考虑转捩影响的DLR_F4翼身组合体阻力计算

针对现代民用飞机设计巡航阻力预测中不考虑转捩影响很难完善计算与试验对比的问题,采用固定转捩和全湍流方法完成DLR_F4翼身组合体阻力计算,并与试验及相关文献结果进行对比.结果表明:采用设置有小范围层流区的固定转捩计算与相应全湍流计算得到的压力因数分布基本一致,二者之间总阻力的差量主要表现在黏性阻力方面;在大多数迎角下,...     

Coupled adjoint aerostructural wing optimization using quasi-three-dimensional aerodynamic analysis

This paper presents a method for wing aerostructural analysis and optimization, which needs much lower computational costs, while computes the wing drag and structural deformation with a level of accuracy comparable to the higher fidelity CFD and FEM tools. A quasi-three-dimensional aerodynamic solver is developed and connected to a finite beam element model for wing aerostructural optimization. In a quasi-three-dimensional approach an inviscid incompressible vortex lattice method is coupled with a viscous compressible airfoil analysis code for drag prediction of a three dimensional wing. The accuracy of the proposed method for wing drag prediction is validated by comparing its results with the results of a higher fidelity CFD analysis. The wing structural deformation as well as the stress distribution in the wingbox structure is computed using a finite beam element model. The Newton method is used to solve the coupled system. The sensitivities of the outputs, for example the wing drag, with respect to the inputs, for example the wing geometry, is computed by a combined use of the coupled adjoint method, automatic differentiation and the chain rule of differentiation. A gradient based optimization is performed using the proposed tool for minimizing the fuel weight of an A320 class aircraft. The optimization resulted in more than 10 % reduction in the aircraft fuel weight by optimizing the wing planform and airfoils shape as well as the wing internal structure.     

基于近似模型的车身气动外形优化

以英国汽车工业研究协会(Motor Industry Research Association,MIRA)阶背模型为基本模型,用参数化建模方法建立其纵向对称面的二维模型.运用优化拉丁超立方方法对每组参数化方案生成600组样本点;将MATLAB与Gambit结合,自动快速生成其网格模型;用FLUENT计算每个样本点的气动阻力.建立径向基神经网络(Radial Basis Function Neural Network,RBFNN)近似模型,以阻力最小为优化目标,采用多岛遗传算法优化外形参数;对优化后的结果进行数值模拟,结果表明阻力减少31.9%.三维验证结果表明:二维优化结果不能完全代表三维结果,直接进行三维优化设计的效果更好.     

Numerical sensitivity analysis for aerodynamic optimization: A survey of approaches

The calculation of the derivatives of output quantities of aerodynamic flow codes, commonly known as numerical sensitivity analysis, has recently become of increased importance for a variety of applications in flow analysis, but the original motivation came from the field of aerodynamic shape optimization. There the large numbers of design variables needed to parameterize surfaces in 3D necessitates the use of gradient-based optimization algorithms, and hence efficient and accurate evaluation of gradients. In this context over the last 20 years a variety of approaches have been developed to supply these gradients, raising particular challenges that have required novel algorithms. In this paper, we examine the historical development of these approaches, describe in some detail the theoretical background of each major method and the associated numerical techniques required to make them practical in an engineering setting. We give examples from our own experience and describe what we consider to be the state-of-the-art in these methods, including their application to optimization of complex 3D aircraft configurations.     

乘波体组合高压捕获翼构型的性能分析

针对高速飞行器大容积、高升力、低阻力和高升阻比的设计需求,提出高压捕获翼(High pressure zone Capture Wing,HCW)的概念.在高速巡航条件下,合理配置HCW可以充分利用来流压缩产生的高压气体,从而提高飞行器升力;HCW采用与来流平行的薄板装置,其附加阻力较小,可以大幅提高升阻比.采用CFD分析工具,比较不同容积的乘波体构型与HCW组合前后的气动性能.结果表明,在不同容积构型下升阻比均有明显提高,最小提升量可达10%.此外,容积越大,升力和升阻比增加效果越明显.     

跨声速机翼阻力分解的一种改进方法

提出黏性区域探测器的一种改进形式，并用于捕捉激波和翼梢涡的熵增阻力；给出尾迹平面的可压缩涡动力学诱导阻力表达式，并与基于热力学的诱导阻力对比。在跨声速来流状态下，对ONERA M6和某民用飞机巡航状态下的机翼阻力进行分解，同时分析该民用飞机机翼安装翼梢小翼前、后的远场阻力构成。结果表明：新的区域探测器合理可靠，黏性阻力与伪熵阻力的计算结果更加准确；2种诱导阻力计算方式的计算结果一致，但基于涡动力学的诱导阻力计算方法受积分平面位置的影响更小；安装翼梢小翼基本不影响整个流场的黏性阻力，减阻的主要效果体现为诱导阻力的减小。     

