Experimental Investigation into the Aerodynamic Stability of a Suborbital Reusable Booster Concept

Transcription

1 50th AIAA Aerospace Sciences Meeting including the New Horizons Forum and Aerospace Exposition January 2012, Nashville, Tennessee AIAA Experimental Investigation into the Aerodynamic Stability of a Suborbital Reusable Booster Concept Andrew R. Miller 1 and Martiqua L. Post 2 U.S. Air Force Academy, Colorado, A responsive, next-generation reusable booster system is a vital part of the United States future in a world of rapidly advancing technology. As part of an Air Force Research Laboratory s (AFRL) program, the aerodynamic stability of a scale model of a suborbital reusable booster concept is evaluated. Three different methods are used to quantify and evaluate the aerodynamic stability of the model with different control surface configurations. A wind tunnel experiment is used to quantify the forces and moments acting on the model; pressure sensitive paint and water tunnel visualizations are used to obtain a more complete understanding of the airflow around the model and its effects on aerodynamic stability. The two vortices, which form on the booster concept s double-delta wing design as angle of attack increases, affect the model s stability. Wind tunnel results show unstable trends in pitching stability beginning between 18 and 22 angle of attack. Pressure Sensitive Paint (PSP) analysis and water tunnel visualizations show the mechanism behind the instabilities is a vortex breakdown off the outboard section of the wing. PSP visualizations indicate the outboard vortex breakdown occurs between 15 and 25 angle of attack. Water tunnel flow visualizations confirm the vortex breakdown and indicate spanwise flow induced by the vortices. C P C L C M CFD CCD GUI I 1 I 2 K M m P 1 P 2 psia PSP UV α Nomenclature = pressure coefficient = lift coefficient = pitching moment coefficient = computational fluid dynamics = charge-coupled device = graphical user interface = emission intensity at reference pressure (wind off) = emission intensity during tunnel run (wind on) = Stern-Volmer constant = Mach number = meter = surface static pressure at reference pressure (wind off) = surface static pressure during tunnel run (wind on) = pounds per square inch absolute = pressure sensitive paint = ultra-violet = angle of attack I. Introduction The USAF is in need of a, routine, reliable, and affordable flight of diverse fully reusable space access vehicles, to, enable aircraft-like space access capabilities. 1 In response to this need AFRL contracted a series of ground experiments to test new technologies designed to, enable affordable and responsive space access. 2 An experimental investigation into the aerodynamic forces and stability on a suborbital reusable booster concept 1 2 nd Lt, USAF, Columbus AFB. 2 Associate Professor, Department of Aeronautics, Senior AIAA Member. 1 This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States.

2 concluded that the concept and has stability issues at angles of attack greater than 20 degrees. 3 The report recommended that, Additional research should be conducted to investigate the longitudinal instabilities at higher angles of attack to better understand what aspects of the flow could cause the instability. 3 In order to investigate and understand the causes of the instabilities, this research paper presents and experimental investigation into the forces and moments acting on a suborbital booster in the approach landing phase of flight as well as a Pressure Sensitive Paint (PSP) investigation and water tunnel visualization of the flow field around the model during the same phases of flight at Mach 0.3. A B Figure 1. (A) Control Surface Locations on Test Model and (B) Painted with PSP and Illuminated with UV Light. Model fabricated via sterolithography and the wing surfaces were coated with PSP after it was used to obtain force and moment data. II. Objectives The objectives of the initial research into the forces and moments acting on the model were to validate CFD data provided by AFRL. This research expands on the initial research and uses two flow visualization techniques to explain why the forces and moments on the model act and change the way they do. Additionally, the research also sought to develop reliable software for future PSP research at the United States Air Force Academy (USAFA). A. Experimental Approach III. Methods Test Model The model of the reusable booster concept model is shown in Figure 1. The model is 0.37m in length excluding the body flap and 0.185m in span. As illustrated in Figure 1, there are four control surfaces: the body flap, inboard elevons, outboard elevons, and rudders. Positive control deflections for the body flap and both sets of elevons is trailing edge down and positive rudder deflection is trailing edge left, causing a positive sideslip angle with wind from the right. The model was fabricated using stereo-lithography. The various control surface configurations were fabricated on separate wings which can be replaced on the model. Wind Tunnel Testing of the reusable launch vehicle was performed in the Subsonic Wind Tunnel Test Facility in the Aeronautics Laboratory at the USAFA. The Subsonic wind tunnel is a closed circuit wind tunnel with a speed range of feet per second. The free-stream Reynolds number was approximately 750,000 for tests at M = 0.3. Data Acquisition The test model for the reusable launch vehicle is placed on a force balance mounted on a sting in the wind tunnel test section. A TASK lb force balance was used to measure five forces (two normal forces, two side forces, and one axial force) and one rolling moment acting on the model. A total pressure probe located upstream of the test section and four static ports located at the entrance of the test section were used to monitor tunnel velocity. The total pressure probe was connected to an absolute pressure transducer. The static pressure ports and the total pressure probe were connected to a differential pressure transducer. Both pressure transducers, as well as the force balance, were connected to a 16-bit analog-to-digital converter. Data was collected at 100 Hertz and saved to a computer for analysis. 2

