Dawei Liu

Investigation on the Multi-Objective Optimization of Supercritical Airfoil Based on Nondominated Sorting Genetic Algorithm

This paper aimed to investigate the multi-objective optimization method of supercritical airfoil. To achieve the optimal design of supercritical airfoil Rae2822, an improved NSGA-2 (Nondominated Sorting Genetic Algorithm) method was utilized, while the cross-operator and adaptive-variation operator were introduced to improve the convergence speed of the algorithm. During the optimization, the airfoil parametric modeling was achieved based on the Bezier-Bernstein method, and the objective function was obtained through solving the N-S equations. Considering the parallel computation characteristics of the algorithm, the computation was conducted in large-scale Linux computer system to reduce the solving time. Optimization results showed that the undominate solution with high quality obtained through the NSGA-2 method distributed evenly, which provided the designer a wider choosing space. It was also showed that the multi-objective optimization method presented in this paper was feasible and reliable.

INVESTIGATION ON THE CORRELATION OF CFD AND EFD RESULTS FOR A SUPERCRITICAL WING

The correlation of EFD (experimental fluid dynamics) and CFD (computational fluid dynamics) results for a typical supercritical wing was investigated to predict its aerodynamic characteristics well in this paper. A wind tunnel test was conducted in European Transonic Wind tunnel (ETW) to obtain the pressure distribution on a typical supercritical wing test model; also integrated lift force coefficients and pitching moment coefficients were achieved. Computational results were obtained through the debugged RANS solver based on the experimental results. It was showed that a better correlation of EFD and CFD pressure distribution could be achieved, through debugging the CFD method with EFD results, including investigations of grid divergence, turbulence model influence and test model deformation simulation. Test model deformation simulation could apparently improve the correlation of pressure distribution for outboard wing sections. The optimization of turbulence model parameters based on the generic algorithm could improve the correlation of CL~α curve at small angles, but the prediction of separation point was poorly correlated with EFD results. However, the turbulence model parameters optimization implied a promising way to further improve the correlation for supercritical wing especially for multipoint optimization.

Numerical Investigation on the Support Interference of Vane Support System in High Speed Wind Tunnels

The vane support system (VSS) causes less distortion to the rear fuselage. It is an ideal support mechanism for large aircraft wind tunnel tests. Investigating on its support interference will contribute much to the conduction and development of VSS tests. This paper presents a numerical investigation on the support interference of VSS in high speed wind tunnel using Ty154 model. CFD results are compared with experiment results to make the numerical method reliable. Then Ty154 model with and without vanes are simulated. The magnitude of interference achieved is quite small. Its interference characteristics vary with Mach numbers and incidences. The vanes would also do some interference to the flow field.

Numerical Investigation on the Reynolds Number Effects of Supercritical Wing

Supercritical wing is widely used in modern transport aircrafts, whose aerodynamic characteristics are rather sensitive to Reynolds number. The aerodynamic design of large airplane mainly depends on the wind tunnel, which is time consuming and costly especially for high Reynolds number test. This paper aimed to investigate the transonic Reynolds number effects of supercritical wing by numerical simulation method. Flows over a supercritical wing were numerically investigated, the three-dimensional Navier-Stokes equations were solved with structure grids by utilizing the Spalart-Allmaras (S-A) turbulence model, with Reynolds numbers varying from 3.0×106 to 50×106 per mean aerodynamic chord and attack angles from-2 degree to 8 degree at cruise Mach numbers 0.76. Simulated results showed that the aerodynamics of supercritical wing including lift, drag and pitch moment changed linearly with the logarithm of Reynolds number, and the upper surface pressure distribution are much more sensitive than lower surface.

