Flow Characteristics of the University of Florida-REEF Supersonic Wind Tunnel

Transcription

1 6th AIAA Aerodynamic Measurement Technology and Ground Testing Conference<BR> - 6 June 008, Seattle, Washington AIAA Flow Characteristics of the University of Florida-REEF Supersonic Wind Tunnel Jonathan Dudley, George Shumway, Charles Tinney and Lawrence Ukeiley The University of Florida Research and Engineering Education Facility (REEF) Unsteady Fluid Dynamics Group recently designed and installed a Supersonic Wind Tunnel. The purpose of the facility is to provide supersonic flow conditions for performing research on active and passive control techniques of cavity flows. Research will be conducted using surface pressure measurements, Schlieren and planar optical flow measurement techniques such as Particle Image Velocimetry. The initial nozzle was constructed for M=1.4 flow providing approximately 00 s of run time at a stagnation pressure of 17.9 kpa. Several static and stagnation pressure measurements have been recorded including streamwise and spanwise pressure gradients along with bottom and side wall boundary layer profiles. Imaging of the compressible flow field was accomplished using a classical Z-Schlieren arrangement providing qualitative information of the reflected Mach waves. I. Introduction Studies of the aeroacoustic environment encountered in weapons bays have been ongoing since the early 1950 s., 5 The flow field resulting from the flow over relatively simple geometries (e.g. rectangular) has proven to be very complicated. The flow involves shear layer instabilities and acoustic resonance which result in significantly high levels of the fluctuating surface pressure inside the cavity. The characteristics of this type of flow field and resultant surface pressures have been shown to be strongly dependent on many parameters such as cavity length-to-depth ratio (L/D), approaching boundary layer and freestream conditions. The necessity to reduce the fluctuating surface pressures is predicated by the delicate electronics that might exist in the environment of real aircraft and requires a clearer understanding of the flow dynamics. To this end a supersonic wind tunnel (SWT) was developed by The Unsteady Fluid Dynamics Group at the University of Florida Research and Engineering Education Facility (REEF) in Shalimar, FL so experiments involving supersonic flow over open cavities can be conducted. The initial design of the wind tunnel has a fixed geometry nozzle designed for Mach 1.4 which can be replaced with different nozzles for desired Mach numbers. The test section has optical access on all four flow-surfaces to facilitate advanced optical measurement techniques such as Particle Image Velocimetry (PIV) and Schlieren techniques. In addition to velocity and density measurements, the surface pressure at various locations can be measured via dynamic pressure transducers. With the completion of this facility detailed studies of sources in the flow field will be conducted using synchronous qualitative schlieren and surface pressure measurements and synchronous PIV and surface pressure measurements with passive or active control. II. The Wind Tunnel Design The initial major design criteria for the SWT included a modular design to optimize flexibility and a test section cross-section of 76. mm x mm with optical access on all four sides to mount a L/D=6 cavity. The selection of the test section cross-sectional area and the operating Mach number were the key fixed parameters that drove the design of the SWT. The selection of the 44 m storage tank and consequently the compressor, dryer and control valve were dictated by this decision. The high pressure air for the tunnel Graduate Student, University of Florida MAE, Shalimar, Fl 579, Member Visiting Assistant Professor, University of Florida MAE, Shalimar, Fl 579, Senior Member. Assistant Professor, University of Florida MAE, Shalimar, Fl 579, Senior Member. 1 of 1 Copyright 008 by the American Institute of Aeronautics and American Astronautics, Institute Inc. All of rights Aeronautics reserved. and Astronautics