Aerodynamic shape optimization of supersonic aircraft configurations via an adjoint formulation on distributed memory parallel computers

This work describes the application of a control theory-based aerodynamic shape optimization method to the problem of supersonic aircraft design. A high fidelity computational fluid dynamics (CFD) algorithm modelling the Euler equations is used to calculate the aerodynamic properties of complex three-dimensional aircraft configurations. The design process is greatly accelerated through the use of both control theory and parallel computing. Control theory is employed to derive the adjoint differential equations whose solution allows for the evaluation of design gradient information at a fraction of the computational cost required by previous design methods. The resulting problem is then implemented in parallel using a domain decomposition approach, an optimized communication schedule, and the Message Passing Interface (MPI) Standard for portability and efficiency. In our earlier studies, the serial implementation of this design method, was shown to be effective for the optimization of airfoils, wings, wing–bodies, and complex aircraft configurations using both the potential equation and the Euler equations. In this work, our concern will be to extend the methodologies such that the combined capabilities of these new technologies can be used routinely and efficiently in an industrial design environment. The aerodynamic optimization of a supersonic transport configuration is presented as a demonstration test case of the capability. A particular difficulty of this test case is posed by the close coupling of the propulsion/airframe integration.     

Three-dimensional aerodynamic computations on unstructured grids using a Newton-Krylov approach

A Newton-Krylov algorithm is presented for the compressible Navier-Stokes equations in three dimensions on unstructured grids. The algorithm uses a preconditioned matrix-free Krylov method to solve the linear system that arises in the Newton iterations. Incomplete factorization is used as the preconditioner, based on an approximate Jacobian matrix after the reverse Cuthill-McKee reordering of the unknowns. Several approximate viscous operators that involve only the nearest neighboring terms are studied to reduce the cost of preconditioning. The performance of the algorithm is demonstrated through numerical studies of the ONERA M6 wing and the DLR-F6 wing-body configuration. A ten-order-of-magnitude residual reduction for the wing and wing-body configurations can be obtained with a computing cost equivalent to 5500 and 8000 function evaluations, respectively, on grids with a half million nodes.     

Global optimization using a multipoint type quasi-chaotic optimization method

This paper proposes a new global optimization method called the multipoint type quasi-chaotic optimization method. In the proposed method, the simultaneous perturbation gradient approximation is introduced into a multipoint type chaotic optimization method in order to carry out optimization without gradient information. The multipoint type chaotic optimization method, which has been proposed recently, is a global optimization method for solving unconstrained optimization problems in which multiple search points which implement global searches driven by a chaotic gradient dynamic model are advected to their elite search points (best search points among the current search histories). The chaotic optimization method uses a gradient to drive search points. Hence, its application is restricted to a class of problems in which the gradient of the objective function can be computed. In this paper, the simultaneous perturbation gradient approximation is introduced into the multipoint type chaotic optimization method in order to approximate gradients so that the chaotic optimization method can be applied to a class of problems for which only the objective function values can be computed. Then, the effectiveness of the proposed method is confirmed through application to several unconstrained multi-peaked, noisy, or discontinuous optimization problems with 100 or more variables, comparing to other major meta-heuristics.     

应用等离子体实现流场主动控制技术的研究

飞行器设计的一个重要目标,就是要优化流场分布,减少阻力,增加升力,提高飞行器的升阻比,飞行器在高、亚音速巡航时,摩擦阻力超过了总阻力的一半,1%阻力的降低,将大约提高10%的有效负荷或飞行距离,传统的方法,特别是先进翼型的普遍采用,大大提高了飞行器的飞行性能,带来了巨大的经济效益和社会效益.但是,随着设计要求的进一步提高,传统的设计方法越来越显示了它的局限性.目前,国际上开始考虑通过等离子体对流场特性的影响来达到减阻这一目的.该文介绍了一个大气压下辉光放电等离子体发生装置的研制方法,并通过已成功研制的等离子体发生装置主动产生表面等离子体,揭示表面等离子体对流场以及电磁场的影响.     

Multi-fidelity wing aerostructural optimization using a trust region filter-SQP algorithm

A trust region filter-SQP method is used for wing multi-fidelity aerostructural optimization. Filter method eliminates the need for a penalty function, and subsequently a penalty parameter. Besides, it can easily be modified to be used for multi-fidelity optimization. A low fidelity aerostructural analysis tool is presented, that computes the drag, weight and structural deformation of lifting surfaces as well as their sensitivities with respect to the design variables using analytical methods. That tool is used for a mono-fidelity wing aerostructral optimization using a trust region filter-SQP method. In addition to that, a multi-fidelity aerostructural optimization has been performed, using a higher fidelity CFD code to calibrate the results of the lower fidelity model. In that case, the lower fidelity tool is used to compute the objective function, constraints and their derivatives to construct the quadratic programming subproblem. The high fidelity model is used to compute the objective function and the constraints used to generate the filter. The results of the high fidelity analysis are also used to calibrate the results of the lower fidelity tool during the optimization. This method is applied to optimize the wing of an A320 like aircraft for minimum fuel burn. The results showed about 9 % reduction in the aircraft mission fuel burn.     

