AERODYNAMIC CHARACTERISTICS OF THE MISSILES MOVABLE WING IN THE PRESENCE OF THE FIXED WING

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1 AERODYNAMIC CHARACTERISTICS OF THE MISSILES MOVABLE WING IN THE PRESENCE OF THE FIXED WING ALİ AKGÜL ROKETSAN Missile Industries Inc., Ankara-Türkiye, roketsan.com.tr EMRAH GÜLAY ROKETSAN Missile Industries Inc., Ankara-Türkiye, roketsan.com.tr JOVAN ISAKOVIĆ Military Technical Institute, Belgrade, SLOBODAN MANDIĆ Military Technical Institute, Belgrade, Abstract: In order to investigate the effect of canards on the wing aerodynamic performance, Computational Fluid Dynamics (CFD) simulations and wind tunnel experiments were performed at Mach number 0.6 and 1.9. Cylindrical body with ogival nose, wing section and canard geometry is used in this study. By changing span dimension and the location of the canard on the body, 5 different body-wing-canard and 1 body-wing configuration are obtained. Wind tunnel measurements and computations of normal force, hinge and bending moment coefficients are done for those configurations. The CFD calculations are performed by the FLUENT software package. Comparing the results from analysis and the experiment, it shows that predictions are in good agreement. Key words: Missile wing aerodynamics, CFD calculation, Wind Tunnel Experiments, FLUENT. 1. INTRODUCTION The aerodynamic performance associated with the canardwing interaction is of particular interest in modern fighter and missile design. The presence of canard in close proximity to the wing results in a highly coupled canardwing aerodynamic flowfield which can include downwash/upwash effects, vortex-vortex interactions and vortex-surface interactions. It is observed by experimental and numerical studies that presence of canards reduces the local angle of attack of wings. Because the trailing canard control surface has a smaller local angle of attack, it is more effective at higher control surface deflection and higher angle of attack. Aerodynamic characteristics of canard-wing interaction are investigated intensively by various experimental and numerical studies. Effect of canard position on wing surface pressure is searched for subsonic speeds at various angles of attack [4]. Results show a remarkable increase in the wing suction peak in the presence of canard. In another study presented that use of canards can delay wing vortex breakdown at high angle of attack and result in increased lift [5]. In order to investigate the effect of canards on the wing aerodynamic performance, Computational Fluid Dynamics (CFD) simulations were performed and compared with the existing experimental data. Cylindrical body with ogival nose, wing section and different canard geometries are is used in this study. By changing span dimension and the location of the canard on the body, different Body-Wing- Canard configurations are obtained. As an experimental data, normal force, hinge and bending moment coefficients of wing at Mach numbers 0.6 and 1.9 for various angles of attack is existing. 2. DESCRIPTION OF MODELS Basic dimensions of the missile models, which are composed of cylindrical body with ogival nose, wing section and different canard geometries, are given in Pictures 1-6. Six different solid models are generated. Main difference between the models is the location and dimension of canards [1]. 8

2 Picture 6. Body Wing Canard-X20-2L/3 Solid Model (BWC-X20-2L/3) Picture 1. Body Wing Solid Model (BW) Hinge moment coefficient reference line distance is 30.98mm from the root chord leading edge. Wing root is used as bending moment reference line. 3. CFD SIMULATION Picture 2. Body Wing Canard-X0-L Solid Model (BWC-X0-L) Viscous computational fluid dynamic simulations were used to calculate the flowfield around the different Body- Wing-Canard configurations. Computations were performed at Mach numbers 0.6 and 1.9 for angle of attack ranging from 0º to 10º. Angle of attack range for the wing alone configuration is between 0º to 20º at the same Mach numbers Solid Model and Computational Mesh Picture 3. Body Wing Canard-X20-L Solid Model (BWC-X20-L) Picture 4. Body Wing Canard-X40-L Solid Model (BWC-X40-L) Six different missile geometries were generated for CFD studies. Each model has different body, wing and canard combination. The both solid model and unstructured hybrid meshes were generated using the GAMBIT software [2],[6]. The computational domain inlet was located 15 model body length upstream from the tip of model nose and the computational domain outlet was located 20 model body length downstream from the model base. Distance between the side boundaries of the computational domain is about 20 model body length. In generating the meshes, boundary layer mesh spacing was used near the model surface. The two-layer zonal model was used for the near-wall equations and the first point off the surface was chosen to give y + value of about layers of prismatic cells were generated to adequately resolve the boundary layer. The remaining part of the solution domain was completely composed of tetrahedral elements. The mesh growth rate was kept below Sample pictures of volume and surface mesh cross-sectional views are shown respectively in Pictures 7-9. Picture 5. Body Wing Canard-X20-L/3 Solid Model (BWC-X20-L/3) Picture 7. Surface and Computational Domain Grid for BW and BWC-X0-L model 9

