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。     

Transonic drag prediction on a DLR-F6 transport configuration using unstructured grid solvers

A second international AIAA Drag Prediction Workshop (DPW-II) was organized and held in Orlando Florida on June 21-22, 2003. The primary purpose was to investigate the code-to-code uncertainty, address the sensitivity of the drag prediction to grid size and quantify the uncertainty in predicting nacelle/pylon drag increments at a transonic cruise condition. This paper presents an in-depth analysis of the DPW-II computational results from three state-of-the-art unstructured grid Navier-Stokes flow solvers exercised on similar families of tetrahedral grids. The flow solvers are USM3D - a tetrahedral cell-centered upwind solver, FUN3D - a tetrahedral node-centered upwind solver, and NSU3D - a general element node-centered central-differenced solver. Overall, grid refinement did not consistently improve the correlation with experimental data for either the wing/body or the wing/body/nacelle pylon configuration. Although, the range in total drag for the wing/body fine grids was only 5 counts, a code-to-code comparison of surface pressures and surface restricted streamlines indicated that the three solvers were not all converging to the same flow solutions- different shock locations and separation patterns were evident. Similarly, the wing/body/nacelle/pylon solutions did not appear to be converging to the same flow solutions. Although the absolute values of total drag predicted by two of the solvers for the medium and fine grids did not compare well with the experiment, the incremental drag predictions were within ±3 counts of the experimental data. Although, the sources of code-to-code variation in force and moment predictions for the three unstructured grid codes have not yet been identified, the current study reinforces the necessity of applying multiple codes to the same application to assess uncertainty.     

Efficient nonlinear static aeroelastic wing analysis

The objective of this work is to demonstrate a computationally efficient, high-fidelity, integrated static aeroelastic analysis procedure. The aerodynamic analysis consists of solving the nonlinear Euler equations by using an upwind cell-centered finite-volume scheme on unstructured tetrahedral meshes. The use of unstructured grids enhances the discretization of irregularly shaped domains and the interaction compatibility with the wing structure. The structural analysis utilizes finite elements to model the wing so that accurate structural deflections are obtained and allows the capability for computing detailed stress information for the configuration. Parameters are introduced to control the interaction of the computational fluid dynamics and structural analyses; these control parameters permit extremely efficient static aeroelastic computations. To demonstrate and evaluate this procedure, static aeroelastic analysis results for a flexible wing in low subsonic, high subsonic (subcritical), transonic (supercritical), and supersonic flow conditions are presented.     

Optimal airfoil shapes for viscous transonic flows of Bethe-Zel’dovich-Thompson fluids

High-performance airfoils for transonic viscous flows of dense gases are constructed using an efficient high-order accurate flow solver coupled with a multi-objective genetic algorithm. Dense gases are theoretically characterized by reversed behavior of the speed of sound in isentropic perturbations for a range of temperatures and pressures in the vapor phase. A class of dense gases, namely the so-called Bethe-Zel’dovich-Thompson (BZT) fluids, might exhibit nonclassical gasdynamic behaviors in the transonic and supersonic regimes, such as the disintegration of compression shocks. Utilizing BZT gases as working fluids may result in low drag exerted on airfoils operating at high transonic speeds thanks to an increase in the airfoil critical Mach number. This advantage can be further improved by a proper design of the airfoil shape, also leading to the enlargement of the airfoil operation range within which BZT effects are significant. Such a result is of particular interest in view of the exploitation of BZT fluids for the development of high-efficiency turbomachinery.     

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.     

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.     

Observations on three-dimensional nonlinear transonic wall-interference assessment and correction using semispan wing data

The current interest in semispan wing testing generally places a greater demand on understanding wall-interference effects than is required in conventional sting-mounted model testing. As part of NASA Langley Research Center's development and evaluation of transonic wind-tunnel wall-interference assessment and correction (WIAC) ideas and methods, a nonlinear, three-dimensional, transonic WIAC code has been applied to four sets of transonic semispan wing data. These data were acquired nearly 20 years ago for the specific purpose of evaluating three-dimensional transonic computational fluid dynamics methodologies and included the measured wall (or near wall) pressure data that are required by the WIAC procedure. Previous papers have focused on the evaluation of nonlinear WIAC and other correction-method results by monitoring the correlations of corrected data and by using a Navier–Stokes code as an independent “free air” check. In the present paper, synoptic observations are made in regard to the transonic WIAC results and the factors that affect these results.     

Parallel three-dimensional direct simulation Monte Carlo method and its applications

The development of a parallel three-dimensional direct simulation Monte Carlo (DSMC) method using unstructured cells is reported. Variable hard sphere molecular model and no time counter method are used for the molecular collision kinetics, while the cell-by-cell ray-tracing technique is implemented for particle movement. Developed serial code has been verified by comparing the results of a supersonic corner flow with those of Bird’s three-dimensional structured DSMC code. In addition, a benchmark test is performed for an orifice expanding flow to verify the parallel implementation of DSMC method by comparing with available experimental data. Static physical domain decomposition is used to distribute the workload among multiple processors by considering the estimated particle weighting distribution. Two-step multi-level graph partitioning technique is used to perform the required domain decomposition. Completed code is then applied to compute a hypersonic flow over a sphere (external flow) and the flow field of a spiral drag pump (internal flow), respectively. Results of the former are in good agreement with previous numerical results using axisymmetric DSMC method and experimental data. Results of the latter also agree well with previous numerical results.     

