Aerodynamic design on high-speed trains

Compared with the traditional train,the operational speed of the high-speed train has largely improved,and thedynamicenvironmentofthetrainhaschangedfromoneof mechanical domination to one of aerodynamic domination.The aerodynamic problem has become the key technological challenge of high-speed trains and significantl affects the economy,environment,safety,and comfort.In this paper,the relationships among the aerodynamic design principle,aerodynamic performance indexes,and design variables are firs studied,and the research methods of train aerodynamics are proposed,including numerical simulation,a reducedscale test,and a full-scale test.Technological schemes of train aerodynamics involve the optimization design of the streamlined head and the smooth design of the body surface.Optimization design of the streamlined head includes conception design,project design,numerical simulation,and a reduced-scale test.Smooth design of the body surface is mainly used for the key parts,such as electric-current collecting system,wheel truck compartment,and windshield.The aerodynamic design method established in this paper has been successfully applied to various high-speed trains（CRH380A,CRH380 AM,CRH6,CRH2 G,and the Standard electric multiple unit（EMU）） that have met expected design objectives.The research results can provide an effective guideline for the aerodynamic design of high-speed trains.

Aerodynamic design of transonic fan/compressor by 3D viscous RNS combined with genetic algorithms

This paper presents an aerodynamic design of a small transonic fan by 3D viscous RNS solver combined with genetic algorithms.The aerodynamic design system based on the 3D viscous RNS solver reduces the dependency on the design experience for designers.Furthermore the optimum with genetic algorithms is an effective method for improving the transonic fan performance as a part of the design system.The design result showed that the transonic fan designed by this method reaches the design requirement even with more efficiency value.

Aerodynamic/stealth design of S-duct inlet based on discrete adjoint method

It is a major challenge for the airframe-inlet design of modern combat aircrafts,as the flow and electromagnetic wave propagation in the inlet of stealth aircraft are very complex.In this study,an aerodynamic/stealth optimization design method for an S-duct inlet is proposed.The upwind scheme is introduced to the aerodynamic adjoint equation to resolve the shock wave and flow separation.The multilevel fast multipole algorithm(MLFMA)is utilized for the stealth adjoint equation.A dorsal S-duct inlet of flying wing layout is optimized to improve the aerodynamic and stealth characteristics.Both the aerodynamic and stealth characteristics of the inlet are effectively improved.Finally,the optimization results are analyzed,and it shows that the main contradiction between aerodynamic characteristics and stealth characteristics is the centerline and crosssectional area.The S-duct is smoothed,and the cross-sectional area is increased to improve the aerodynamic characteristics,while it is completely opposite for the stealth design.The radar cross section(RCS)is reduced by phase cancelation for low frequency conditions.The method is suitable for the aerodynamic/stealth design of the aircraft airframe-inlet system.

Application of a PCA-DBN-based surrogate model to robust aerodynamic design optimization

An efficient method employing a Principal Component Analysis(PCA)-Deep Belief Network(DBN)-based surrogate model is developed for robust aerodynamic design optimization in this study.In order to reduce the number of design variables for aerodynamic optimizations,the PCA technique is implemented to the geometric parameters obtained by parameterization method.For the purpose of predicting aerodynamic parameters,the DBN model is established with the reduced design variables as input and the aerodynamic parameters as output,and it is trained using the k-step contrastive divergence algorithm.The established PCA-DBN-based surrogate model is validated through predicting lift-to-drag ratios of a set of airfoils,and the results indicate that the PCA-DBN-based surrogate model is reliable and obtains more accurate predictions than three other surrogate models.Then the efficient optimization method is established by embedding the PCA-DBN-based surrogate model into an improved Particle Swarm Optimization(PSO)framework,and applied to the robust aerodynamic design optimizations of Natural Laminar Flow(NLF)airfoil and transonic wing.The optimization results indicate that the PCA-DBN-based surrogate model works very well as a prediction model in the robust optimization processes of both NLF airfoil and transonic wing.By employing the PCA-DBN-based surrogate model,the developed efficient method improves the optimization efficiency obviously.

Global aerodynamic design optimization based on data dimensionality reduction

In aerodynamic optimization, global optimization methods such as genetic algorithms are preferred in many cases because of their advantage on reaching global optimum. However,for complex problems in which large number of design variables are needed, the computational cost becomes prohibitive, and thus original global optimization strategies are required. To address this need, data dimensionality reduction method is combined with global optimization methods, thus forming a new global optimization system, aiming to improve the efficiency of conventional global optimization. The new optimization system involves applying Proper Orthogonal Decomposition（POD） in dimensionality reduction of design space while maintaining the generality of original design space. Besides, an acceleration approach for samples calculation in surrogate modeling is applied to reduce the computational time while providing sufficient accuracy. The optimizations of a transonic airfoil RAE2822 and the transonic wing ONERA M6 are performed to demonstrate the effectiveness of the proposed new optimization system. In both cases, we manage to reduce the number of design variables from 20 to 10 and from 42 to 20 respectively. The new design optimization system converges faster and it takes 1/3 of the total time of traditional optimization to converge to a better design, thus significantly reducing the overall optimization time and improving the efficiency of conventional global design optimization method.

