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1266 PATEL ET AL. Swept wings of low aspect ratio are commonly used on high-speed aircraft because of their favorable wave-drag characteristics. The LEV is the main feature of the flow over swept wings that provide lift for flight control at high angles of attack. At low angles of attack and lower speeds, however, the aerodynamic behavior of swept wings is vastly different from that of the high-aspect ratio wings. The performance of swept wings outside the high-speed, high-alpha envelope is crucial, because the mission roles of modern aircraft require them to operate at low-speed and low-alpha conditions during different portions in flight (e.g., takeoff, landing, etc.). The formation of the LEV and subsequent vortex breakdown (VBD) phenomena over a swept wing are highly influenced by a number of parameters including angle of attack, leading-edge design, and adverse pressure gradients, which present unique challenges in controlling the vehicle dynamics at different flow conditions. For example, at low angles of attack, the flow remains attached to the surface and the location of the (weak) VBD is usually downstream within the wake of the wing. As the angle of attack increases, the strength of the LEV increases and the location of the VBD begins to move forward. The VBD phenomenon is usually associated with a loss in vortex lift, which has been shown to cause changes in the lift, drag, and pitching moments of the swept air vehicle [21–23]. At large angles of attack, the upper wing surfaces show the presence of complex vortex systems that dominate the leeward flowfield and cause the wingtip separations [24]. In the past decade, several researchers have employed flow control methods to control the LEV and VBD phenomena for improved aerodynamics of a swept wing. For example, Moeller and Rediniotis [25] demonstrated control of the pitching moment of a 60-deg swept- delta-wing model at high angles of attack using a series of surface- mounted pneumatic vortex control actuators. Control was achieved by altering the vortex breakdown phenomena that affected the chordwise lift distribution over the wing, ultimately resulting in an induced pitching moment. Amitay et al. [26] reported an experimental study on the use of synthetic jet actuators on a 1301 UCAV design (nicknamed Stingray). The design of Stingray [26] and the present 1303 UCAV share some similarity in that the leading- edge sweep angle is approximately 50 deg, leading to similar three- dimensional flow patterns over the wing. Amitay et al. [26] showed that at conditions in which the flow was normally separated from the leading edge, between 14- and 24-deg angles of attack, the zero-mass jets were able to produce significant forces and moments on the vehicle. Visser et al. [27–29] employed steady spanwise blowing to control leading-edge vortex breakdown and asymmetric roll- moment conditions. In a more recent effort, a computational study on the aerodynamic performance of a 1303 UCAV design for different leading-edge designs was reported by Zhang et al. [30] The effects of three leading- edge designs [a basic profile, a rounded leading edge (similar to the one used in our study), and a sharp leading edge] were investigated using the NPARC code at a Mach number of 0.25 and at angles of attack ranging from 5 to 20 deg. It was found that there were only minor differences among pressure distributions with the three configurations for both the computed and experimental data. The predicted pressure distributions compared favorably with the wind- tunnel measurements for all regions except near the wingtip, for which the computations did not consistently predict the separations. At small angles of attack, flowfield studies showed attached, smooth, and well-behaved flow. In general, the flying wing aircraft that have been developed and successfully flown rely on multiple control surfaces distributed across the wing to provide control moments for trim and maneuvering. Each control surface is essentially a trailing-edge flap that when deflected, changes the lift, drag, and pitching moment over that portion of the wing. By suitably arranging multiple flaps across the wing, one can create moments to pitch, roll, or yaw the wing and moments to trim the wing at a particular flight condition. The ultimate objective of the present work is to demonstrate hingeless flight control with limited or no use of conventional control surfaces. This paper presents results using plasma actuators placed near the leading edge to provide control at high-angle-of-attack ( 15 to 35 deg) flight conditions. The present work complements subsequent demonstrations on the 1303 UCAV planform for roll control using leading-edge plasma actuators [31] and for lift control at low angle of attack ( to 24 deg) using wind-side plasma actuators [32]. II. Experimental Setup The UCAV planform used in this study is based on a 1303 design with varying cross sections, a 47-deg leading-edge sweep, and a 30- deg trailing-edge sweep (shown in Fig. 2a). The design was originally developed by the U.S. Air Force Research Laboratory (AFRL) in conjunction with Boeing Phantom Works and was recently used as a benchmark for a joint computational fluid dynamics code-validation effort by The Technical Cooperation Program (TTCP). TTCP involved a consortium of governmental interests in five countries to study the performance predictions of various Boeing/AFRL 1303 UCAV configurations with different leading edges [33–37]. In the present work, a 1303 configuration with a (relatively) blunt leading edge was used. The same configuration was later used in other plasma flow control demonstrations [31,32]. Photographs of the different full- and half-span models of the scaled 1303 UCAV used for wind-tunnel tests are shown in Figs. 2– 4. Laser-smoke flow visualization experiments were conducted on a 2.31%-scale full-span model of the UCAV (shown in Fig. 2b), to capture off-surface flowfield information. Fluorescent-oil flow visualization and force-balance experiments were conducted on a 4.16%-scale, 0.4-m root chord, 0.34-m span, half-span model (shown in Fig. 2c). Tests were also conducted on a 4.16%-scale half- span model (shown in Fig. 2d) with traditional control surfaces, flap, and split ailerons, to quantify improvements in the overall control authority and the operational envelope of the wing using the plasma actuators. The half-span model has a root chord of 16 in. (40.64 cm) and a half-span dimension of 13.375 in. (13.97 cm). The models were cast from a numerically machined two-piece aluminum mold. The casting material was a mixture of epoxy and microglass beads that resulted in a very rigid model that accurately duplicated the mold shape. Wind-tunnel experiments were conducted for angles of attack ranging from 0 to 35 deg. Many of the tests were performed from 0 to 25 deg, however, additional tests were later conducted for angles of attack up to 35 deg. Lift and drag measurements on the half- span models were conducted in the 0.42 m (1.39 ft) square by 1.8 m (6 ft) long, cross-sectional, low-speed wind tunnel at the University of Notre Dame. All experiments were conducted at a chord Reynolds number Rec of 4:12 105 based on the root chord, which corresponds to a Mach number of 0.045 and freestream velocity U1 of 15 m=s. The tunnel consists of a removable inlet with a series of 12 screens, followed by a 24:1 contraction that attaches to the test section. The turbulence level in the test section, u0 =U1 , was approximately 0.08%. The test section is equipped with a clear Plexiglas side wall that allows optical access to view the model. The back wall of the test section has removable panels that allow access into the test section. The half-span models were mounted vertically on the support sting of a lift–drag force balance that was mounted on the top of the test section. The model was suspended below a splitter plate that was attached to the ceiling of the test section. The splitter plate was designed to produce a two-dimensional flow with a thin boundary layer leading up to the model. A hole in the ceiling splitter plate accommodated the sting supporting the model. Wiring for the plasma actuator also entered through this hole. This hole was aligned with the support sting so that it would not interfere with angular positioning of the model when setting different angles of attack. A stepper motor on the force balance drove the angular position of the support sting. Its motion was controlled by the data acquisition computer through software; with this, the angular position was repeatable to 0:005 deg. The force balance consists of independent lift and drag platforms. The lift platform was supported on the drag platform by two vertical plates that flex only in the lift direction. The drag platform wasPDF Image | Plasma Actuators for Hingeless Aerodynamic Control of an Unmanned Air Vehicle
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