Aeroelastic tailoring using lamination parameters

The aim of the present work is to passively reduce the induced drag of the rear wing of a Formula One car at high velocity through aeroelastic tailoring. The angle-of-attack of the rear wing is fixed and is determined by the required downforce needed to get around a turn. As a result, at higher velocity, the amount of downforce and related induced drag increases. The maximum speed on a straight part is thus reduced due to the increase in induced drag. A fibre reinforced composite torsion box with extension-shear coupled upper and lower skins is used leading to bending-torsion coupling. Three-dimensional static aeroelastic analysis is performed loosely coupling the Finite Element code Nastran and the Computational Fluid Dynamics panel code VSAERO using ModelCenter. A wing representative of Formula One rear wings is optimised for minimum induced drag using a response surface methodology. Results indicate that a substantial induced drag reduction is achievable while maintaining the desired downforce during low speed turns.     

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.     

Implementation of density-based solver for all speeds in the framework of OpenFOAM

In the framework of open source CFD code OpenFOAM, a density-based solver for all speeds flow field is developed. In this solver the preconditioned all speeds AUSM+(P) scheme is adopted and the dual time scheme is implemented to complete the unsteady process. Parallel computation could be implemented to accelerate the solving process. Different interface reconstruction algorithms are implemented, and their accuracy with respect to convection is compared. Three benchmark tests of lid-driven cavity flow, flow crossing over a bump, and flow over a forward-facing step are presented to show the accuracy of the AUSM+(P) solver for low-speed incompressible flow, transonic flow, and supersonic/hypersonic flow. Firstly, for the lid driven cavity flow, the computational results obtained by different interface reconstruction algorithms are compared. It is indicated that the one dimensional reconstruction scheme adopted in this solver possesses high accuracy and the solver developed in this paper can effectively catch the features of low incompressible flow. Then via the test cases regarding the flow crossing over bump and over forward step, the ability to capture characteristics of the transonic and supersonic/hypersonic flows are confirmed. The forward-facing step proves to be the most challenging for the preconditioned solvers with and without the dual time scheme. Nonetheless, the solvers described in this paper reproduce the main features of this flow, including the evolution of the initial transient.     

可压缩全速湍流流动的边界数值处理方法研究

计算流体力学问题的边界条件处理方法关系到数值仿真结果的精确度。为解决算法的精度,提出了三维可压缩湍流流动的边界条件数值处理方法,对所研究的边界类型包括进口边界、出口边界和固体壁面,流动的速度范围涉及亚音速、跨音速和超音速。流场数值仿真采用SIMPLE算法,湍流采用k-ε模型仿真。将边界总结为沟通型和孤立型边界两种类型,对每一控制方程分别阐述特定的数值处理方法。应用提出的边界处理方法对单圆弧凸包通道进行数值仿真获得了合理的结果,跨音速和超音速情形下准确地计算出了流场中存在的激波。     

反折式二元超声速进气道设计及数值研究

研究固体冲压发动机进气道优化问题,为降低弹用超声速进气道的外部阻力和提高导弹飞行速度,提出了一种具有低外阻特性的反折式二元进气道方案,明确设计流程和主要设计参数的选取,并针对2~3.5Ma速度范围的应用需求开展了方案设计。进一步采用Fluent软件进行数值仿真,研究了反折式进气道的流场特性和性能水平,并与传统设计方案进行了对比。结果表明,在捕获流量相同的条件下,反折式进气道比原方案具有更小的外部阻力及外廓尺寸,还能保持与原方案相当的总压恢复性能,满足工程应用需求,为设计提供依据。     

Fast convergence of inviscid fluid dynamic design problems

A new procedure for robust and efficient design optimization of inviscid flow problems has been developed and implemented on a wide variety of test problems. The methodology involves the use of an accurate flow solver to calculate the objective function and an approximate, dissipative flow solver, which is used only in the solution of the discrete quasi-time-dependent adjoint problem. The resulting design sensitivities are very robust even in the presence of noise or other non-smoothness associated with objective functions in many high-speed flow problems. The design problem is solved using what we term progressive optimization, whereby a sequence of a partially converged flow solution, followed by a partially converged adjoint solution followed by an optimization step is performed. This procedure is performed using a sequence of progressively finer grids for the solution of the flow field, while only using coarser grids for the adjoint equation solution.This approach has been tested on numerous inverse and direct (constrained) design problems involving two- and three-dimensional transonic nozzles and airfoils as well as supersonic blunt bodies. The methodology is shown to be robust and highly efficient, with a converged design optimization produced in no more than the amount of computational work to perform from 0.5 to 2.5 fine-mesh flow analyses.     

