Mehul P. Patel

PLASMA FLAPS AND SLATS: AN APPLICATION OF WEAKLY-IONIZED PLASMA ACTUATORS

The experimental validation of an application of weakly-ionized plasma actuators for improved aerodynamic performance of multi-element wings and wings with movable control surfaces is presented. Two spanwise arrays of plasma actuators, configured to produce a wall-jet effect, were applied on the suction surface of a two-dimensional NACA 0015 wing model, one at the leading edge and the other near the trailing edge to mimic the effects of a wing leading-edge slat and a trailing-edge flap, respectively. Flow control tests were conducted at chord Reynolds numbers, corrected for blockage, of 0.217 x 10 6 and 0.307 x 10 6 in a low-speed wind tunnel at the University of Notre Dame. The leading-edge-separation control resulted in an increase in both the maximum lift coefficient and the stall angle of attack and a lift-to-drag improvement of as much as 340%. An optimum frequency was found to exist for unsteady excitation of the leading-edge separation. Under this condition, the power to the actuator was estimated to be only 2 W. The trailing-edge actuator was found to produce the same effect as a plain trailing-edge flap. This included a uniform shift at all angles of attack of the lift coefficient and a shift toward higher lift coefficients of the drag bucket. In addition, there was a slight decrease in the minimum drag coefficient. The obvious advantages of this approach are its simplicity, as there are no moving parts, and its lack of hinge gaps, which add drag. An example of their use as ailerons for roll control produces a comparable roll moment coefficient to a sample general aviation aircraft.

Scaling Effects of an Aerodynamic Plasma Actuator

We present experimental results to yield insight into the scalability and control effectiveness of single-dielectricbarrier-discharge plasma actuators for leading-edge separation control on airfoils. The parameters investigated are chord Reynolds number, Mach number, leading-edge radius, actuator amplitude, and unsteady frequency. This includes chord Reynolds numbers up to 1:0 � 106 and a maximum freestream speed of 60 m=s corresponding to a Mach number of 0.176. The main objective of this work is to examine the voltage requirements for the plasma actuators to reattach the flow at the leading edge of airfoils at poststall angles of attack for a range of flow parameters in order to establish scaling between laboratory and full-flight conditions. For the full range of conditions, an optimum unsteady actuator frequency f is found to minimize the actuator voltage needed to reattach the flow, such that F� � fLsep=U1 � 1. At the optimum frequencies, the minimum voltage required to reattach the flow is weakly dependent on chord Reynolds number and strongly dependent on the poststall angle of attack and leading-edge radius. The results indicate that the voltage required to reattach the flow scales as the square of the leading-edge radius.

Plasma actuators for hingeless aerodynamic control of an unmanned air vehicle

The use of dielectric barrier discharge plasma actuators for hingeless flow control over a 47-deg 1303 unmanned combat air vehicle wing is described. Control was implemented at the wing leading edge to provide longitudinal control without the use of hinged control surfaces. Wind-tunnel tests were conducted at a chord Reynolds number of 4.12 x 105 and angles of attack ranging from 15 to 35 deg to evaluate the performance of leading-edge plasma actuators for hingeless flow control. Operated in an unsteady mode, the actuators were used to alter the flowfield over the lee-side wing to modify the aerodynamic lift and drag forces on the vehicle. Multiple configurations of the plasma actuator were tested on the lee side and wind side of the wing leading edge to affect the wing aerodynamics. Data acquisition included force-balance measurements, laser fluorescence, and surface flow visualizations. Flow visualization tests mainly focused on understanding the vortex phenomena over the baseline uncontrolled wing to aid in identifying optimal locations for plasma actuators for effective flow manipulation. Force-balance results show considerable changes in the lift and drag characteristics of the wing for the plasma-controlled cases compared with the baseline cases. When compared with the conventional traditional trailing-edge devices, the plasma actuators demonstrate a significant improvement in the control authority in the 15- to 35-deg angle-of-attack range, thereby extending the operational flight envelope of the wing. The study demonstrates the technical feasibility of a plasma wing concept for hingeless flight control of air vehicles, in particular, vehicles with highly swept wings and at high angles of attack flight conditions in which conventional flaps and ailerons are ineffective.