Reynolds Number Effect Investigation of Shock Wave on Supercritical Airfoil

The supercritical airfoil has been widely applied to large airplanes for sake of high aerodynamic efficiency. But at transonic speeds, the shock wave on upper surface of supercritical airfoil may induce boundary layer separation, which would change the aerodynamic characteristics. The shock characteristics such as location and intensity are sensitive to Reynolds number. In order to predict aerodynamic characteristics of supercritical airfoil exactly, the Reynolds number effects of shock wave must be investigated.The transonic flows over a typical supercritical airfoil CH were numerically simulated with two-dimensional Navier-Stokes equations, and the numerical method was validated with test results in ETW(European Transonic Windtunnel). The computation attack angles of CH airfoil varied from 0oto 8o, Mach numbers varied from 0.74 to 0.82 while Reynolds numbers varied from 3×106 to 50×106 per airfoil chord. It is obvious that shock location moves afterward and shock intensity strengthens as Reynolds number increasing. The similar curves of shock location and intensity is linear with logarithm of Reynolds number, so that the shock location and intensity at flight condition could be extrapolated from low Reynolds number.

Influencing Factors Analysis of the Shock-Induced Separation for Supercritical Airfoil

The shock-induced separation easily occurred on the upper surface of supercritical airfoil at transonic speeds, which would change the aerodynamic characteristics. The problem of the shock-induced separation was not solved completely for the complicated phenomena and flow mechanism. In this paper, the influencing factors of shock-induced separation for supercritical airfoil CH was analyzed at transonic speeds. The Navier-Stokes equations were solved, in order to investigate influence of different attack angles, Mach numbers and Reynolds numbers. The computation attack angles of CH airfoil varied from 0oto 7o, Reynolds numbers varied from 5×106 to 50×106 per airfoil chord while Mach number varied from 0.74 to 0.82. It was shown that the shock-induced separation was affected by attack angles, Mach numbers and Reynolds numbers, but the influence tendency and areas were quite different. The shock wave location and intensity were affected by the three factors, and the boundary layer thickness was mainly affected by Reynolds number, while the separation structure was mainly determined by the attack angle and Mach number.

Engineering Extrapolation to Flight Reynolds Number for Supercritical Airfoil Pressure Distribution Based on CFD Results

Pressure distribution of supercritical airfoil at flight Reynolds number could not be fully simulated except in cryogenic wind tunnel such as NTF (National Transonic Facility) and ETW (European Transonic Wind tunnel), which is costly and time resuming. This paper aimed to explore an engineering extrapolation to flight Reynolds number from low Reynolds number wind tunnel data for supercritical airfoil pressure distribution. However, the extrapolation method requiring plenty of data was investigated based on the CFD results for the reason of low cost and short period. Flows over a typical supercritical airfoil were numerically simulated by solving the two dimensional Navier-Stokes equations, with applications of ROE scheme spatial discretization and LU-SGS time march. Influence of computational grids convergence and turbulent models were investigated during the process of simulation. The supercritical airfoil pressure distribution were obtained with Reynolds numbers varied from 3.0×106 to 30×106 per airfoil chord, angles of attack from 0 degree to 6 degree and Mach numbers from 0.74 to 0.8. Simulated results indicated that weak shock existed on the upper surface of supercritical airfoil at cruise condition, that the shock location, shock strength and trailing edge pressure were dependent of Reynolds number, attack angles and Mach numbers. A similar parameter describing the Reynolds number effects factors was obtained by analyzing the relationship of shock wave location, shock front pressure and trailing edge pressure. Based on the similar parameter, airfoil pressure distribution at Reynolds number 30×106 was obtained by extrapolation. It was shown that extrapolated result compared well with simulated result at Reynolds number 30×106, implying that the engineering method was at least promising applying to the extrapolation of low Reynolds number wind tunnel data.

Numerical Investigation on Shock-Induced Separation Structure of Supercritical Airfoil

The supercritical airfoil has been widely applied to large airplanes for sake of high aerodynamic efficiency. But at transonic speeds, the complicated shock-induced separation on the upper surface of supercritical airfoil will change the aerodynamic characteristics. The transonic flows over a typical supercritical airfoil CH were numerically investigated in this paper, in order to analyses different shock-induced separation structure. The two-dimensional Navier-Stokes equations were solved with structure grids by utilizing the S-A turbulence model. The computation attack angles of CH airfoil varied from 0oto 4o, Mach numbers varied from 0.74 to 0.82 while Reynolds numbers varied from 3×106 to 50×106 per airfoil chord. It is shown that with the attack angle increases, the separation bubble occurred on the upper surface first, then the trailing-edge separation occurred, the trailing-edge would separate totally at last. The different separation structure would result in different pressure coefficient distribution and boundary layer thickness.