3 In-house software is used for data processing. The acquired voltages from the force balance and pressure transducers are converted to aerodynamic forces, moments, and velocity. In-house software also uses the moment offset and characteristic chord to calculate force and moment coefficients. A model with known aerodynamics characteristics verified the force balance calibration. Once verified, the reusable booster model was secured to the force balance and sting. The tunnel door was secured, and data was collected in an air off test. This air off data was used as the zero reading for all model configurations. Test Matrix Data was collected with a clean configuration, all control surfaces at zero degree deflections, and then with each surface deflected in a systematic manner. Table 1 summarizes the test matrix. Once each data run was complete, the in-house software reduced the voltage output file into forces and coefficients. Data was acquired at low angles of attack using a straight sting and at angles of attack higher than 28 degrees with a bent sting. Table 1. Test Matrix. Mach # Control Surface Deflection 0.3 Clean configuration N/A 0.5 Clean configuration N/A 0.3 Body Flap -10,15, Rudder -30, Outboard Elevon -30, Inboard Elevon -30, Speed Brakes 15, 30 B. PSP Visualization Pressure Sensitive Paint (PSP) allows visualization and non-intrusive measurement of static surface pressures on a model coated with the paint. PSP consists of a luminescent molecule suspended in an oxygen permeable binder. 4 The following process allows for pressure data collection; a PSP coated surface is illuminated by an ultraviolet light which excites the luminescent molecules. The excited molecules re-emit absorbed energy which is captured by a camera and recorded on a computer. The luminescence of the molecules is reduced when there is a higher concentration of oxygen molecules near the surface of the model. The oxygen interacts with the luminescent molecules in a quenching, 4 effect on the light emitted. The amount of oxygen at the surface of the model is directly proportional to the static surface pressure, allowing a relationship between luminescence and surface pressure to be derived. This relationship is derived from the Stern-Volmer relation between luminescence and pressure seen in Eq. (1). 5 = + P (1) Since P 1 is a reference pressure it can be considered a constant which results in a linear relationship between P 2 and the luminescence ratio on the left-hand side of Eq. (1). The relationship can be determined by through calibration relating known pressure to a given luminescence ratio. Test Procedure The wings used on the model for wind tunnel testing are coated with PSP. The model is then configured with the desired control surface deflections. Once configured, two ultraviolet lights are positioned to provide good illumination of the surface of interest. A CCD Camera with a variable filter is positioned and focused on the surface of interest. The camera and filter switch are attached to a computer with OMS Acquire software, which is used to capture images at two filter settings. The camera s exposure time is adjusted to ensure the luminescence captured falls within the intensity range the camera s sensors can record. With the wind tunnel off and room lights out, two sets of images are captured: a background image with the ultraviolet lights off and another with the ultra violet lights on. These two images will be combined later to form the wind-off image. The wind tunnel is then run at M = 0.3 and once stabilized, the same two sets of images are recorded with the ultraviolet lights on and off. The tunnel is deactivated and the model is reconfigured with the next test point. The control configurations investigated are highlighted in blue in Table 1. This investigation involved 27 wind tunnel tests of seven control surface configurations at three angles of attack: 0, 15, 25. 3