Development of a MEMS-based control system for compressible flow separation

A MEMS-based sensor and actuator system has been designed and fabricated for separation control in the compressible flow regime. The MEMS sensors in the system are surface-micromachined shear stress sensors and the actuators are bulk-micromachined balloon vortex generators (VGs). A three-dimensional (3-D) wing model embedded with the shear stress sensors and balloon VGs was tested in a transonic wind tunnel to evaluate the performance of the control system in a range of Mach number between 0.2 and 0.6. At each Mach number tested, the shear stress sensors quantify the boundary layer on the surface of the wing model while the balloon VGs interact with the boundary layer in an attempt to provide flow control. The shear stress measurements indicate the presence of a separated flow on the trailing ramp section of the wing model at all Mach numbers tested when the balloon VGs are not activated. This result is confirmed by total pressure measurements downstream from the wing model where a wake profile is observed. When the balloon VGs are activated, the shear stress level on the trailing ramp increases with the Mach number. At the highest Mach number tested, this increase elevates the shear stress on the ramp to almost the same level as the unseparated flow, suggesting the possibility of a boundary layer reattachment. This result is supported by the downstream pressure measurements which show a large pressure recovery when the balloon VGs are activated. The wind tunnel experiment successfully demonstrated two aspects of the MEMS flow control system: the effectiveness of the microshear stress sensors in measuring the separation characteristics of a high-speed compressible flow and the ability of the microballoons in positively enhancing the aerodynamic performance of a high-speed wing through boundary layer modification.     

Robust control design of an air‐breathing engine for a supersonic vehicle using backstepping and UKF

This paper presents an efficient robust control design approach for an air‐breathing engine for a supersonic vehicle using the Lyapunov stability theory based nonlinear backstepping control, augmented with unscented Kalman filter (UKF). The primary objective of the control design is to ensure that the thrust produced by the engine tracks the commanded thrust by regulating the fuel flow to the combustion chamber. Moreover, as the engine operates in a supersonic range, an important secondary objective is to manage the shock wave location in the intake for maximum pressure recovery with adequate safety margin by varying the throat area of the nozzle simultaneously. To estimate the states and parameters as well as to filter out the process and sensor noises, a UKF has been incorporated for robust output feedback control computation. Furthermore, independent control designs for the actuators have been carried out to assure satisfactory performance of the engine. Additionally, a guidance loop is designed to generate a typical flight trajectory of the representative vehicle using a nonlinear suboptimal input constrained model predictive static programming formulation for testing the performance of the engine. Simulation results clearly indicate quite successful robust performance of the engine during both climb and cruise phases.     

On the numerical analysis of compressible flow problems by the “modified FLIC method”

This paper presents a new computational technique for transonic flow problem analysis. This method, named Modified FLIC Method, is based on a time-marching technique of FLIC (fluid in cell) method and employs triangular elements conventionally used in finite element method. This technique can be applied to transonic flows with any complicated boundary shapes. Three problems were solved in this paper, the first was a supersonic flow around a circular cylinder, the second was a transonic flow between tubrine blade cascades and the last was an unsteady flow in a duct with a junction. The calculated results showed a good agreement with the experimental data.     

遗传算法在跨超声速风洞总压控制中的应用

总压作为风洞控制中的重要流场参数,其调节性能是风洞控制系统能否满足试验要求的重要指标,为提高跨超声速风洞的总压控制水平,需对总压控制策略进行设计。针对某跨超声速风洞对总压控制系统提出的快速性和精确性要求,提出串级控制、智能PID控制和总压分段控制等方法,并利用MATLAB系统辨识工具箱对流场调节阶段的总压系统模型进行了辨识。提出将遗传算法应用于风洞流场调节阶段的PID控制器参数整定中,重点对基于遗传算法的PID控制原理和参数整定步骤进行介绍,并针对遗传算法的遗传算子进行了设计。系统仿真和风洞实际运行情况表明：该方法较常规PID参数整定与优化方法,具有更好的控制性能指标,满足总压控制系统精确性、快速性、鲁棒性等要求,为后续风洞建设和设备改造提供了新方法。     

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.     

迷你翼梢小翼增升减阻效应

选取某窄体客机的翼梢小翼为研究对象,采用Spalart Allmaras模型对无翼梢小翼、全尺寸翼梢小翼和迷你翼梢小翼3种机翼构型进行数值模拟,通过流场分析和速度分解等手段,研究翼梢小翼的增升减阻机理。结果表明：迷你翼梢小翼有恢复涡核流速、减弱涡流掺混程度和梳理翼梢气流的作用;增升减阻的关键在于迷你翼梢小翼对气流方向的修正;翼梢小翼的局部流动差异会对整体机翼近场造成影响。由于尺寸较小,迷你翼梢小翼能在较大攻角范围内改善传统翼梢小翼的性能,具有一定的实践意义。