An experimental study of pilots' control characteristics for flight of an STOL aircraft in backside of drag curve at approach and landing.

In general, most vehicles can be modelled by a multi-variable system which has interactive variables. It can be clearly shown that there is an interactive response in an aircraft's velocity and altitude obtained by stick control and/or throttle control. In particular, if the flight conditions fall to backside of drag curve in the flight of an STOL aircraft at approach and landing then the ratio of drag variation to velocity change has a negative value (delta D/delta u less than 0) and the system of motion presents a non-minimum phase. Therefore, the interaction between velocity and altitude response becomes so complicated that it affects to pilot's control actions and it may be difficult to control the STOL aircraft at approach and landing. In this paper, experimental results of a pilot's ability to control the STOL aircraft are presented for a multi-variable manual control system using a fixed ground base simulator and the pilot's control ability is discussed for the flight of an STOL aircraft at backside of drag curve at approach and landing.     

Aeroelastic simulations using gridless boundary condition and small perturbation techniques

This paper examines the use of stationary Cartesian mesh for non-linear flutter computations involving complex geometries. The surface boundary conditions are implemented using reflected points which are determined via a gridless approach. The method uses a cloud of nodes in the vicinity of the surface to get a weighted-average of the flow properties using radial basis functions. To ensure computational efficiency and for local grid refinements, multigrid computations within an embedded grids framework are used. As the displacements of moving surfaces from their original position are typically small for flutter problems, a small perturbation boundary condition method is used to account for the moving surfaces. The method therefore does not require repeated grid re-generation for the deforming surfaces. The overall method is both accurate and robust. Computations of the well-known Onera M6 wing, RAE wing-body configuration, the AGARD 445.6 wing flutter test case show good accuracy and efficiencies. Simulations of the aeroelastic behavior of a complete fighter-type aircraft with wing tip missiles at high transonic speeds further demonstrate the practical usefulness of the present boundary conditions technique.     

面向分级设计优化的飞行器参数化建模方法

针对飞行器气动隐身外形综合设计优化问题,提出合适的面向分级设计优化流程,建立适应该流程的渐进分层参数化建模方法;用基于敏度分析的参数影响程度分析方法筛选复杂设计变量;采用多学科设计优化（Multidisplinary Design Optimization,MDO）理论和差分进化算法进行飞行器气动隐身外形的综合设计优化.将该方法用于某飞行器外形设计优化,结果表明：该方法合理可行,可为飞行器外形多学科设计优化提供一定参考.     

Blended wing body structures in multidisciplinary pre-design

The structural analysis of blended wing body (BWB) aircraft configurations is presented in the context of a preliminary, multidisciplinary aircraft design process by means of the aircraft design and optimization program (PrADO) of the Institut of Aircraft Design and Lightweight Structures of the TU Braunschweig. A multidisciplinary process is described that enables parametric creation of detailed finite element models and its loads. Iteratively different flight conditions are trimmed and corresponding pressure distributions calculated by the higher-order subsonic and supersonic panel code HISSS. Each defined loading condition is used for the iterative structural sizing of the primary structure. Based on finite element idealization, a mass estimation of all structural masses is performed. The primary and secondary masses are fed back into the closed overall aircraft optimization loop of PrADO until this iterative procedure shows convergence on calculated aircraft variables (e.g., aircraft masses and static engine thrust). This automated process enables further configuration improvements using manual parametric variations or optimization features of PrADO with an objective function being defined by fuel consumption, aircraft mass, or direct operating costs. Different structural solutions and their integration in the global model are presented for passenger versions of a 700 passenger BWB with special consideration of a pressurized cabin. As an example, structural masses and parametric studies on the influence of the center body rib spacing are presented and compared by weight breakdowns.     