3 Picture 8. Surface and Computational Domain Grid for BWC-X20-L and BWC-X40-L model Picture 9. Surface and Computational Domain Grid for BWC-X20-L/3 and BWC-X20-2L/3 model 3.2. Flow Solver and Boundary Conditions FLUENT commercial flow solver was used to compute the normal force coefficient, hinge and bending moment coefficients and the flowfield around missile geometries. The density-based, implicit, compressible, unstructuredmesh solver was used. The three-dimensional, Reynolds- Average Navier-Stokes (RANS) equations were solved using the finite volume method. Second-order discretization was used for all flow variables. A modified form of the k- ε two-equation turbulence model (realizable k- ε) was used in this study. This turbulence model solves transport equations for the turbulence kinetic energy, k, and its dissipation rate, ε. The boundary conditions were as follows. Downstream, upstream, and outer radial boundaries were set as far-field (characteristics-based inflow/outflow), with sea-level temperature and pressure free stream conditions (300 K, Pa). All the solid surfaces was modeled as a noslip, adiabatic wall boundary conditions. The model reference length was m which is the wing aerodynamic chord and the reference area was taken as m Solution Strategy CFD simulations were performed in 48 CPUs parallel supercomputer. Each case was started with a lower CFL value of 1.0 and ramped up to 16 during iterations of the simulation. Convergence was determined by tracking the change in the flow residuals and the aerodynamic coefficients during the solution. The solution was deemed converged when the flow residuals had reduced at least 3 orders of magnitude and the aerodynamic coefficients changed less than about 1% over the last 100 iterations. 4. TEST FACILITY AND MEASUREMENTS The T-38 test facility at the Military Technical Institute is a blow-down pressurized wind tunnel with a 1.5m x 1.5m square test section. For subsonic and supersonic tests, the test section is with solid walls, while for transonic tests, a section with porous walls is inserted in the wind tunnel configuration, [3]. Mach number in the range 0.2 to 4.0 can be achieved in the test section, with Reynolds numbers up to 110 million per meter. Mach number is set and maintained to within 0.5% of the nominal value by means of either a flexile nozzle or sidewall flaps and/or sidewall blowoff, depending on the test speed range. Stagnation pressure in the test section can be maintained between 1.1 bars and 15 bars, depending on Mach number, and regulated to 0.3% of nominal value. Run times are in the range 6s to 60s, depending on Mach number and stagnation pressure. Model was supported in the test section by a tail sting mounted on a pitch-and-roll mechanism by which desired aerodynamic angles can be achieved. The facility supports both step-by-step model movement and continuous movement of model (sweep) during measurements. The stagnation pressure in the test section was measured by a Mensor absolute pressure transducer pneumatically connected to a pitot probe in the settling chamber of the wind tunnel. The static pressure in the test section was measured by a transducer of the same type (but lower range) pneumatically connected to an orifice on the test section sidewall. The nonlinearity and hysteresis of the transducers used are typically 0.02% F.S. The stagnation temperature was measured by a RTD probe in the settling chamber. The accuracy of this transducer was approximately ±0.5K. The pitching angle of the model support was measured by a resolver which is a part of the mechanism. Roll angle was measured by an absolute optical encoder into the drive unit. The overall accuracy of measurements of model position, including the calculation of sting deflections under load, was about Aerodynamic forces and moments on the movable wing were measured by a VTI-produced internal, threecomponent strain gauge balance. The accuracy of the balance was approximately 0.20% F.S. 5. RESULTS In this section, the results of the computed and measured normal force, hinge and bending moment coefficients comparisons are presented. All CFD results showed good agreement with the experimental data from VTI T-38 wind tunnel. Each wing-canard configuration results are compared in different sub-sections Body Wing (BW) presented for Body Wing (BW) configuration at Mach 0.6 and Mach 1.9 with respect to angle of attack. CFD computations accurately predicted the normal force, hinge and bending moment coefficients. 10

4 Picture 10. Variation of CN w.r.t Angle of Attack for WB Configuration Picture 13. Variation of CN w.r.t Angle of Attack for WBC-X0-L Configuration Picture 11. Variation of CHM w.r.t Angle of Attack for WB Configuration Picture 14. Variation of CHM w.r.t Angle of Attack for WBC-X0-L Configuration Picture 12. Variation of CBM w.r.t Angle of Attack for WB Configuration Picture 15. Variation of CBM w.r.t Angle of Attack for WBC-X0-L Configuration 5.2. Body Wing Canard-X0-L (BWC-X0-L) 5.3. Body Wing Canard-X20-L (BWC-X20-L) presented for Body Wing Canard-X0-L (BWC-X0-L) configuration at Mach 0.6 and Mach 1.9 with respect to angle of attack. presented for Body Wing Canard-X20-L (BWC-X20-L) configuration at Mach 0.6 and Mach 1.9 with respect to angle of attack. 11