2 is supplied from a rotary screw Quincy Compressor (model QSI 1000 A/C) capable of producing 6.4 m /min at kpa through 50.8 mm piping. From the exit of the compressor the air flows through a desiccant dryer to lower the dew point of the air to. K before entering the storage tank. (a) Compressor (b) Dryer (c) Tank Figure 1. SWT Compressor, Dryer and Storage Tank In order to minimize the generation of excess flow-noise inside the room a mm diameter pipe was run from the tank into the building which maintained airspeeds under 0.5 m /s. After a mm isolation valve the pipe diameter expands to 15.4 mm and enters the 15.4 mm Fisher control valve. The control valve was selected with an appropriate flow coefficient (CV) to maintain the desired flow conditions based on typical upstream and downstream conditions. In order to minimize turbulence levels entering and exiting the control valve at least 4 valve diameters of straight pipe was placed before and after the valve. Some basic preliminary calculations using 1-D gas dynamics were performed to determine key design parameters such as maximum run times, maximum allowable L/D ensuring shock reflections clear the cavity and maximum cavity length which keep frequencies which one desires to control below 4 khz. The maximum frequency constraint is dictated by the active control hardware. The Mach angles were determined from similar Mach number experiments and simulations of cavity flows. From these computations the wind tunnel was anticipated to run at a stagnation pressure of 17.4 kpa, a test section static pressure of 48.6 kpa and a mass flow rate of 0.45 kg /s with a run-time of approximately 00 s. Figure and illustrate the layout of the blow down SWT. (a) Facility Layout (b) Supersonic Wind Tunnel Figure. Facility Layout Diagram and the Supersonic Wind Tunnel of 1

3 Figure. Supersonic Wind Tunnel A. The Components of the SWT In the rest of this section we will discuss the individual components of the tunnel in the order the air flows through them. To maintain simplicity and cost effectiveness the plenum and flow-conditioner sections were designed using 04.8 mm internal diameter schedule 40 and 60 stock carbon steel piping as seen in Figure 4 and were treated with an epoxy based paint. The plenum is a 04.8 mm x 04.8 mm x 15.4 mm reducing T-section that is mm long. Subsonic air from the control valve enters the plenum through the 15.4 mm leg and settles in the chamber reducing the velocity and turbulence acquired through the delivery system. One 04.8 mm run is sealed with a blind flange which allows for the mounting of a Resistance Temperature Detector (RTD) to measure the stagnation temperature. The second 04.8 mm run of the plenum joins the honeycomb-seeder section which is a 81 mm long, 04.8 mm diameter pipe. The first 17 mm of the inlet was bored out to produce a positive stop for a Corning 400 (or 600) cell per square inch Celcor ceramic honeycomb substrate. The substrate is used to condition the airflow to reduce the scale of any remaining coherent turbulence in the flow. The remaining 54 mm of the pipe allow the flow to settle after the honeycomb. Swagelok fittings with. mm inner diameter were welded 17 mm upstream of the exit of the honeycomb-seeder in -alternate azimuthal patterns to allow for flexibility when introducing seed into the flow for PIV experiments. Three ports were mounted on the top at 0 spacing while four additional ports at 45 were mounted on the bottom of the chamber. A pressure port was added 5.4 mm downstream of the seeding location to measure the stagnation pressure for the control system using a 44.7 kpa Druck pressure transducer. (a) Stagnation Chamber (b) Seeder (c) Substrate Figure 4. Stagnation Chamber After the honeycomb-seeder section there is a round-to-rectangular transition piece (04.8 mm diameter 76. mm x 54 mm) which was designed using a Matlab algorithm detailed in Matthew et al. 4 Schematics and photographs of this piece can be seen in Figure 5. The code supplied cross-sections at user defined constant intervals. The xy cross-section point data was then imported into Pro-Engineer to generate a solid -D geometry using variable sweeps. Computational fluid dynamics was used in conjunction with the Matlab code to verify the effectiveness of the code by examining boundary-layer build-up and the resulting pressure of 1