4 B C Figure 2. In-Situ PSP Calibration. Pressure ports on images in upper right capture surface static pressures which are correlated with the intensity ratios around the ports in the lower right. Calibration equation used in MATLAB for image processing. Image Processing Once the images are recorded and saved, the images are processed using a GUI MATLAB program developed by the researcher. Screenshots and brief explanation of the program can be seen in the appendix. The image processing begins with four images: wind-off, wind-off background, wind-on and wind-on background. All images are loaded into the GUI and the background images are subtracted, pixel by pixel, from their respective illuminated images to eliminate the interference of any ambient light. Due to a slight shift in the model s position between windon and wind-off conditions, the user is then prompted to select control points on each image which will be used to align the images. These control points allow the user to select physical features on the model which are used to align the images with a built-in MATLAB control point command through a local weighted mean algorithm. Once aligned, the luminosity intensity ratio on the left hand side of Eq. (1) can be found by dividing the intensity of each pixel in the wind-off image by its corresponding value in the wind on image. This results in a single image for each model configuration. Once the luminosity image is defined, the intensities are converted to pressures through the linear relationship in Eq. (1). The numeric values associated with the equations are determined using and in-situ method of calibration shown in Figure 2A. Direct measurement of surface static pressures through the pressure ports on top of the wing, shown in Figure 2B, allowed the relationship between luminosity intensity ratio (Figure 2C) and pressure to be plotted (Figure 2A) and fit with the linear relationship in Eq. (1). The calibration used five separate tests with four separate static pressures recorded on the surface of a wing during every test for a total of twenty calibration points. The luminosity intensity ratios are found by averaging the values of the intensity ratios around a pressure port. Pressures recorded from the pressure ports are then plotted with respect to the corresponding intensity ratio. A linear fit is also plotted which becomes the calibration equation, described by the Stern-Volmer relationship, Eq. (1), and is used to convert all intensity ratios to pressures. Images of the pressure distribution over the area of interest are then processed and saved. C. Water Tunnel Visualization In addition to surface static pressure visualization of the flow, a water tunnel visualization of the flow around the model was performed in the USAFA Water Tunnel. Manufactured by Eidetics, the water tunnel in the USAFA Aeronautics Laboratory is a closed circuit, free surface tunnel with a speed range of zero to one feet per second in a 15 wide, 20 high and 60 long test section. 6 The same model used for the wind tunnel and pressure sensitive paint investigations was used in the water tunnel visualization. While the tunnel was running at a velocity of feet per second, colored dye was injected into the flow to allow flow visualization. Images were recorded at 0, 15 and 22 angles of attack. Due to tunnel size constraints, the maximum angle of attack achievable by the model was 22 with a Reynolds number of approximately 27,000. 4

5 IV. Results A. Experimental Aerodynamics Lift Coefficient As with any airfoil, as angle of attack increases the lift coefficient increases. Above 22 angle of attack, Figure 3 shows the lift curve slope becomes negative. Normally the slope continues to drop off and the wing stalls past the local maximum critical angle of attack; however, the lift coefficient of this double-delta wing design drops but continues to increase at a shallower. When the slope of the lift curve shallows and becomes negative at some points, it is difficult to determine a clearly defined stall point as increasing the angle of attack both increases and decreases the lift coefficient. Figure 3. M=0.3 Lift Curve and Pitching Moment Coefficient. Lift Coefficient as a function of angle of attack and pitching moment coefficient as a function of angle of attack with different symmetric inboard elevon deflections. The stall characteristics of the model are similar to all double-delta wing aircraft in that the exact angle of attack at which stall occurs is undefined. 1 Pitching moment When an aircraft is longitudinally trimmed, pitching moment coefficient will be zero. A stable aircraft will have a negative derivative, where perturbation in angle of attack from the trimmed condition will produce a stronger restoring moment opposite the rotation of the aircraft as a result of the perturbation. It is also interesting to note that for all control surface the deflections, the slope of the moment coefficient curve becomes longitudinally unstable at approximately α=20. This instability will need to be managed through feedback flight control systems if the aircraft operates at higher angles of attack. Similar stability trends can be observed as there is a slight upward trend in the slope at all deflections at approximately 20 AoA; however, this unstable trend increases in magnitude as the control surfaces transition from negative to positive deflections. Using only inboard elevon deflections between 0 and -15, it is possible to trim the vehicle anywhere between α=4 and α=26. 5