Autonomous Sensing and Control of Wing Stall Using a Smart Plasma Slat

DOI: 10.2514/1.24057 The concept of a self-governing smart plasma slat for active sense and control of flow separation and incipient wing stall is presented. The smart plasma slat design involves the use of an aerodynamic plasma actuator on the leading edge of a two-dimensional NACA 0015 airfoil in a manner that mimics the effect of a movable leading-edge slat of a conventional high-lift system. The self-governing system uses a single high-bandwidth pressure sensor and a feedback controller to operate the actuator in an autonomous mode with a primary function to sense and control incipient flow separation at the wing leading edge and to delay incipient stall. Two feedback control techniques are investigated. Wind tunnel experiments demonstrate that the aerodynamic effects of a smart actuator are consistent with the previously tested open-loop actuator, in that stall hysteresis is eliminated, stall angle is delayed by 7 deg, and a significant improvement in the lift-to-drag ratio is achieved over a wide range of angles of attack. These feedback control approaches provide a means to further reduce power requirements for an unsteady plasma actuator for practical air vehicle applications. The smart plasma slat concept is well suited for the design of low-drag, quiet, highlift systems for fixed-wing aircraft and rotorcraft.

Electric Circuit Model for Aerodynamic Plasma Actuator

The aerodynamic plasma actuator has shown considerable promise as a o w control device in dieren t applications. It has been shown previously that the lumped-element circuit model correctly describes the behavior of the aerodynamic plasma actuator. To incorporate this model into the Navier-Stokes solver, it was modied to include the spatial distribution of the discharge within the plasma. To model the behavior of the single dielectric barrier discharge aerodynamic plasma actuator, we represent it as a network of electric circuit elements. The electric circuit consists of N elementary subcircuits, each representing a small physical domain with nite width and length. Each subcircuit consists of the air capacitor, dielectric capacitor, plasma resistive element, and diodes which govern the presence of the plasma. The results of the simulation are compared to the experimental data of the plasma spatial distribution obtained with a photomultiplier tube.

Plasma Flaps and Plasma Slats: An Application of Weakly-Ionized Plasma Actuators

The experimental validation of an application of weakly-ionized plasma actuators for improved aerodynamic performance of multi-element wings and wings with movable control surfaces is presented. Two spanwise arrays of plasma actuators, configured to produce a wall-jet effect, were applied on the suction surface of a two-dimensional NACA 0015 wing model, one at the leading edge and the other near the trailing edge to mimic the effects of a wing leading-edge slat and a trailing-edge flap, respectively. Flow control tests were conducted at chord Reynolds numbers, corrected for blockage, of 0.217 x 10 6 and 0.307 x 10 6 in a low-speed wind tunnel at the University of Notre Dame. The leading-edge-separation control resulted in an increase in both the maximum lift coefficient and the stall angle of attack and a lift-to-drag improvement of as much as 340%. An optimum frequency was found to exist for unsteady excitation of the leading-edge separation. Under this condition, the power to the actuator was estimated to be only 2 W. The trailing-edge actuator was found to produce the same effect as a plain trailing-edge flap. This included a uniform shift at all angles of attack of the lift coefficient and a shift toward higher lift coefficients of the drag bucket. In addition, there was a slight decrease in the minimum drag coefficient. The obvious advantages of this approach are its simplicity, as there are no moving parts, and its lack of hinge gaps, which add drag. An example of their use as ailerons for roll control produces a comparable roll moment coefficient to a sample general aviation aircraft.

Experimental Investigations of Reconfigurable Porosity for Aerodynamic Control

Experimental results to demonstrate the application of an innovative, low-power, flow control technique – reconfigurable porosity – for aerodynamic control is presented. The control is based on selective actuation and de-actuation of discrete holes that form unique porous-patterns to enable small amounts of mass transfer in-and-out of the surface. Since the technique relies on flow transpiration caused due to natural pressure loading around the aerodynamic surface, power is required only for the control of hole-patterns, and not for plumbing, as in the case of suction or blowing. Low-power, MEMS-based microvalves can be used to form discrete hole-patterns to further minimize the total power consumption. Transpiration of air at certain optimal locations on the aerodynamic surface is hypothesized to energize the boundary layer, forcing a delay in flow separation. Controlled threedimensional perturbations caused by patterned-porosity can be rapidly re-configured to adapt to the changing flow conditions so as to continually operate in an optimal configuration for improved effectiveness. To assess the control performance, low-speed experiments are conducted on two aerodynamic models, a multi-wedged supersonic wing model and a complete projectile model (with two four fins arranged in cruciform (+) configuration). Force data as well as flow visualization results are presented to illustrate the patterned-porosity concept. Quantitative results show that control forces of varying magnitude can be generated using different porous patterns for a range of alpha conditions. This control technique has demonstrated the ability to generate adequate levels of control forces for course correction and maneuvering of air vehicles.

