Flow Control Over the Wing of a Fighter-Type Configuration Using Boundary Layer Suction

ICAS 2002 CONGRESS FLOW CONTROL OVER THE WING OF A FIGHTER-TYPE CONFIGURATION USING BOUNDARY LAYER SUCTION M. R. Soltani, T. Khadivi, A. R. Davari Department of Aerospace Engineering, Sharif University of Technology, Tehran, I. R. Iran Keywords: Cropped Delta Wing, Vortex, Flow Control, Suction, Sweep Angle Abstract performance. Various concepts have been used to prevent or delay separation. The objective of this study is to determine the Boundary layer control remains one of the effectiveness of using an active suction most promising avenues delaying for transition technique to delay flow separation over the and separation, thus improving aircraft wing of a fighter type configuration model. A performance. A primary objective of the number of experiments were conducted on a 40 boundary layer control has always been that of degree swept cropped delta wing, similar to the allowing the flow over the wing surface to be High Alpha Research Vehicle (HARV) model, diffused to the trailing edge without massive operating at two subsonic speeds and at low to separation. moderate angles of attack. Both spanwise The second object being pursued very surface pressure distribution and velocity actively at present is delaying the boundary profiles at various angles of attack for suction layer transition phenomenon hence forcing the on and off cases were measured. Smoke and flow to remain laminar instead of turbulent over tufts were used to visualize the flow over the the vehicle [1]. This process will cause wing at various angles of attack. The results substantial reduction in the skin friction drag of indicate formation of a relatively weak vortex the immersed body. over the wing surface at low angle of attack. As There are three methods for stabilizing a alpha increases, this vortex widens, covering a boundary layer and delaying separation: 1) large portion of the wing, disappearing at shaping the surface to provide long runs of moderate angle of attack. Suction affects favourable pressure gradient, 2) providing more surface pressure distribution at low to moderate stable boundary layer through suction, and 3) angle of attack, while its effect diminishes at a providing more stable boundary layer through high angle of attack. It modifies the velocity surface cooling [2]. profile shape near the surface. Boundary layer control as an aerodynamic art has been practiced through the 20th century. 1 Introduction The state of the art as of 20 years ago has been In designing a fighter aircraft, it is most subject of several studies [3]. The last 25 years important to increase its maneuverability at have witnessed a fairly continuous effort at moderate to high angles of attack without developing the technology for laminar flow deteriorating its stability and control criterion. aircrafts using surface suction. The difficulties Highly sweep-back wings have been used for of maintaining the aerodynamic surfaces with these aircrafts mainly to prevent flow separation their numerous suction slots have triggered when operating at high angles of attack. many thoughts and developments [3, 4]. However, depending on the wing sweep angle, In the experimental studies presented the flow field over the wing surface separates at herein a suction system has been used to delay a certain angle of attack, degrading aerodynamic flow separation over the wing of a fighter type configuration operating at subsonic speeds and 2105.1 M. R. Soltani, T. Khadivi, A. R. Davari at low to moderate angles of attack. The wing wind tunnel at the Sharif University of has a 40 degrees sweepback at the leading edge. Technology, Department of Aerospace Both spanwise surface pressure distribution and Engineering. The tunnel is of blow-down type velocity profile at various angles of attack for with an approximately 45x45 cm test section, suction on and off cases have been measured at 100 cm in length. It operates at speeds ranging two different Reynolds numbers. Some flow from 5 to 45 m/s with the diffuser installed [8] visualization tests including smoke visualization and uses 4 screens and a honeycomb in its and tufts were carried out to investigate the flow settling chamber to break down the incoming patterns at various angles of attack. The results free stream vortices. Figure 2 shows the test confirm presence of a relatively weak vortex on section with model installed in it. In figure 3 the wing at angles of attacks of about 6 degrees. variations of the test section turbulence intensity with speed measured by a hot wire anemometer 2 Experimental Apparatus is shown. In this paper the flow field measurement data for clean model are Figure 1 shows the model used for these presented. investigations. It consists of an axisymmetric body with a conical nose and