Multidisciplinary aircraft design and evaluation software integrating CAD,analysis, database,and optimization

The development of an overall strategy to design the aircraft using analysis codes is presented. The procedure is automated through the integrated software MADE (multidisciplinary aircraft design and evaluation) developed in this research that enables the determination of an optimum set of aircraft configurations. The core ingredients of MADE are (i) analysis codes, which utilizes initial sizing, aerodynamics, mass, stability and control, propulsion, performance, and RCS (radar cross section) analyses, (ii) optimization utilizing gradient-based optimization technique, response surface modeling, and Carpet plot, (iii) database utilizing commercial Oracle 8i database management system (DBMS) and OCI (Oracle call interface) to save design parameters and aircraft configurations. The complexity of input and output containing massive amounts of data has previously placed severe limitations on the use of aircraft analysis codes by the general aerospace engineering community. By using C++ and the Microsoft Foundation Classes to develop a graphical user interface (GUI) and using DBMS, it is believed that use of MADE can avoid the additional cumbersome, and sometimes tedious, task of preparing the required input files manually and can make the transition to general usability in an aerospace engineering environment. For detail explanation, examples of E–R diagram, class diagram, and data flows between codes for the MADE are also presented.     

The implementation of a knowledge-based framework for the aerodynamic optimization of a morphing wing device

In the field of aerospace engineering currently a lot of research effort is directed towards the reduction of cruise drag of civil transport aircraft in order to reduce fuel burn, and hence environmental impact and costs. In order to reduce cruise drag, a promising method is under consideration by adjusting, or rather morphing the rear part of the aircraft’s wing during cruise flight. Given the premature state of knowledge of such a design implementation, a knowledge-based computational framework is developed. The purpose of this framework is to allow for an aerodynamic optimization of a section of the wing. The framework is set up in such a way that all relevant design knowledge generated in the process can be captured and used in a subsequent mechanical design process. In this fashion, the complex design process of a novel morphing wing device can be automated to a certain degree. This automation can be used to construct a large number of different feasible and optimized designs with varying boundary conditions of a complex experimental device.This article describes the initial 2-dimensional aerodynamic design step of the morphing device under consideration and how it is implemented in a knowledge-based optimization framework. It describes the initial stage of the development of this tool, as it will be expanded by a number of design steps that each adds more detail to the design in all relevant aspect fields (aerodynamic, structural, actuation, etc.). Ultimately, this tool will be used to obtain a thorough evaluation of a number of different proposed structural solutions and allow for a comparison between them.     

Optimization process for configuration of flexible joined-wing

An optimized configuration design utilizing both structural and aerodynamic analyses of a flexible joined-wing configuration is presented in this paper. The joined-wing aircraft concept fulfills a proposed long-endurance surveillance mission and incorporates a load-bearing antenna structure embedded in the wing skin. Aerodynamic, structural, and optimization analyses are completed a number of times. A range of joined-wing configurations were trimmed for critical flight conditions and then structurally optimized for trimmed flight and gust loads to achieve a minimum weight for each configuration. A response surface statistical analysis was then applied to determine optimal joined-wing aircraft configurations. The response surface showed trends in the design of lightweight joined-wing aircraft. The revised version of this paper was presented at the 10th MAO Multidisciplinary Analysis and Optimization Conference, August 30–September 1, 2004.     

A parallel aerostructural optimization framework for aircraft design studies

Preliminary aircraft design studies use structural weight models that are calibrated with data from existing aircraft. Computing weights with these models is a fast procedure that provides reliable weight estimates when the candidate designs lie within the domain of the data used for calibration. However, this limitation is too restrictive when we wish to assess the relative benefits of new structural technologies and new aircraft configurations early in the design process. To address this limitation, we present a computationally efficient aerostructural design framework for initial aircraft design studies that uses a full finite-element model of key structural components to better assess the potential benefits of new technologies. We use a three-dimensional panel method to predict the aerodynamic forces and couple the lifting surface deflections to compute the deformed aerodynamic flying shape. To be used early in the design cycle, the aerostructural computations must be fast, robust, and allow for significant design flexibility. To address these requirements, we develop a geometry parametrization technique that enables large geometric modifications, we implement a parallel Newton–Krylov approach that is robust and computationally efficient to solve the aeroelastic system, and we develop an adjoint-based derivative evaluation method to compute the derivatives of functions of interest for design optimization. To demonstrate the capabilities of the framework, we present a design optimization of a large transport aircraft wing that includes a detailed structural design parametrization. The results demonstrate that the proposed framework can be used to make detailed structural design decisions to meet overall aircraft mission requirements.     

Transition prediction for three-dimensional flows using parallel computation

A computational method to predict transition lines for general three-dimensional configurations is presented. The method consists of a coupled program system including a 3D Navier-Stokes solver, a transition module, a boundary layer code and a stability code. The newly developed transition module has been adapted to be used with parallel computation to account for the high computational demand for three-dimensional configurations. Detailed computations have been performed to show the ability of the Navier-Stokes code to provide three-dimensional boundary layer data of high accuracy needed for the stability analysis. A comprehensive investigation on general computational and parallel performance identifies the numerical effort for the transition prediction method. The procedure has been validated comparing the numerical results with experiments for the flow around an inclined prolate spheroid. Feasibility studies on generic transport aircraft have been performed to show the code’s capability to predict transition lines on general complex geometries.