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

Wind-tunnel experiments were cone ucted on a 47-deg sweep, scaled 1303 unmanned air vehicle model to assess the performance of an innovative windwiird-surface plasma actuator design for flight control at low angles of attack. Control was implemented by altering the flow past an aft separation ramp on the windward side using a single dielectric barrier discharge plasma actuator. The influence of ramp-expansion angles (20, 30, and 40 deg) on the plasma actuators ability to affect flow separation and aerodynamic lift was examined. Both steady and unsteady actuations of the plasma actuator were examined, and their effects were captured using lift measurements and flow visualizations. Results reveal that the plasma actuator effects are highly dependent on the ramp angle and actuator parameters such as duty cycle and modulation frequency. The actuators produced significant shifts in the lift curve, up to 25% for the most effective ramp angles of 20 and 30 deg, in the 0-20-deg a range. Flow visualization results, confirmed that the plasma actuator causes the flow to reattach over a region downstream of the separation ramp. For all ramp cases examined, the unsteady (pulsed) actuator was more effective than the steady actuator in controlling flow separation and influencing the aerodynamic lift. The aerodynamic effect of plasma actuators was found to be highly dependent on the ramp angle and the separation strength over the ramp. Significant control forces were obtained using windward-surface plasma actuators and, indirectly, these control forces can be implemented to generate substantial control moments for maneuvering air vehicles.

Deployable Flow Effectors for Phantom Yaw Control of Missiles at High Alpha

A high alpha phantom yaw control system for enhanced missile maneuverability and stabilization control has been developed. Open and closed-loop experiments on a fin-less 3:1 tangent ogive missile model were conducted to quantify the control effectiveness of the high alpha phantom yaw control system. The flow control system utilizes co-located actuators and sensors modules, force balance data, and a closed-loop controller. The co-located actuators and sensors modules incorporate eight deployable flow effectors and eight corresponding dynamic pressure sensors located circumferentially near the tip of the missile nose cone. Deployable flow effectors are active micro-vortex generators that control and manipulate pressure distribution along the forebody to produce significant side forces and yawing moments for missile control. Significant side forces caused by crossflow separation and natural baseline asymmetries were observed between 40° and 60° alpha. Deployable flow effectors were efficacious between 40° to 55° alpha in generating large side forces that cancelled the baseline flow asymmetry and producing yawing moments on either side of the slender body. Results demonstrate that flow effectors can be used to achieve a wide spectrum of control forces and to modulate the side forces around the missile forebody for desired effect. Dynamic test results showed that the closed-loop controller was successfully able to control the yawing moment on the missile during sweeps from 0° to 60° alpha at various sweep rates. Closed-loop experiments demonstrated the control system’s ability to maintain a desired side force corresponding to zero, left and right yaw. Nomenclature α (alpha) angle of attack, deg Red freestream Reynolds number based on model diameter, ρUod/μ U0 freestream velocity μ absolute viscosity d missile model diameter β angle of sideslip, deg θ DFE radial location on nose cone, deg Cn yawing-moment coefficient, MZ/qSrefd CY Side-force coefficient, FY/qSref MZ yawing moment FY side force Sref cross-sectional area of cylindrical portion of missile model, πd/4 PID Proportional Integral Derivative Kp proportional gain of the PID control law Ki integral gain of the PID control law Kd derivative gain of the PID control law

Active Boattailing and Aerodynamic Control Fins for Maneuvering Weapons

An active flow control concept utilizing miniature deployable structures for advanced weapons control is presented. The ultimate goal is to provide pitch and yaw control to weapons (slender bodies) that operate at low angles of attack, where the baseline control is primarily provided by tail-fins. In this study, miniature-spoilers, integrated on the weapon boattail and fins, act as flow control devices to provide aerodynamic control. The targeted application is the U. S. Army’s 105 mm smart cargo projectile. Primary objectives of the present work are to enhance the range and end-game maneuvering of the projectile. The basic idea is to replace the traditional hinged-control surfaces (tail-fins) with Aero Control Fins (fins integrated with miniature-spoilers) to generate the required control authority for projectile maneuvering. Spoilers are also tested on the boattail for further enhancing the total control authority, as well as, for providing control on a finless projectile. Spoiler configurations of varying height, length, and position on the aeroboattail and the aero fins are examined. Low-speed experiments are conducted on a 70% scale projectile model at Mach 0.1. Results demonstrate significant control-force generation and effective force modulation by varying spoiler-configurations for the 0-20 deg alpha range. This technology offers an active, compact, light-weight, flight control system for aerodynamic maneuvering, and is applicable to all types of slender-bodied weapons including missiles, projectiles, and munitions.