a flat plate cropped delta wing with a leading edge sweep 3 Experimental results angle of 40°. The model is similar to the HARV Figures 4.a-g shows the flow field over a delta model used in various European wind tunnels wing with leading edge sweep of 70°, two swept for force and moment measurements [5-7]. wings with sweeps of Λ=30° and Λ=45°, a However, to authors' knowledge, studies cropped delta wing with ΛLE=30° and the tuft including velocity field, boundary layer and visualization results of the present surface pressure measurements over this investigations. As can be seen, a pair of counter- configuration have not been done yet. An rotating vortices primarily dominates the flow extensive low speed experimental investigation over a delta wing at moderate to high angle of including the aforementioned studies as well as attack. These vortices contain a large amount of canard and its position effects on the surface energy and remain stable for a wide range of static pressure has been conducted at Sharif angles of attack. The surface oil-flow patterns University of Technology, Department of shown over two swept wings, figures 4.b and c Aerospace Engineering. The selected wing has a indicate the existence of a vortex sheet, an chamfered edge and is equipped with small attachment line, and vortex burst phenomenon, holes for measuring surface pressure at various figure 4.b. As the wing sweep angle increases, angles of attack. figure 4.c, Λ=45°, the vortex is seemed to align Static pressure as well as total pressure has itself with the free stream, similar to that of been measured using highly sensitive pressure delta wing, figure 4.a. The vortices formed over transducers. Data for each transducer was these wings, figures 4.b and c, are somehow collected via a multiplexer and transferred to the similar to delta wing vortex but do not contain computer through a 16 bit analog to digital as much energy as the delta wing vortices do. (A/D) board. The 16-bit A/D board was selected This vortex is shed from the wing apex and to increase the system accuracy. Various moves towards the wing tip. From there it sampling rates were performed and finally the becomes parallel to free stream flow. The best one was selected. Tests were performed for movement of these shed vortices towards the clean model and model with two types of wing tip is probably due to the cross flow surface roughness using sand papers located at velocity component over the wing surface. Also the wing leading edge. shown in figure 4 is the low speed flow pattern All tests were conducted for suction on and over a cropped delta wing with leading edge off cases at various angles of attack. These sweep of ΛLE=30°, visualized by oil strakes at experiments were carried out in the subsonic two angles of attack, α=6°, and 12°, figures 4.d 2105.2 FLOW CONTROL OVER THE WING OF A FIGHTER-TYPE CONFIGURATION USING BOUNDARY LAYER SUCTION and e. At low angle of attack, α=6°, the flow is and 6.b), while at angle of attack of 18 degrees, characterized by completely attached flow with the entire wing surface is covered by the a weak tip vortex and a weak leading edge vortical type flow (figures 5.c and 6.c). bubble, figure 4.d. The oil accumulation along Figures 7 and 8 show spanwise static the entire leading edge is an indication of the pressure distribution over the wing surface at leading edge bubble, which is formed by the various angles of attack, α=2°-18°, and at two flow separation from the leading edge of the different stations, X/C=0.6 and 0.8. As seen outer panel and reattaching on the surface again. from these figures, at low angles of attack, α=2° However, for this angle of attack, α=6°, no and α=6°, pressure distribution along the span separation is observed as shown by the straight decreases slightly from the wing root to the attached lines. At 12° angle of attack, figure 4.e, wing tip, indicating potential flow, similar to the the flow pattern is completely different. The oil data shown in figure 4.d. However, for wing tip vortex is still visible, but the attached higher angles of attack, α=10° and α=14°, a flow region is less than that shown in figure 4.d. large portion of the wing surface is dominated Instead there exists a large region of reversed by the low pressure, indicating the presence of flow and a large leading edge and wing tip the vortex. The surface covered by this vortex separation regions. This phenomenon causes a increases as the angle of attack increases from large decrease of lift along with a change in the 10° to 18°. Also, from both figures, figures 7 pitching moment variation. Shown in figure 4 is and 8, note that as alpha increases the point of also the tuft flow visualization result of the minimum pressure moves close to the wing present experiments.

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