4 loss. The transition piece was fabricated from 6061-Al using a combination of Wire-EDM, milling and hand finishing to produce a smooth sealed one piece design. (a) TP End View (b) TP Internal Volume (c) TP Centerline Cross Section Figure 5. Transition Piece The rectangular exit of the transition piece mates up with the Mach 1.4 converging-diverging nozzle which was designed using method of characteristics and validated with CFD. The nozzle was constructed from two machined pieces of 7075-Al, a nozzle block and a covering plate. The converging-diverging nozzle block was end-milled and finished with a 16 surface finish to minimize disturbances and viscous effects caused by surface roughness in the flow. The inlet of the nozzle was designed with a 76. mm x 54 mm cross section to match the transition piece. The throat area is contracted vertically to mm and expands to match the test-section of 76. mm x mm by the exit. A cover plate was used to bolt over the nozzle block leaving the nozzle assembly with only one sealing surface. (a) Nozzle (b) Nozzle Wire (c) RANS CFD Figure 6. Nozzle and RANS CFD of Nozzle The nozzle exit is joined directly to the test section, a rectangular frame with a flow volume of 76. mm x mm x 04.8 mm (H x W x L) and is illustrated in Figure 8. The test section assembly was designed with a single piece frame which allows inserts to be mounted on the 4 exposed surfaces. These inserts will typically be fabricated with PlexiglasR G cell-cast acrylic for optical purposes. The test section cross-sectional area is unconventional for supersonic wind tunnels of this scale with the major dimension in the vertical direction of mm and the minor dimension in the horizontal direction of 76. mm. Previous experimental and CFD simulations provided the angle of waves from the leading edge of a cavity in a similar flowfield as shown in Figure 7. Using trigonometry and the Mach angle while assuming a perfect reflection off the top wall, the position where the reflected wave strikes the floor was calculated as a function of the test section height. The test section height was chosen so that the distance between the leading edge 4 of 1

5 of the cavity and the reflected wave incidence on the floor was two cavity depths greater than the maximum desired cavity length. Figure 7. Cavity Shock Reflection Diagram The floor (area where the cavity is) can accept various inserts with various cavity configurations as well as an insert with no cavity to make a straight duct for calibration purposes. The symmetry of the test section allows for the cavity to be top or bottom mounted. The cavity will be geometrically limited to maximum dimensions of 8.6 mm x 76. mm x 5.4 mm (L x W x D). A 0.4 mm diameter pressure tap was used to minimize boundary layer disturbance and acoustical resonance and is located upstream of the cavity insert to measure the static pressure for the control system. (a) Test-Section with Flat Wall (b) Test-Section with Cavity Mount (c) Cavity Figure 8. Test-Section The first designed -D cavity is 76. mm long and 1.7 mm deep providing an L/D=6 and is a four piece construction as seen in Figure 8. Two plexiglas blocks are held in place by an aluminum frame containing an O-ring cutout. The assembly is either top or bottom mounted in the test section and is lastly covered with a plexiglass plate to complete the seal. After leaving the test section the air enters the diffuser which incorporates.75 diverging half angles and was manufactured with 1.7 mm thick 6061-aluminum plates welded along the seam lines. The diffuser is capable of sliding into the exhaust-pipe to allow access into the end of the test section. A picture of the diffuser is located in Figure 9. The exhaust-pipe accepts the diffuser and passes through the building wall where it then turns 90 upward and exhausts safely to the environment. The end is chamfered at 45 and screened to protect the tunnel from environmental hazards. 5 of 1