6 Figure 4. M=0.3 Surface Pressure Distributions of Outboard Elevon 30 Deflection. The left wing s outboard elevon is deflected while the right wing has no control surface deflections. Two vortices develop on both wings as angle of attack increases. The outboard vortex on the left wing is stronger at α=15 due to the control surface deflection. Both outboard vortices disappear at α=25. PSP Analysis Pressure distributions are the driving force behind aerodynamic forces and moments. Visualization and analysis of the distributions can help explain why aerodynamic forces and moments produce both expected and unexpected results. Using pressure distributions generated through PSP analysis in Figure 4, it is possible to generate an explanation for this unexpected behavior discussed in the previous section. At α = 15 the right wing in a clean α=25 Outboard Vortex Inboard Vortex Figure 5. Spanwise Distribution of Pressure Coefficients Spikes in the C p curves indicate the presence of vortices. The outboard vortex breaks down between 15 and 25 degrees AoA. 6

7 configuration appears to have vortices developed off the inboard leading edge and outboard leading edge which are of equal strength. At α = 25 the vortex on the outboard leading edge of the right wing is no longer apparent from the static pressures on the surface of the wing. The vortex on the inboard leading edge has become stronger by approximately 0.5 psia at the strongest point in the vortex. This same trend is apparent at α = 15 on the left wing where the outboard elevon is deflected down 30 creating a nose down pitching moment. Unlike the clean configuration, the vortex on the outboard leading edge is stronger than the vortex on the inboard leading edge due to the elevon deflection. When the angle of attack increases to 25 the effects of the outboard vortex are no longer visible and the inboard vortex strengthens. Figure 5 provides a graphical portrayal of this change. Increasing the angle of attack decreases the pressure coefficient and the two vortices can be seen in the green, α = 15, curve. At α = 25, the outboard vortex is no longer creating a low pressure region on the outboard section of the wing. This vortex has most likely broken down as the adverse pressure gradient has overtaken the kinetic energy of the flow over the wing, causing the vortex to break down. The lift curves of the model begin to indicate a decrease in lift coefficient as the outboard vortex breaks down between 15 and 25 angle of attack. The model with the outboard elevon deflection exhibits a larger loss in lift at α = 22, in Figure 3, as the outboard vortex is stronger than in a clean configuration. When the outboard vortex breaks down, the outboard section of the wing stalls. The stronger the vortex when it breaks down, the more pronounced effect it will have on the lift coefficient as evidenced in the lift curves in Figure 3. This outboard vortex breakdown also influences pitching moment. The pitching moment coefficient of a stable aircraft should decrease as angle of attack increases; however, Figure 3 indicates the longitudinally stable trend of a negative derivative in this model becomes unstable somewhere between 15 and 25 angle of attack, precisely where the outboard vortex breaks down. The outboard vortex is further aft of the model s center of gravity than the inboard vortex. When the outboard vortex breaks down, the lift on this portion of the wing is also reduced. The moment arm on this portion of the wing is significantly larger than that at the location of the inboard vortex leading to a reduction in pitching moment coefficient. B. Water Tunnel Images from the water tunnel runs validate the presence of vortical flow indicated by the PSP analysis. The images also show that as the angle of attack increases, the vortices become stronger and induce a span-wise flow over the surface of the wing. Notice the similarities between the surface static pressure measurements using PSP (Fig. 4) and the location and strength of the vortices captured in the water tunnel (Fig. 6). At α = 22, the outboard vortex appears to be breaking down and less clearly defined than at α = 15. Though the outboard vortex is still visible at α = 22, the rear of the vortices appear to be breaking down which suggest that complete outboard vortex breakdown will occur at a slightly higher angle of attack as the adverse pressure gradient advances further up the vortex. Figure 6. M=0.3 Water Tunnel Visualization with 30 Outboard Elevon Deflection on Left Wing. The left wing s outboard elevon is deflected while the right wing has no control surface deflections. Notice the similarities between the vortex strength and location with the PSP surface pressures in Figure 4. Each image for a single α is a combination of six separate water tunnel images. 7