5 Picture 16. Variation of C N w.r.t Angle of Attack for WBC-X20-L Configuration Picture 19. Variation of C N w.r.t Angle of Attack for WBC-X40-L Configuration Picture 17. Variation of C HM w.r.t Angle of Attack for WBC-X20-L Configuration Picture 20. Variation of C HM w.r.t Angle of Attack for WBC-X40-L Configuration Picture 18. Variation of C BM w.r.t Angle of Attack for WBC-X20-L Configuration 5.4. Body Wing Canard-X40-L (BWC-X40-L) presented for Body Wing Canard-X40-L (BWC-X40-L) configuration at Mach 0.6 and Mach 1.9 with respect to angle of attack. Picture 21. Variation of C BM w.r.t Angle of Attack for WBC-X40-L Configuration 5.5. Body Wing Canard-X20-L/3 (BWC-X20-L/3) presented for Body Wing Canard-X20-L/3 (BWC-X20- L/3) configuration at Mach 0.6 and Mach 1.9 with respect to angle of attack. 12

6 Picture 22. Variation of C N w.r.t Angle of Attack for WBC-X20-L/3 Configuration Picture 25. Variation of C N w.r.t Angle of Attack for WBC-X20-2L/3 Configuration Picture 23. Variation of C HM w.r.t Angle of Attack for WBC-X20-L/3 Configuration Picture 26. Variation of C HM w.r.t Angle of Attack for WBC-X20-2L/3 Configuration Picture 24. Variation of C BM w.r.t Angle of Attack for WBC-X20-L/3 Configuration 5.6. Body Wing Canard-X20-2L/3 (BWC-X20-2L/3) presented for Body Wing Canard-X20-2L/3 (BWC-X20-2L/3) configuration at Mach 0.6 and Mach 1.9 with respect to angle of attack. Picture 27. Variation of C BM w.r.t Angle of Attack for WBC-X20-2L/3 Configuration 5.7. Effect of Canard Position on Wing Aerodynamic Performance CFD results of BW, BWC-X0-L, BWC-X20-L and BWC- X40-L configurations are used examine the effect of canard position on the wing aerodynamic performance. Presence of canards decreases the local angle of attack of the wings so the normal force. But as the distance between canard and wing increases, effect of canards on the wings decreases at Mach number 0.6. However 13

7 variation in canard position does not have any effect on wing aerodynamic performance at Mach number 1.9. Picture 28. Effect of Canard Position on the C N at Mach number 0.6 Picture 29. Effect of Canard Position on the CN at Mach number Effect of Canard Size on Wing Aerodynamic Performance CFD results of BW, BWC-X20-L, BWC-X20-2L/3 and BWC-X20-L/3 configurations are used to examine the effect of canard size on the wing aerodynamic performance. As expected, small canard span value has small effect on wing aerodynamic performance. Picture 31. Effect of Canard Size on the C N at Mach number CONCLUSION CFD calculations and wind tunnel measurements were performed for 5 different body-wing-canard and 1 bodywing configuration. Experimental and numerical normal force, hinge and bending moment coefficients are compared with each other. The CFD calculations are performed by the FLUENT software package and experiments are conducted in VTI. Comparing the results from analysis and the experiment, it shows that overall predictions are in good agreement. CFD accurately predicted the normal force, hinge and bending moment coefficient characteristics. However in some cases, there is a difference between numerical and experimental hinge moment coefficients results for angles of attack bigger than 6º at subsonic regime. This situation has to be inspected. In subsonic flow, presence of canard decreases the local angle of attack of wing so the normal force of wing. Also canard causes the center of pressure of wing to move through trailing edge, hence wing hinge moment increases. Similarly in supersonic flow, canard causes the decrease in wing normal force but position and size of canard do not affect the results. Moreover wing pressure center location does not influenced by the presence of canard in supersonic flow. As an extension of this work, it is planned to develop a new methodology for aerodynamic design of split-canard missile using these experimental and numerical results based on equivalent angle of attack method. ACKNOWLEDGMENTS This work is supported by ROKETSAN and VTI as a part of research activity. The authors thank VTI for providing wind tunnel test data. Picture 30. Effect of Canard Size on the C N at Mach number 0.6 References [1] Stojković,S., Mandić,S.: Uticaj destabiliyatora na koeficijent normalne sile i šarnirnog momenta krila, VTI , Belgrade, [2] Fluent Users Guide, Fluent, Inc. 14

8 [3] Elfstrom,G.M., Medved,B.: The Yugoslav 1.5m trisonic blowdown wind tunnel, AIAA Paper CP, [4] Soltani,M.R., Askarif,F., Davari,A.R., Nayabzadeh,A.: Effect of Canard Position on Wing Surface Pressure, Scientia Iranica, April 2010, Vol.17, No.2, pp [5] Tu,E.L.: Numerical Study of Steady and Unsteady Canard-Wing-Body Aerodynamics, August 1996, NASA Technical Memorandum [6] Gülay,E., Akgül,A., Isaković,J., Mandić,S.: Computational Fluid Dynamics and Experimental Investigaion of Wrap-Around-Fins Missile Rolling Moment, Scientific Technical Review, 2011, Vol.61, No.3-4, pp