6 (a) Diffuser (b) Exhaust Figure 9. Diffuser B. Instrumentation The wind tunnel is instrumented to monitor static pressure, stagnation pressure and stagnation temperature. A Druck PMP4015 pressure transducer is used to monitor the absolute stagnation pressure upstream of the nozzle at the surface of the honeycomb-seeder where the flow is assumed to be slow enough that static and stagnation pressures are nearly equivalent. The transducer has a pressure range of kpa with an accuracy of 0.08% full scale. The over pressure rating of the transducer is four times the full scale and it outputs 0 to 10 VDC. A second pressure transducer measures the absolute static pressure at the entrance of the test section and its specifications are identical except the pressure range is from kpa. A four lead wire 100 Ω RTD supplied from R.T.D. Company is mounted in the plenum end flange to monitor the stagnation temperature. The RTD sensor is 50.8 mm long with an exposed platinum element to increase the response time and supplies measurements with +/- 0.06% accuracy of the measured temperature. C. Control The tunnel is controlled by a 15.4 mm Fisher Control valve operated by an PC-based LabVIEW algorithm implemented on a FPGA card. A National Instruments PXI-10 chassis and PXI-781R card uses measurements from Druck 10.4 kpa and 44.7 kpa pressure transducers to measure the nozzle exit static pressure and the tunnel stagnation pressure in the seeder assembly. The 781R has 8 analog inputs and 8 analog outputs, each with a 16 Bit resolution and a Voltage range -10 to 10 VDC. It also has a digital input/output interface with 96 channels with a 0 to 5.5 Volt range. The SWT is controlled using a LabVIEW virtual instrument (vi). The vi runs in a loop that is 0.5 seconds long, during which it averages 50 measurements from two pressure transducers and one RTD. It is capable of manual control, where a user controls the tunnel by opening and closing the control valve directly, and automatic control, where the vi compares the measured stagnation pressure to a target value and operates the control valve accordingly. The auto control is a basic control scheme which opens if the measured pressure is below the target and waits if the measured pressure is above the target. It has built in delays and offsets to account for the response time of the control valve, and the response time of the air flow itself. These delays and offsets were determined empirically and optimized for a stagnation pressure range from 14.1 kpa to 17.4 kpa. However they are not perfect, and in the future a more rigorous and systematic control scheme will be implemented, most likely a PID controller. It should be noted that we are not measuring the stagnation pressure in the flow conditioning section as this measurement is acquired from a static port to eliminate an intrusive probe that would create flow disturbances. The differences between the static and stagnation pressures with slow moving air are expected to be small and were checked empirically and and analytically presented in the results section. 6 of 1

7 D. Characterization The wind tunnel flow was characterized using qualitative Schlieren imaging, surface pressure gradients, and boundary layer measurements. Particle image velocimetry will be utilized in the future to measure the velocity distributions in the flow field. The Z-type Schlieren setup is shown in Figure 10 which consists of two 15.4 mm inch parabolic mirrors with 11.9 mm, a mm arc length Xenon flashlamp with a µs flash duration, a Sony XCD-V60 digital video camera, a simple focusing lens and a trigger synchronizing the light source. Figure 10. Qualitative Schlieren Measurements Static pressure gradients on the bottom and side wall of the test section were measured in three directions: streamwise (dpdx), spanwise (dpdy) and wall-normal (dpdz) corresponding to 11. The locations for static and stagnation pressure measurements are given in Figure 1 along with Table 1. The X and Y coordinates are non-dimensionalized following equation 1 given below where X ts,l and Y ts,w are the internal dimensions of the test section: length, width and height respectively. Pressure taps with 0.4 mm diameter were drilled for static pressure measurements in the test section walls while the stagnation pressure probe ports were. mm in diameter. A United Sensor stainless steel boundary layer probe was used with a sensing head diameter of 0.64 mm flattened down to 0. mm. Boundary layer measurements were taken from the leading edge of each plate at the P0 and P16 locations of the top and side wall respectively as shown in Figure 1. The test section symmetry allowed the bottom and side plates to be reversed for measuring the static and stagnation pressures at the test section entrance. A kpa Druck pressure transducer was used to measure the directional pressure gradients. A kpa Druck sensor was used for boundary layer stagnation pressure readings on the bottom and side walls of the test section. To evaluate the pressure gradients and boundary layer measurements the above mentioned transducers were connected to a PXI-447 card which has 8 simultaneously sampled analog inputs with 4 bit A/D resolution. The pressure transducers were sampled at 1 khz and data was averaged over a 0 s window. (a) Coordinate System (b) Boundary Layer Measurement Figure 11. Coordinate System and Boundary Layer Measurement 7 of 1

