Optimization of Delta Wing Geometry Parameters for Sonic Boom Mitigation

Abstract

The effects of maximum thickness ratio, maximum camber ratio and wing nose angle on mitigating the sonic -boom ground signature have been investigated earlier for a delta wing . In the present work , f urther studies on sonic boom mitigation are carried out to investigate the effect of the wing dihedral a ngle on same delta wing . Optimal shape design for low sonic -boom ground signature is carried out with constraints of least degradation effect on the aircraft aerodynamic p erformance. The computations of the near -field region are carried out using a modified CFL3D code and t he far -field region computations are carried out using the full -potential code and the Thomas ray code as well. Design -Expert software is used for optimization of individual and combined effects. In the present study the design variable s are the maximum wi ng thickness ratio , the maximum wing camber ratio, the wing nose angle and the wing dihedral angle . T he base aerodynamic component is a delta wing with a biconvex profile with a maximum thickness ratio of 5% . The wing is at zero angle of attack and a Mach number of 2 . The wing is at 40,000 ft altitude. The results show additional decrease in the sonic -boom ground signature with the inclusion of dihedral angle with a small decrease in lift coefficient and a small increase in the drag coefficient. INTRODUCT ION It is well identified that the development of high speed supersonic jet has increasing demand in both the commercial and military sectors. Recently, substantial work has been published to gain deep insight of the sonic -boom noise problem as related wi th supersonic commercial jets. The prime barrier in the development of supersonic jet is the mitigation of sonic -boom ground signature . It is acknowledg ed that shape optimization is the proper approach for t he mitigation of ground sonic boom to as low as 0.3 psf, the objective FAA has set for the Quiet Supersonic Platform (QSP) program. One need to consider the multi -design variable approach to meet the goal set by FAA. For a supersonic aircraft, the near -field shock structure is a complex array of shocks and pressure waves originating from various pa rts of the airframe and engine. In the far -field, these waves coalesce into the ch aracteristic N -shaped wave . The initial rise in pressure, or shock, is due to the coa lescence of various shock waves emanating from the forward components of the aircraft, while the aft pressure rise usually stems from shocks (including recompression shocks) emanating from the aft regions of the aircraft. Research work focused at reducing ground sonic boom should attempt to redu ce the magnitude of both the forward and aft pressure “jumps ”.