Control of shock-induced vortex breakdown in Transonic regime on a slender delta wing body configuration using blowing

Abstract

Modern high speed aircrafts feature highly swept and low aspect ratio wings, mainly to increase the critical Mach number and reduce the wave drag. A special case of swept wings is wings with triangular planform commonly known as delta wings. The delta wings provide advantageous aerodynamic characteristics such as small lift curve slope, non-linear lift and a higher stalling angle due to it’s ability to form two counter rotating leading edge vortices. These vortices are strong and are a source of high energy which enables higher suction on the lee-ward side of wing and results into an increase in lift. However, as the angle of attack is increased beyond certain value, these vortices are affected by changes in the flow behaviour, which causes them to become unstable and breakdown into an incoherent form. This angle of attack is termed as critical angle of attack, beyond which the organised vortical structures are disintegrated leading to loss in lift and wing stall, a phenomenon called as vortex breakdown (VB). The vortex breakdown is detrimental to the aerodynamic characteristics of the wing and can cause instability of the aircraft. Due to this adverse effect, it is important to understand the behaviour of such flows. The behaviour of the flow over slender delta wings under transonic conditions is highly complex. With the occurrence of a number of shocks in the flow, vortex breakdown is abrupt and the overall behaviour is quite different to that for subsonic flow. In fact, the critical angle of attack is significantly lower in transonic speed range compared to low speeds due to interaction of vortex and cross flow/normal shock. The lower critical angle of attack in transonic flow regimes puts serious limitation on overall flight envelope for high speed aircraft. To consider this, the flow over a AGARD-B configuration having a delta wing of 60◦ sweep angle with sharp leading edge, is investigated experimentally in the transonic Mach number range of 0.8 to 0.95 around the critical angle of attack in the range of 10◦ to 15◦. The experimental results are complimented with numerical simulations performed at Mach number of 0.85 at selected angles of attack. Also, a detailed experimental investigation is carried out on a pneumatic control to delay the vortex breakdown. This thesis presents an experimental study of vortical flows and vortex breakdown over an AGARD-B configuration in transonic Mach number regime. The effect of vortex breakdown on overall aerodynamic characteristics and the variation of critical angle of attack with free stream Mach number is determined. The experimental aerodynamic data of AGARD-B is validated with the literature. Few test cases at free stream Mach number of 0.85, are analysed with numerical steady state simulations performed using High Resolution Flow Solver for Unstructured meshes (HiFUN). The numerical simulation results are validated with the present experimental data. A brief literature review of the control strategies adapted for delaying the vortex breakdown is presented. The literature suggests that ‘along the core blowing’ or ‘spanwise blowing’ method for control are found to be effective especially in subsonic Mach number regime. However, evaluation of these methods in transonic Mach number regime is found to be limited in literature. Also, the basic knowledge of a jet interacting with the transonic cross flow and it’s effect on the oncoming transonic flow is found to be limited in literature hence an experimental study of the sonic and supersonic jet interaction with transonic cross flow is carried out on a flat plate. An empirical correlation is suggested to predict the stream wise pressure variation ahead and behind the round jet injection on a flat plate, for a range of transonic Mach numbers, jet exit conditions and momentum ratios. A judicious selection of jet exit condition is made based on these experimental results and is implemented on AGARD-B configuration. The numerical simulations are utilised to evaluate the effectiveness of ‘along the core blowing’ on the vortex breakdown for AGARD-B configuration for few test conditions. Numerical study reveal a number of spanwise shocks present on the leeward side of wing for a baseline case and with the control case. Implementation of the control jet injection seems to modify the shock shape and reduces the severity of the vortex breakdown and improves the suction on leeward side of the wing. The jet used for the ‘along the core blowing’ requires very low levels of blowing (≈ 0.026%) and significantly enhances the lift to drag ratio by about 17 % near critical angle of attack. A detailed experiments are carried out to understand the effect of ‘spanwise’ and ‘along the core blowing’, control jet exit conditions, the control jet injection location etc. using force and moment measurement, pressure sensitive paint technique and unsteady pressure measurements. Analysis of pressure sensitive paint images, unsteady pressure fluctuations is used to identify the vortex breakdown phenomenon and effect of jet injection and is compared with baseline case. The span wise control jet injection from very close to the apex and tangentially along the leeward surface of the wing is found to be enhancing the lift to drag ratio by 4 to 9% depending upon the free stream Mach number. Also, the critical angle of attack is found to be pushed by approximately 2◦ due to the ‘spanwise blowing’ as compared to the baseline case. The primary and secondary vortex are found to be energized due to the span wise control jet injection. It is felt that the oncoming vortex entrains the fluid from the control jet and eventually enhances the axial velocity of vortex core which helps in pushing the vortex breakdown location of primary vortex by about 40% and that of the secondary vortex by 22%. The overall or cumulative coefficient of fluctuating pressure CPrms measured in the rear portion of the wing shows that the control jet injection reduces the overall coefficient of fluctuating pressure significantly by 20% compared with the baseline case indicating a reduction in the buffet loads. It is also observed that the jet exit condition and thus jet momentum coefficient (J) or coefficient of blowing (Cμ) plays an important role in pushing the vortex breakdown (VB) without disturbing the oncoming vortex significantly, especially since the control jet is injected very close to the apex of the wing. The experience gained from these experiments reveal that, for an effective postponement of vortex breakdown, a sonic jet should be blown either ‘Spanwise (SWB)’ or ‘along the core (ACB)’ from very close to the apex of the delta wing and a moderately low momentum coefficient (J) should be chosen for the injection such that the oncoming flow is minimally disturbed.