline. Light hill [9] likened the 2-D unsteady problem to that of a gas flow in a tube driven by a piston and termed it “Piston Analogy”. Ghosh [10] has developed a large incidence 2-D hypersonic similitude and piston theory. It includes Light hill‟s [9] and Mile‟s [11] piston theories. Ghosh and Mistry [12] have applied this theory of order of ¢ 2 where ¢ is the angle between the attached shock and the plane approximating the windward surface. For a plane surface, ¢ is the angle between the shock and the body. The only additional restriction compared to small disturbance theory is that the Mach number downstream of the bow shock is not less than 2.5. Abstract: In the Present paper effect of sweep angle on roll of damping derivative of a delta wing with straight leading edges for an attached shock case in supersonic/hypersonic flow has been studied analytically. A Strip theory is used in which strips at different span wise location are independent. This combines with similitude which leads to give a piston theory. The Present theory is valid for attached shock case only. The results of the present study reveals that with the increase in the sweep angle; 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 very large as compared to the higher values of the sweep angles due to the drastic change in the plan form area. Roll damping derivative progressively increases with the angle of attack, however, with the increase in the Mach number it results in the decrement in the damping derivative and later conforms to the Mach number principle. Effects of wave reflection, leading edge bluntness, and viscosity have not been taken into account. Results have been obtained for supersonic/hypersonic flow of perfect gases over a wide range of angle of attack, plan form area, and the Mach number. Keywords: delta wing, Hypersonic flow, Piston theory, Rolling derivative,Supersonic Flow, sweep angle.