Supersonic Wing Design Research Articles

Aerodynamic Optimization Design of Supersonic Wing Based on Discrete Adjoint

Reducing fuel consumption and improving the economy by effectively reducing cruising drag is the main objective of the aerodynamic design of supersonic civil aircraft. In this paper, the aerodynamic optimization design system based on the Reynolds-Averaged Navier–Stokes (RANS) equation and discrete adjoint theory is applied to supersonic wing design. Based on this system, a single-point optimization design study of aerodynamic drag reduction in cruise conditions was carried out for two typical supersonic wing layouts, subsonic leading edge and supersonic leading edge, and the drag reduction reached 3.78% and 4.53%, respectively. The aerodynamic design characteristics of different types of supersonic wings were explored from the perspectives of wing load, twist angle distribution, pressure distribution, airfoil shape characteristics, and flow field characteristics. The optimization results show that the drag reduction of the subsonic leading edge configuration is dominated by the induced drag, while the optimizer mainly focuses on reducing the shock wave drag for the supersonic leading edge configuration. By comparing the sensitivity analysis of lift and drag coefficients to airfoil deformation with the optimization results, the optimized dominant directions of two types of supersonic wings are qualitatively analyzed. The derivatives obtained from discrete adjoint equations are useful to elaborate the design tendency and the reason for the trade-off generation of supersonic wings under specific layouts and engineering constraints, which provides a reference for the design of supersonic wings in the future.

高効率手法を適用した静粛超音速航空機の多分野融合最適設計

In this study, high efficient design tool is developed with several informatics approaches for Multidisciplinary Design Optimization (MDO) and knowledge discovery of supersonic wing design. In this design, Multi-Objective Genetic Algorithm (MOGA) is applied as an optimizer, while Kriging model is also used to reduce computational cost. To obtain the information of the design space, functional ANalysis Of VAriance (ANOVA) and Parallel Coordinate Plot (PCP) is applied. For Kriging model construction, 107 sample points are evaluated. This tool is applied to the multidisciplinary design problem of supersonic wing. The objective functions are to maximize lift to drag ratio and to minimize sonic boom intensity at supersonic cruise, and to minimize wing weight. According to the results, there are trade-off relationships among three objective functions. The ANOVA results indicate that the cambers of the wing section at the root and the kink have an influence on the lift to drag ratio, the inner wing sweep back angle affects the sonic boom intensity, and the camber of wing section at the kink and aspect ratio affect the wing weight. The design space information could be visualized quantitatively from the sampling results with PCP technique. Since the design space exploration using MOGA is carried out based on Kriging surrogate models, the proposed MDO process is effective in terms of computational cost.

데이터 마이닝을 통한 공력설계공간 지식습득

본 연구에서는 공력설계공간의 지식습득을 위해 분산해석법과 자기조직화지도의 이용을 제안하였다. 이 기법들은 각각의 설계변수가 목적함수에 미치는 영향을 예측 가능하게 한다. 더욱이, 분산해석법은 설계변수들의 상호관계가 목적함수에 미치는 영향도 예측 가능하게 하며, 자기조직화지도는 목적함수들 사이에 어떠한 trade-off관계가 있는지도 예측 가능하게 한다. 본 논문에서 72개의 설계변수와 4개의 목적함수를 가진 초음속 날개 설계의 결과에 대하여 이들의 데이터 마이닝 기법들을 적용하였다. Two data mining techniques, analysis of variance (ANOVA) and self-organizing map (SOM), are applied to knowledge discovery in aerodynamic design space. These methods make it possible to identify the effect of each design variable on the objective functions. Furthermore, ANOVA shows the effect of interaction between design variables on the objective function and SOM visualizes the trade-off among objective functions. Present methods are applied to the result of the supersonic wing design which includes 72 design variables and 4 objective functions.

704 多目的超音速翼設計パレート解の自己組織化マツプ(O.S.7-1 複雑CFD-1)(O.S.7 複雑CFD)

704 多目的超音速翼設計パレート解の自己組織化マツプ(O.S.7-1 複雑CFD-1)(O.S.7 複雑CFD)

Inverse optimization of supersonic wing design with twist specification

A design optimization system for a supersonic wing has been developed. The present design system consists of the supersonic inverse design method and the Multiple-Objective Genetic Algorithm (MOGA) code for target pressure optimization. Furthermore, a wing twist specification technique is introduced to the present inverse design method by taking advantage of the non-uniqueness of three-dimensional aerodynamic inverse problems. With this method, a designer can select a wing that is easy to be manufactured and aerodynamically optimal. The designed wing shows better performance than the second baseline wing of National Aerospace Laboratory's (NAL) scaled supersonic experimental airplane program.

Euler/Navier-Stokes Optimization of Supersonic Wing Design Based on Evolutionary Algorithm

This paper presents aerodynamic shape optimization of a supersonic wing for supersonic civil transportation (SST) using an Evolutionary Algorithm (EA) coupled with an Euler/Navier-Stokes code. To overcome enormous computational time necessary for the design, aerodynamic evaluations are parallelized on Numerical Wind Tunnel (NWT) at National Aerospace Laboratory, a parallel vector machine with 166 processing elements. Parallelization of function evaluations in EA is straightforward and its performance is extremely good since most of computational time is used by flow computations. The design result indicates that the present EA successfully minimizes both the induced drag and the volume wave drag in the given design space.

Inverse Design of Supersonic Airfoils Using Integral Equations

Introduction T HE next-generation supersonic transport (SST) is under consideration today because the Concorde has been in service for 20 years and is not fully successful economically. There are many ongoing researches in this Ž eld in the United States, Europe, and Japan. To guarantee its economic success, the next-generationSST is required to have higher lift-to-drag ratio than that of the Concorde. To achieve this goal, developmentof a new design technique for supersonicwings is of great interest. For the transonicwing design,Takanashi1 proposedan inversemethod that uses the “residualcorrection” concept and has shown successful design results. In this Note, this inverse method is extended to the supersonic wing design. In the supersonic  ow, the governing equation can be linearized. Therefore,the integrationbecomessimpler than thatof the transonic case, but the integration region must be selected carefully because the region contributing to the point on the wing surface is limited to the Mach forecone. Sample results of two-dimensional airfoil designs conŽ rm the validity of the present method.

Euler/Navier-Stokes optimization of supersonic wing design based on evolutionary algorithm

Euler/Navier-Stokes optimization of supersonic wing design based on evolutionary algorithm

The influence of leading-edge load alleviation on supersonic wing design

A theoretical and experimental program to assess the effect of leading-edge load constraints on wing design and performance was conducted. For a planform characterized by a highly swept leading edge on the inboard region, linear theory was used to design camber surfaces which produced minimum drag-due-to-lift at the design lift coefficient of 0.08 and a design Mach number of 2.4. In an effort to delay the formation of leading edge vortices which often occur on highly swept wings, two approaches were used in the design criteria to limit the loadings on the leading edge. One wing was constrained to have the normal Mach number less than one everywhere along the leading edge and the second wing was constrained to have a pressure coefficient of zero on the leading edge. Force tests were run on the two constrained wings, on a flat reference wing and on an optimized wing with no leading edge constraints. All wings had identical planforms and thicknesses and were tested over a range of Mach numbers from 1.8 to 2.8 and a range in angles of attack from -5 deg to 8 deg. A comparison of the experimental performance of these four models is shown. Correlations of these results with theoretical predictions and flow visualization photographs are also included.