8 (a) Bottom Wall Pressure Taps (b) Side Wall Pressure Taps Figure 1. Side Wall and Bottom Wall Pressure Tap Locations X = X X ts,l, Y = Y ( Yts,w ) (1) Bottom Wall Side Wall P X Y P X Y P X Z P X Z Table 1. SWT Pressure Tap Locations III. Results The Mach number in the facility was obtained by measuring the static pressure in the test section and in the seeder assembly in conjunction with Equation. ( p t p = 1 + γ 1 ) γ M The nozzle exit stagnation pressure was assumed to be equal to the static pressure measured in the honeycomb/seeder assembly. Provided the settling chamber is large enough to adequately slow the flow resulting in a sufficient contraction ratio for the nozzle chamber, this estimation is typically sufficient for flow where M The calibration of the nozzle, due to the above mentioned assumption, is presented in Figure 1. The corrected Mach number was found to be ± 0.5% of the recorded value. The difference between the static and stagnation pressures in the the seeder assembly was also verified analytically using an iterative method assuming isentropic flow and knowing the flow has to be choked at the nozzle throat. γ 1 () 8 of 1

9 Figure 1. Correction for Stagnation Pressure The transition of the start up shock through the test section for boundary layer measurements is illustrated in Schlieren images in Figure 14. The shock is completely outside of the viewing area by a stagnation pressure of 14.1 kpa thus representing the minimum stagnation pressure the tunnel can be operated at. For a cavity mounted in the test section the tunnel starts and can be run at a stagnation pressure of 17.9 kpa. (a) Start Up Shock Position1 (b) Start Up Shock Position (c) Start Up Shock Position Figure 14. Test Section Start Up Shock Figure 15 illustrates a controlled run for the facility when measuring a pressure gradient point in the test section at a set-point of 17.9 kpa. The assumed stagnation pressure is represented by P 0 and the measured static pressure at the exit of the nozzle is illustrated by P ref. The resulting Mach number and stagnation temperature are also shown in the plot. The error in determining the Mach number variations due to calibration accuracy using static and stagnation pressures given by Pope 1 was found to be % from Equation. dm M = 1 + γ 1 M ( dpt γm dp ) () p t p 9 of 1

10 Figure 15. Typical Controlled Experiment The drop in stagnation temperature due to the pressure decrease in the storage tank is known to effect the test section Reynolds number. The Re x (per mm) was calculated following Equation 5. The absolute viscosity was corrected for temperature changes using Sutherland s law and is shown in Equation 4. Density was corrected for both pressure and temperature assuming an ideal gas. ( ) T T ref + S µ = µ ref T ref T + S (4) Re x = ρu µ = ρm γrt µ (5) ρ = P RT (6) Figure 16 illustrates a typical change in Reynolds number for running pressures of 17.9 kpa and 06.8 kpa over a 0 s run-time. The rise in Re x due to dropping stagnation temperature in a typical run are on the order of 5%. At the current time, the facility offers no way of controlling Reynolds number or Reynolds variations due to temperature change in the system. This could be overcome by controlling the stagnation temperature with the use of a inline heater or thermal mass inside the tank. Figure 16. Re x Variation with Stagnation Pressure Figure 17 shows the bottom and side wall pressure variations and repeatability when altering the stagnation pressure. The pressures are non-dimensionalized per run according to Equation 7 while the X and Y 10 of 1

11 coordinates follow Equation 1 given above. The runs were consistent to within ± 4% on repeated runs and and higher dynamic pressure runs. The stagnation pressure was raised to demonstrate the tunnels capability at higher dynamic pressures. p = p measured P ref (7) (a) Bottom Wall dpdx (b) Side Wall dpdx (c) Bottom Wall dpdy Figure 17. Bottom Wall Static Pressure Gradients The presence of waves are apparent in the test section as shown in Figures 17 and 18. The waves originate from the joint of the test section to nozzle, the joint between the optical plexiglass and test section frame and there are some Mach waves from inside the nozzle. The measured static pressures on the bottom wall rise 14% over the entire length of the wall and 4% over the length of the -D cavity. The side wall pressure rise in the streamwise direction is approximately 9% while the pressure variation in the lateral direction is to %. The pressure rise in the streamwise direction is consistent with a thickening boundary layer leading to a reduced Mach number. 11 of 1

