Juntao Xiong

Computational Analysis of Conical Forebody Flow at High Alpha with Transitional Model

Borui Zheng∗ Northwestern Polytechnical University, 710072 Xi’an, People’s Republic of China Zhongyi Wang Beijing University of Aeronautics and Astronautics, 100191 Beijing, People’s Republic of China Chao Gao Northwestern Polytechnical University, 710072 Xi’an, People’s Republic of China Hong Fang Beijing Institute of Nearspace Vehicle’s System Engineering, 100076 Beijing, People’s Republic of China and Juntao Xiong, Feng Liu, and Shijun Luo University of California, Irvine, Irvine, California 92697-3975

Investigation of NACA 0012 Airfoil Periodic Flows in a Transonic Wind Tunnel

An experimental study of static and dynamic pressure distributions over a NACA 0012 airfoil in the two-dimensional test section 0.8 × 0.4-meter of a transonic wind tunnel is presented. Tests are conducted at freestream Mach numbers from 0.35 to 0.89, and angle of attack between −2.0 and 15, and Reynolds number based on the airfoil chord length of 3.0 million. Tests are performed at boundary-layer transition free and transition fixed at 5% chord length conditions. Firstly, sublimating naphthalene visualization is performed to locate the transition position of the boundary layer over the airfoil at upper surface. Secondly, the chordwise static pressure distributions with transition fixed are validated by well-documented data in the literature. Thirdly, the dynamic pressure data at Mach numbers from 0.60 to 0.89 are analyzed. Finally, a numerical simulation at freestream Mach numbers from 0.82 to 0.89 are done. The unsteady computation results show a close trends as experimental measurements.

Experimental Study of Shock Wave Oscillation on SC(2)- 0714 Airfoil

An experimental study executed in the NPU NF-6 transonic wind tunnel on shock oscillations over SC(2)-0714 airfoil is presented. Tests are conducted at freestream Mach number from 0.72 to 0.82, Reynolds number based on the airfoil chord change from approximating 3.0 million to 5.0 million with transition strip fixed at 28% chord length. Limited span wise pressure distributions are obtained to verify the two-dimensionality of the flow field. The unsteady pressures on upper surfaces are analyzed and transonic buffet onset points are given. The results show that the phenomenon is essentially two dimensional in the middle of the model. The effect of the Reynolds number is negligible when the difference is about 1 million. Finally Lee’s self-sustained shock oscillation model is examined.

Computation of NACA0012 Airfoil Transonic Buffet Phenomenon with Unsteady Navier-Stokes Equations

A computational study of transonic buffet phenomenon over NACA0012 airfoil is presented using unsteady Reynolds-averaged Navier-Stokes Equations. Computations are conducted at freestream Mach numbers from 0.82 to 0.89 at zero angle of attack for Reynolds numbers of 3.0 and 10.0 million with transition fixed at 5% chord length. Numerical simulation reveals the transonic buffet occurs in a narrow range of freestream Mach numbers of 0.85-0.87. It is close to the range of freestream Mach numbers of 0.84-0.86 in which multiple solutions occur for the steady, inviscid flow. 1 The unsteady flow fields around the airfoil are examined and characterized. The shock motion frequency is found to depend on the distance between time-averaged shock wave position and trailing edge. The shock motion frequency is larger when the time-averaged shock position is closer to the trailing edge. The computational results are verified using present transonic wind tunnel experiment data. Finally the effect of a trailing edge splitter for buffet suppression is studied.

Multiple Solutions and Buffet of Transonic Flow over the NACA0012 Airfoil

Multiple solutions of the small-disturbance potential equation and the full potential equation were known for the NACA0012 airfoil in a certain range of transonic Mach numbers and at zero angle of attack. However the multiple solutions for this particular airfoil were not observed using Euler or Navier-Stokes equations under the above flow conditions. In the present work, both the Unsteady Reynolds-Averaged Navier-Stokes (URANS) computations and transonic wind tunnel experiments are performed to further study the problem. The results of the two methods reveal that buffet appears under certain Reynolds number and transition position in a narrow Mach number range where the potential flow methods predict multiple solutions. Boundary layer displacement thickness computed from URANS at the same flow condition is used to modify the geometry of the NACA0012 airfoil. Euler equations are then solved for the modified geometry. The results show that the addition of the boundary layer displacement thickness creates multiple solutions for the NACA0012 airfoil. Global linear stability analysis is also performed on the original airfoil and the modified airfoil. This shows a close relationship between the viscous unsteady shock buffet phenomenon of transonic airfoil flow and the existence of multiple solutions of the external inviscid flow.

Analysis of the Effects of Wall Interference on 0.4×0.8-Meter Transonic Wind Tunnel Airfoil Tests

This article presents analytical and computational analysis of the influence of the wind tunnel walls on transonic airfoil tests. The goal is to quantify the effects of the wind tunnel walls and characterize the key parameters affecting the flow field in order to construct a numerical model to realize experimental data corrections. Firstly, the experimental data are compared with well-documented data in the literature to perform preliminary corrections. Secondly, an analytical method based on small perturbation theory is used to estimate the wall interference. Finally, 2D and 3D numerical investigations of the influence of the wind tunnel walls on the transonic airfoil tests are performed to confirm the wall interference corrections. Results show the experiment Mach number in NF-6 wind tunnel can be corrected with Murthy sidewall correction method.

Non-unique Numerical Solutions and Their Stability of the Euler Equations for Transonic Flow over Airfoils

Multiple numerical solutions of inviscid transonic flow over several airfoils exist at certain Mach number and angle of attack. Global linear stability analysis of the multiple solutions is conducted in this paper. Linear perturbation equations of the Euler equations around a steady-state solution are formed, an eigenvalue problem is then constructed using the modal analysis approach, then it is discretized using a numerical scheme that is similar to the JST scheme for the Euler equations. Only a small portion of the eigen spectrum is needed and thus can be found efficiently by using implicit restarted Arnoldi’s algorithm. The eigenvalue that has the largest real part determines stability of the steady-state solution, the corresponding eigenmode can also be found. Multiple numerical solutions and their stability characteristics are studied for the Hafez airfoils with flat and wavy surfaces, original and modified NACA 0012 airfoil. For the Hafez airfoils, the numerical solutions and stability are consistent with previous studies in the literature, which suggest the validity of the present method. Analysis of the NACA 0012 airfoil indicates stability of symmetric solutions of the Euler equations at conditions where buffet is found from unsteady Navier-Stokes equations. Euler solutions of the same airfoil but modified to include the displacement thickness of the boundary layer computed from the Navier-Stokes equations, however, exhibit instability based on the present linear stability analysis. The asymmetry J78 airfoil exhibits different type of multiple solutions and stability mechanism.

Boundary-Layer Transition Effects on Aerodynamic Characteristics of AGARD-B Model

The AGARD-B calibration model is tested in a continuous closed-circuit 0.8 × 0.6 transonic wind tunnel. The lift, drag, pitching-moment and the base pressure are measured at freestream Mach number of 0.60 and 0.75, the Reynolds number based on the wing mean aerodynamic chord about 2 × 10 and angle of attack between −6 and 14. Both free boundary-layer transition and transition fixed near the wing leading edge and the fuselage apex are performed. The fixed and free boundary-layer transitions over the upper surface of the wing are visualized by a sublimating naphthalene technique at angles of attack −2, 0, 2 and 6. The effects of the boundary-layer transition on the aerodynamic characteristics are analyzed, and the results are compared with data available in the literature.