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Publication Title | Aerodynamic Control Using Windward-Surface Plasma Actuators on a Separation Ramp

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Search Completed | Title | Aerodynamic Control Using Windward-Surface Plasma Actuators on a Separation Ramp
Original File Name Searched: JA-2007-Windward Plasma UCAV Lift Control.pdf | Google It | Yahoo | Bing



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1890 LOPERA ET AL.
plasma slats for autonomous sensing and control of wing stall by Patel et al. [9], plasma optimized airfoil by Corke et al. [10], and plasma wings for hingeless flight control of a UAV by Patel et al. [1]. A more detailed background on the physics and behavior of plasma actuators are provided by Enloe et al. [11,12], and an overview of some of the recent developments of the SDBD actuator are presented by Corke et al. [13].
A majority of past work on the use of a plasma actuator as a flow control device has focused on flat-plate and two-dimensional geometries. A recent study by Patel et al. [1] has shown experimental evidence on the use of a plasma actuator for flight control of a three- dimensional geometry: a 47-deg-sweep 1303 UAV. Although the control demonstration was good at high angles of attack (between 􏰃 􏰇 15 and 25 deg), no effects were produced at low angles of attack (􏰃 < 14 deg). Results from a flow visualization study conducted on the1303UAVmodelsuggestthatatbelow􏰃􏰴15 deg,theleading- edge vortices are intact and thus too strong to be affected by plasma. Above􏰃􏰴15 deg,thevorticesbreakdownandsotheleewardflow is essentially that of a stalled wing, which enables the effect of plasma to be more pronounced.
Additionally, results from flow visualization experiments conducted in a water tunnel revealed that the windward surface (pressure side) provided a relatively simpler (two-dimensional) flow structure that is fully attached [1]. There appears to be little, if any, crossflow component on the windward surface of the wing section. This is in strong contrast to the leeside (suction side) flow, which exhibited a strong crossflow component. These results suggest that the windward-side (pressure-side) flow is potentially more receptive to control than the suction side.
Because of the predominant complex 3-D flow structure on the suction side, which is difficult to control, and the simpler 2-D flow structure on the pressure side, seemingly controllable via simple plasma actuator geometries, the current study was undertaken to investigate the effect of a novel windward-ramp plasma flow control concept for achieving lift control at low angles of attack. The purpose of the present work is to investigate the behavior of steady and unsteady plasma-induced flow over different separation-ramp angles on the windward surface for lift control. The concept of manipulating the flow past a separation ramp to influence the aerodynamic performance of a control surface using a flow control device is not new; however, the application of this technique on the windward surface to achieve lift control has not been examined before.
The main objectives of the present work were to 1) assess the performance of plasma actuators in conjunction with a windward separation ramp for providing lift control at low angles of attack and 2) improve the control effectiveness by optimizing the backward ramp angle and actuator parameters such as unsteady (modulation) frequency and duty cycle.
Fig. 1 Schematic of the UAV wing tested. Dimensions are in millimeters.
Force balance
Wind direction
Fig. 2
(Not to scale)
High-precision rotary stage
Wind-tunnel ceiling
Wind-tunnel floor
Wind-tunnel experimental setup.
A.
II. Experimental Setup UAV Test Model
Two tempered-glass sidewalls and a large Plexiglas window on the ceiling provided convenient access for flow visualization from different viewing angles. The flow in the test section was uniform, with a turbulence level of less than 0.2% outside of the wall boundary layers.
For every angle of attack, 30,000 samples were taken at a sampling rate of 2000 samples per second. Initial tests were composed of one data sample, and later, three data samples were collected and ensemble-averaged for each test case. Lift forces were measured, and the accuracy of lift data was verified by repeat measurements. The average standard deviation in the lift coefficient was 0.00496.
B. Plasma Actuators
The SDBD plasma actuator configuration consists of two electrodes that are separated by a dielectric material. One of the electrodes is usually exposed to the surrounding air and the other is fully encapsulated by a dielectric material. When a large ac potential, hereafter referred to as ac carrier frequency, is applied to the electrodes at sufficiently high amplitude levels, the air in the region of the largest potential ionizes and plasma is produced. The ionization typically occurs at the edge of the electrode that is exposed to the air and spreads out over the area projected by the covered electrode, directing momentum into the surrounding air.
In the present study, two strips of plasma actuators were placed immediately upstream of the separation ramp: one in the inboard section and the other in the outboard section. The actuators were fabricated using two 0.05-mm-thick copper electrodes, which were separated by two layers of 0.1-mm-thick Kapton film. Both inboard and outboard actuators were operated in a similar configuration. The electrodes were arranged in an asymmetric configuration, as shown in Fig. 3. The two electrodes were overlapped by a small amount, approximately 1 mm, to ensure uniform plasma was formed in the
The design of the 1303 UAV test model is described in detail in the paper by Patel et al. [1]. The UAV model has a leading-edge sweep of 47 deg with varying cross sections and a trailing-edge sweep of 􏰊30deg, as shown in Fig. 1. The present experiments were conducted on a 4.16%-scale half-span UAV model with a 0.4-m root chord and 0.34-m span. Wind-tunnel experiments were conducted at the University of Toledo’s low-speed closed-return wind tunnel with a 0:9 􏰉 0:9 m (3 􏰉 3 ft) test section. Experiments were conducted at a chord Reynolds number of 4:33 􏰉 105, based on the mean chord c of 0.2 m and freestream velocity of 15 m=s for angles of attack ranging from 􏰈10 to 26 deg at 2-deg increments.
A force balance mounted on top of the wind-tunnel ceiling was used for the present experiments. The wing model was connected to a high-precision 495-Series 3-in. rotary stage that was controlled using a Newport 855C programmable controller and a remote controller for model positioning. Using this system, the measurement accuracy of the angle of attack was 􏰊0:001 deg. A schematic of the experimental setup is shown in Fig. 2.

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