Control of unsteady separated flow associated with the dynamic pitching of airfoils

Although studies have been done to understand the dependence of parameters for the occurrence of deep stall, studies to control the flow for sustaining lift for a longer time has been little. To sustain the lift for a longer time, an understanding of the development of the flow over the airfoil is essential. Studies at high speed are required to study how the flow behavior is dictated by the effects of compressibility. When the airfoil is pitched up in ramp motion or during the upstroke of an oscillatory cycle, the flow development on the upper surface of the airfoil and the formation of the vortex dictates the increase in lift behavior. Vortex shedding past the training edge decreases the lift. It is not clear what is the mechanism associated with the unsteady separation and vortex formation in present unsteady environment. To develop any flow control device, to suppress the vortex formation or delay separation, it is important that this mechanism be properly understood. The research activities directed toward understanding these questions are presented and the results are summarized

Steady pressure measurements in the strap-on booster interference Region of 1/20 scale aslv configuration

Wind tunnel studies were carried out to obtain pressure distribution in the strap-on booster interference region of 1/20th scale Augmented Satellite Launch Vehicle model configuration. Tests were done in the 1 .2m tunnel at NAL in the Mach number range of 0 .5 to 2.5 for the clean configuration as well as with spring housing attachments on the strap-on boosters. Both the model configurations with the boosters strapped on to the core vehicle in the horizontal plane (pitch) and in the vertical plane (yaw) were tested for incidences at 0, 4 and -4 deg. In addition pressure measurements were also done on the core vehicle alone at Mach numbers 2.1, 2.5 and 3.0 for 0, +4 degree incidences. The test Reynolds number was varied from 0.7 to 1.3 millions based on the maximum diameter of the model. The pressure distribution showed significant interference effects of boosters on the core vehicle. It is observed that the positive pressure peak associated with flow compression at the flare junction increases with increase in Mach number. In the pitch plane the normal force distribution remains positive along the core vehicle whereas in the yaw plane it is of less magnitude

Effectiveness of vortex plate on the aerodynamic characteristics of MiG-21M aircraft model

Wind tunnel studies were carried out to determine the effect of vortex plate fixed at the leading edge of wing, on the aerodynamic characteristics of MIG-21M aircraft model. Tests Reynolds number based on the mean aerodynamic chord varied from 3.9 to 6.0 million in the Mach number range of 0.5 to 1.8. Angle of attack varied from -4 to 20 degrees. The results have demonstrated that vortex plate can be used as an effective device in increasing the lift to drag ratio of the aircraft model at moderate to large lift coefficients at subsonic and transonic speeds. The effectiveness is observed to decrease with increase in Mach number

Surface pressure measurements in the strap-on boosters interference region of 1/180 scale PSLV configuration at supersonic speeds

Pressure measurements were made on the core vehicle in the strap-on interference region of a 1/100 scale PSLV configuration at supersonic Mach numbers of 2.1, 2.5 and 3.0 and incidences of 0 and +/- 5 degree. The test Reynolds number varied from 33.8 to 46.9 millions per meter. Pressure coefficients are compared with those obtained in 1.2M tunnel on a 1/40 scale PSLV configuration at a free stream Mach number of 2.5. The pressure variations are in agreement except in a short region where the pressure coefficients obtained in 0.3M tunnel are lower compared to those of 1.2M tunnel. This discrepancy could be due to the presence of link in the strap-on region and also due to viscous effects arising out of difference in Reynolds number

Investigation of flow field on a hammerhead nose configuration at transonic speeds AIAA-91-1711

Flow field at transonic Mach numbers of a hammerhead13; nose configuration with boat tail an les in13; face flow visualisation and surface pressure measurements.13; Surface steady pressure data shows presence13; of a shock on the cylindrical portion of the body that13; shifts downstream with increase in Mach number and13; boat tail angle. Maximum travel of the shock wave is13; observed at boat tail angle of 90 degrees. Analysis of13; unsteady pressure data shows evidence of shock oscillations13; with multiple frequency content at M=0.9.13; Results of the tests on a scaled model in two different13; tunnels show that the oscillations are due to the13; nature of unsteadiness in the tunnel. Variation of the13; boat tail angle does not appear to alter the unsteadiness13; in the cylindrical region at transonic speeds

Pressure measurements in the heat shield region of 1/80 scale PSLV model In the mach number range of 0.8 to 4.0