8 V. Conclusions Vortex breakdown is the critical mechanism behind the undefined stall and longitudinal instabilities of this model noted during wind tunnel tests. PSP analysis of the surface static pressure distribution allowed researchers to visualize the pressure distribution created by airflow over the vehicle. This visualization showed how the pressure distribution changed as the aircraft s angle of attack increased. Water tunnel runs verified the cause of the change in low pressure regions on the top of the wing to be primarily influenced by vortical flow. The current research can only see visual evidence that the outboard vortex breaks down somewhere between α = 15 and α = 25. Water tunnel visualizations indicate the outboard vortex is beginning to lose its definition, and is in the process of bursting at α = 22. Results from this study can only conclude that the vortex breakdown occurs somewhere between 15 and 25 angle of attack. Analysis of lift and moment data from earlier studies would suggest the outboard vortex breakdown occurs at approximately α = 23 ±1. The vortex breakdown will only present a potential problem to safe flight if at angles of attack exceeding 15. 8

9 Appendix The MATLAB GUI PSP processing interface, created by the researcher, opens with the left window in figure A1. Here the user can load the images and subtract the background images. Once this is accomplished, the user will select Align Images which will bring up the control point selection tool on the right of figure A1. Selecting control points allows the user to properly align the images in MATLAB to account for any shift due to the wind tunnel running. Figure A1. PSP Analysis Screenshots. The left window shows the main interface. The right window shows the control point selection tool. Once the images are aligned, the user may input test conditions to generate final pressure and pressure coefficient plots. Selecting Generate Final Images will bring up the PSP Visualization interface shown in the left window of figure A2. This allows the user to visualize intensity ratio, surface pressures or pressure coefficients. The user may select color maps and display range for each image. By selecting Crop Area of Interest the user may select part of the image in order to remove data that is unwanted or to focus on a particular section of the image. In addition to a 2-D color plot of the pressures or pressure coefficients as in the right window of figure A2, the user can also plot these values over a specified path. Selecting Define Line will bring up the right window of figure A2 where the user can define a line (or lines) of interest which will generate a plot of the, in this case, pressure coefficient along that line as seen in the left window of figure A3. Figure A2. PSP Visualization Screenshots. The left window shows the visualization interface where the user can select to visualize the intensity ratio, surface pressures or pressure coefficients. The right window shows the line selector which will generate plots of the surface values, in this case pressure coefficient shown in the left window of figure A3. 9

10 Figure A3. PSP Visualization Output. The left window shows C p plotted along the line defined in the right window of figure A2. The right window shows a 3-D plot of the pressure coefficient plotted above the planform of the wing selected from the image. Vortices can clearly be seen using this three dimensional plot. The final output of the PSP visualization interface is a three dimensional plot of the surface values, in this case, pressure coefficients over the area of the wing planform which was selected from the main image. Both this and the line plot are useful for analyzing the lift distribution over the surface of interest. In the case of the line plot a strong vortex can be seen on the inner part of the wing; however, in the three dimensional plot we can see that this outboard vortex is actually stronger closer to the leading edge of the wing while the inboard vortex is larger, though not as strong. 10

11 Acknowledgments The authors would like to thank Mr. Jeff Falkenstine, USAFA, for model fabrication and construction, Mr. Ken Ostasiewski, USAFA, for wind tunnel operation and PSP application to model and SSgt Danny Washburn for Water Tunnel Operation. Thank you to Mr. Barry Hellman at AFRL for the support for this project. References 1 Doupe, Lt. C., Sponable, J., Zweber, J., Cohn, R., Fully Reusable Access to Space Technology, AFRL-PR-ED- TP , 22 nd National Space Symposium, Colorado Springs, CO, April Jordan, H., FAST program seeks to mature hypersonic air vehicles and space launch technologies. 12 Feb Wright-Patterson Air Force Base. < 29 Jan Miller, A., Mills, R., Post, M., and Hellman, B., Experimental Investigation of a Suborbital Reusable Booster with a Body Flap, AIAA Paper , 49th AIAA Aerodynamic Sciences Meeting, Orlando, FL, January Jackson, D., Pressure Sensitive Paint, < 05 December Oglesby, D., Puram, C., and Upchurch, B., Optimization of Measurements With Pressure Sensitive Paints, NASA TM Department of Aeronautics: Water Tunnel < aero_research_center/facilities/water.cfm>. 11 March