12 (a) Test Section Schlieren (b) Pitot Tube Schlieren Figure 18. Test Section Schlieren Images The boundary layer thickness and profile was estimated by measuring the stagnation pressure at a given X (bottom wall) or Y plane (side wall) and comparing to the average static pressures obtained at x = If the pressure ratio indicated subsonic flow Equation was used to compute the subsonic Mach number. If a supersonic ratio was found the Rayleigh Pitot Equation given by James in Equation 8 was used to correct for the bow shock in front of the tube as seen in Figure 18(b). An iterative Matlab script was written to solve for the supersonic Mach number. The Mach number is normalized by the freestream Mach number and y is non-dimensionalized by the 1 the vertical height of the test section. P t P s = ( γ+1 M ) γ γ 1 ( γ γ+1 M γ 1 γ+1 ) 1 γ 1 (8) Figure 19. Boundary Layer Mach Profiles for Bottom and Side Plates A line was plotted for the 99% Mach number to determine the cutoff for measuring the boundary layer height. From Figure 19 the boundary layer height was measured as.5 mm for the sidewall and.8 mm on the top wall. Included in Figure 0 (with flow left to right) are instantaneous snapshots demonstrating Schlieren capability for a supersonic L/D=6 -D cavity. The images indicate little dependence of the shear layer formation due to the interaction of waves generated by the nozzle or test section mounts penetrating the 1 of 1

13 shear layer. The shear layer and the strong shock formed by the leading edge of the cavity are easily noted. The shock reflects off the upper wall of the test section and hits the floor clearly downstream of the cavity as expected and seen in the upper right corner of Figure 0. Figure 0. Random Schlieren Images of the Supersonic Flow Over the L/D=6 Cavity IV. Conclusion The supersonic wind tunnel at The University of Florida Research and Engineering Education Facility has been installed and has demonstrated acceptable flow characteristics for the future study of supersonic cavity flows. The tunnel s boundary layer profile and pressure gradients detailing the flow quality in the test section have been presented. The static pressure variations in the flow, lateral and vertical directions were found to be under 4% over the length of the installed -D cavity. Qualitative Schlieren capabilities have been demonstrated through imaging of the test section at supersonic speeds and for an L/D=6 -D supersonic cavity. The imaging has demonstrated that there is little interference of reflected waves with the installed cavity. A more rigorous and systematic PID control scheme is in the design process and will be measured against the existing algorithm for effectiveness. Additional instrumentation on the storage tank including an RTD and pressure transducer will likely be installed. It may be necessary in the future to implement a filtration system for rust particulate due to the carbon steel piping. Acknowledgments Support from the authors has been provided for by AFOSR and USAFSEO through an IHAAA grant while support for the tunnel construction has been provided for by the University of Florida. These sources of funding are greatly appreciated. The authors also acknowledge the support of our colleagues at Florida State University for the time touring and monitoring their facilities. References 1 Pope A. and Kenneth G. L. High Speed Wind Tunnel Testing. John Wiley and Sons, Inc, New York, NY, first edition, Roshko A. Some measurements of flow in a rectangular cutout. Technical Report Technical Note 488, James J. and Keith T. Gas Dynamics. Prentice Hall, New York, NY, rd edition, Mathew J., Bahr C., Carrol B., Sheplak M., and Cattafesta L. Design fabrication and characterization of an anechoic wind tunnel facility. In 11th AIAA/CEAS Aeroacoustics Conference (6th AIAA Aeroacoustics Conference), Monterey, Ca, -5 May 005, Ca, USA, 005. AIAA. 5 Krishnamurty K. Acoustic radiation from two-dimensional rectangular cutouts in aerodynamic surfaces. Technical Report Technical Note 487, NACA, of 1