Pressure measurements were carried out in the heatshield region of a 1/80 scale PSLV pressure model in the Mach number range of 0.8 to 4.0 and at incidences of 0 and 5 degrees in the 1.2M tunnel to supplement the pressure data obtained in 0.3M tunnel where tests on the model covered transonic Mach numbers only and also to have a comparison of the pressure data with that of 0.3M tunnel for any tunnel/support system interference. Results show a good comparison of the pressures with respect to those of 0.3M tunnel at M=0.78

Horizontal Tail And Flap Effectiveness Studies On Mig-21 M Aircraft Model

Wind Tunnel Tests Were Carried Out On A 1/17 Scale Mig-21 M Aircraft Model (With No Vortex Plate On The Wing) With Trailing Edge Flaps Deflected At 3,6,9,12 And 15 Degrees And Horizontal Tails Set At 5,0,-5,-10 And -13 Degrees. Tests Covered A Mach Number Range Of 0.5 To 1 .1 And Incidence Range Of -5 To 15 Degrees. Reynolds Number Based On Mac Varied From 3 .9 To 6 .3 Millions. Analysis Of The Test Data Indicate The Following: 1. Flap Effectiveness Studies Show A Constant Value Of Cl B F At .011 Per Degree (0 46f 4 15) Up To M = 0 .8 ; Cl 6 F However Decreases Beyond M = 0 .8 And At Higher Incidences. 2. Results With Tail Deflections Show A Constant Value Of Cm 5 ht At .005 Per Degree (-13 < 6 Ht 4 5)Up To Cl Of 0 .5 In The Range Of Mach Numbers Tested. 3. At A Free Stream Mach Number Of 0.7, Results With The Optimum Flap Deflection Schedule Show An Increase In (L/D) Trim Of 15 Percent Compared To That Of Flap Deflection Zero Degree . Also The Increase In (L/D)T Im For Optimum Flap Deflection Schedule Is Higher When Compared To Thai Obtained With Vortex Plate On The Wing

Development of smart actuators for active flow control at low reynolds number

The flight regime of small aerial vehicles fall under low Reynolds number flows where the phenomenon like laminar separation bubble and hysteresis significantly affect the performance of the airfoil. Active flow control is one of the important research areas that have high relevance to small aerial vehicle like UAVs and MAVs [1]in the context of achieving high aerodynamic efficiency. It is a multidisciplinary research area combining sensing, actuation, flow physics and control. Here flow field is manipulated using a time dependent forcing system, typically to leverage a natural instability of the flow. Advantages include the ability to attain a large effect using a small and localized energy input. The rapid progress in the sensor, actuator and embedded hardware technologies has been significantly contributing to the development of different flow control techniques based on Piezo Actuator (PZT), plasma actuator, magnetostrictive, Electro Active Polymer (EAP) etc. Attempts have been made at NAL in the recent times in the development of smart actuators for the flow control using devices using PZTs, EAPs etc. In order to characterize these devices, the flow field and its modifications are being studied using advanced flow diagnostics tools like Laser Doppler Velocimetry (LDV) and Particle image Velocimetry (PIV). This paper presents work on the development smart devices for flow control and their effective deployment

Evaluation of aerodynamic performance of sidewall compression intake relevant to air breathing propulsion vehicles for intake flow entry at M = 3.5

Wind tunnel studies were carried out on a 1:8.5 scale of side wall compression intake configuration model at free stream Mach number of 3.5. Starting characteristics of the intake was ascertained through wall pressure distribution and flow visualization pictures. The model has a convergent section, a nearly constant area section and divergent section, followed by a mass flow meter with a variable rear flap. Sidewalls with different leading edge sweeps of 30 and 45 degrees have been tested. Effect of sweep on mass flow has been studied for all cowl lengths. Pressure recovery is obtained for maximum cowl length. Results show sidewall sweep and the cowl length affect the mass flow entering the intake; and the maximum mass flow is seen with 30F sweep and the maximum range of mass flow is seen with differential sweep for a minimum cowl length

Model support system interfrenceon zero-lift drag at trasonic speeds AIAA - 78-809

In this paper the support system interference on the zero lift drag of an axisymmetric and an aircraft type models is discussed. Two different techniques evaluate the support sting interference. It is found from these tests that the result in a reduction in the zero lift drag of as much as 20 to 50 precent of the true value. This apparent reduction in drag is found to be a strong function of unity Detailed pressure measurements over the aft-body of the axisymmetric model suggests that due to the positive pressure field imposed by the sting over the best tail region of the model the free stream Rach number at which the shocks appear in the boat-tail region will be higher when the sting is present than that without it. This will result in an increased drag divergence Mach number for the model in the presence of the sting. It is argued that because of this reason the sting effect on zero-lift drag strongly depends on the Mach number close to unity