Yair Guy

VELOCITY MEASUREMENTS IN A SYNTHETIC JET

Detailed measurements of the velocity field in a synthetic jet emanating from a piezoelectric actuator are conducted. A two-dimensional, steady, turbulent jet is synthesized from the ambient air by periodically deflecting a piezoelectric-driven disk, enclosed in a cavity. The cavity is sealed all around, with only a two-dimensional slot, 0.5 mm wide and 30 mm long, on one side. The actuator is excited with square and sinusoidal waveforms, and the time-dependent velocity is measured with a hotfilm anemometer over a wide range of excitation frequencies and spatial locations. Time-averaged values and useful statistics are then computed from these data. It is found that two distinct regions exist in the flow-field created by the actuator. Close to the slotexit, the mean centerline velocity accelerates to a maximum at a distance of approximately 12 slotwidths. The flow is unsteady, and is dominated by large velocity fluctuations induced by the oscillating membrane. At axial distances larger than approximately 12 slot-widths, the flow pattern changes. At greater distances it appears as a steady jet where the flow is no longer affected by the oscillating membrane. The jet exhibits properties that are indicative of a fully developed, twodimensional, turbulent jet. * Visiting Researcher, on Sabbatical from the Israeli Armament Development Authority ^ Research Engineer, Senior Member AIAA # Assistant Professor, Associate Fellow AIAA This paper is declared a work of the U.S. Government and Is not subject to copyright protection in the United States. Measurements of the centerline velocity, at various excitation frequencies, show that there are two resonance frequencies at which the mean exit velocity is maximized. The two resonance frequencies are 700 Hz and 1160 Hz. When the actuator is excited with a square wave, the corresponding mean velocities are 8.7 m/s and 7.0 m/s. When excited with a sinusoidal wave at the same frequencies, the actuator produces mean velocities of 7.7 m/s and 5.0 m/s. Nomenclature b D d E f ma T t U(t)

Numerical simulation of periodic suction and blowing control of vortex breakdown on a delta wing

A subsonic’ flowfield about a delta wing configuration is generated numerically in order to simulate the interaction between a leading-edge vortex and periodic leading-edge suction and blowing with zero-net mass. The computations are for a delta wing with a 70” sweep angle and a leading-edge slot with sinusoidal blowing and suction. Solutions are obtained on an unstructured mesh system by time integration of the unsteady three-dimensional compressible Navier-Stokes equations with a laminar flow assumption. The Reynolds number is 200,000 which matches wind tunnel experiments. Details of the computations for the delta wing without periodic suction/ blowing are summarized, and the accuracy of numerical results ‘are assessed via grid resolution and studies of numerical parameters related to time-step size. The location of vortex breakdown for both with and without maas injection are computed at an angle of attack, a = 35”, and several periodic blowing/suction parameters to assess its utility in controlling the vortex breakdown phenomenon. Features of the Aowfields are described, and comparisons are made between the respective solutions and companion experimental results.

Parametric Investigation of the Effects of Active Flow Control on the Normal Force of a Delta Wing

The efficacy of active flow control in delaying vortex breakdown and enhancing the lift characteristics of a 70°-sweep delta wing is experimentally investigated in a low-speed wind tunnel at the USAF Academy. Periodic blowing and suction with zero net mass flux is applied at the leading edge of the wing. The pressure distribution over the upper surface of the wing is measured at a freestream velocity of 4.3 m/s, corresponding to a chord Reynolds number of 2.1 x 1 05. A parametric study is conducted, aimed at investigating the effect of periodic flow excitation on the pressure distribution on the upper surface of the wing. In particular, the normal force is computed and optimum values of key control parameters are established. The momentum coefficient of the flow excitation is varied from 0 to 0.004 and the nondimensional frequency is varied from 0 to 3.5. Pressure distribution on the upper surface of the wing is measured at angles of attack of 20” to 40’ and the pressure is integrated to yield the normal force coefficient.

Effects of Canard Shape on the Center of Pressure of a Generic Missile Configuration

An experimental investigation is conducted to study the effects of canard shape on the center of pressure location of a generic canard-wing-body configuration. The aspect ratio, sweep angle and taper ratio of canards having constant planform area are systematically varied. The aspect ratio is varied from 2 to 10. The taper ratio is varied form 0 to 1 and the sweep angle is varied from - 15’ to +45’. The aerodynamic loads on the configurations are measured with a 6-component balance. All tests are conducted in the subsonic wind tunnel at the USAF Academy, at a Mach number of 0.5 and at angles of attack of -5’ to 14’. The efficacy of the various canards in shifting the center of pressure forward is evaluated. It is found that high aspect ratio canards are the most effective in shifting the center of pressure forward and reducing the static stability. At low angles of attack, inline canards having an aspect ratio of ten shift the center of pressure forward by 2.2 body diameters, relative to the case without canards. In comparison, inline canards having an aspect ratio of two and the same planform area shift the center of pressure forward by 0.7 body diameters only. High aspect ratio, interdigitated canards are less effective than the inline canards by 0.4 body diameters.

Experimental Study of a Delta Wing with Upper and Lower Flaps

An experimental investigation is conducted to study the effects of deflectable flaps on the aerodynamic characteristics of a TCP-sweep delta wing. Leading edge upper-surface flaps and lower-surface flaps are investigated. The flaps are deflected downward at angles from 0° to 90°, and the aerodynamic loads are measured with a six component internal strain gage balance. All tests are conducted at a freestream velocity of 12 m/sec, corresponding to a chord Reynolds number of 4.9*105, and at angles of attack of-l°to41°. It is shown that both the upper and the lower flaps alter the aerodynamic characteristics of a delta wing by creating suction force that acts normal to the flap. However, the effect of the upper flaps on the aerodynamic characteristics is much more profound, relative to the effect of the lower flaps. Upper flaps greatly reduce the drag and the lift of the wing, throughout the entire range of deflection angles tested. As a result an increase of approximately 20% in the maximum lift-to drag ratio is obtained, with the upper flaps deflected at 35°, Lower flaps only slightly affect the lift, and increase the drag over most of the angle-of attack tested. As a result, only a small increase of 2% in the maximum lift-to-drag ratio is obtained, with deflection angles lower than 40°.