Aerodynamic Design Methodology for Blended Wing Body Transport

This paper puts forward a design idea for blended wing body（BWB）.The idea is described as that cruise point,maximum lift to drag point and pitch trim point are in the same flight attitude.According to this design idea,design objectives and constraints are defined.By applying low and high fidelity aerodynamic analysis tools,BWB aerodynamic design methodology is established by the combination of optimization design and inverse design methods.High lift to drag ratio,pitch trim and acceptable buffet margin can be achieved by this design methodology.For 300-passenger BWB configuration based on static stability design,as compared with initial configuration,the maximum lift to drag ratio and pitch trim are achieved at cruise condition,zero lift pitching moment is positive,and buffet characteristics is well.Fuel burn of 300-passenger BWB configuration is also significantly reduced as compared with conventional civil transports.Because aerodynamic design is carried out under the constraints of BWB design requirements,the design configuration fulfills the demands for interior layout and provides a solid foundation for continuous work.

Airfoil design optimization based on lattice Boltzmann method and adjoint approach

We present a new aerodynamic design method based on the lattice Boltzmann method （LBM） and the adjoint approach. The flow field and the adjoint equation are numerically simulated by the GILBM （generalized form of interpolation supplemented LBM） on non-uniform meshes. The first-order approximation for the equilibrium dis- tribution function on the boundary is proposed to diminish the singularity of boundary conditions. Further, a new treatment of the solid boundary in the LBM is described par- ticularly for the airfoil optimization design problem. For a given objective function, the adjoint equation and its boundary conditions are derived analytically. The feasibility and accuracy of the new approach have been perfectly validated by the design optimization of NACA0012 airfoil.

Hypersonic Waverider Surface Development Using Aerodynamic Flow Around Conical Bodies

Developing the waverider based hypersonic vehicles is an inverse design process in which shape is developed from a known flow field by tracing of streamlines to form a stream surface. The flow field can be based on a solution of Taylor Maccoll equation for a specified shock or cone angle. This Paper discusses the development of waverider shapes for hypersonic reentry vehicles.

Nonlinear uncertainty impact of geometric variations on aerodynamic performance of low-pressure turbine blades with ultra-high loading under extreme operational conditions

Uncertainty impact of random geometric variations on the aerodynamic performance of low-pressure turbine blades is considerable,which is further amplified by the current ultra-high-lift design trend for weight reduction.Therefore,this uncertainty impact on ultra-highly loaded blades under extreme operational conditions near the margins with potential large-scale open separation is focused on in this study.It is demonstrated that this impact is significant,unfavourable,and nonlinear,which is clearly severer under extreme conditions.In addition to the overall attenuation and notable scattering of specific performance,the operational margins with open separation are also notably scattered with great risk of significant reduction.This scattering and nonlinearity are dominated by the variations in leading-edge thickness.The thinning of leading edge triggers local transition,enhancing downstream friction and reducing resistance to open separation,which is further exacerbated by operational deterioration.However,the opposite thickening yields less benefit,implying nonlinearity.This unfavourable impact highlights the need for robust aerodynamic design,where both a safer operational condition and a more robust blade are indispensable,i.e.,a compromise among performance,weight,and robustness.Besides the necessary limitation of loading levels,a mid-loaded design is recommended to reduce adverse pressure gradients in both the leading edge and rear region of the suction side,which helps to decrease the susceptibility of the transition and open separation to random perturbations.Similar improvements can also be achieved by appropriately thickening the leading edge.

Aerodynamic optimization design for high pressure turbines based on the adjoint approach

Abstract A first study on the continuous adjoint formulation for aerodynamic optimization design of high pressure turbines based on S2 surface governed by the Euler equations with source terms is presented. The objective function is defined as an integral function along the boundaries, and the adjoint equations and the boundary conditions are derived by introducing the adjoint variable vec- tors. The gradient expression of the objective function then includes only the terms related to phys- ical shape variations. The numerical solution of the adjoint equation is conducted by a finite- difference method with the Jameson spatial scheme employing the first and the third order dissipa- tive fluxes. A gradient-based aerodynamic optimization system is established by integrating the blade stagger angles, the stacking lines and the passage perturbation parameterization with the quasi-Newton method of Broyden Fletcher Goldfarb-Shanno （BFGS）. The application of the continuous adjoint method is validated through a single stage high pressure turbine optimization case. The adiabatic efficiency increases from 0.8875 to 0.8931, whilst the mass flow rate and the pressure ratio remain almost unchanged. The optimization design is shown to reduce the passage vortex loss as well as the mixing loss due to the cooling air injection.

跨音速三维多升力面气动设计(英文)

An aerodynamic design method and corresponding codes are developed for three-dimensional multi lifting surfaces at transonic flow. It is based on the ＂iterative residual correction＂ concept that is successfully used for transonic wing design and subsonic multi-lifting surface design. The up-wind scheme is introduced into governing equations of multi-lifting surface design method and automatically acted when supersonic flow appears on the surface. A series of interface codes are programmed, including a target-pressure modification tool. Using the improved inverse aerodynamic design code, TAU code and interface codes, the transonic multi-lifting aerodynamic design software system is founded. Two cases of canard-wing configuration have been performed to validate the method and codes. The results show that the convergence of analysis/design iteration is very good at higher speed transonic flow.

Performance-Driven Multi-Objective Optimization Method for DLR Transonic Tandem Cascade Shape Design

Transonic tandem cascades can effectively increase the working load,and this feature conforms with the requirement of the large loads and pressure ratios of modern axial compressors.This paper presents an optimization strategy for a German Aerospace Center(DLR)transonic tandem cascade,with one front blade and two rear blades,at the inlet Mach number of 1.051.The tandem cascade profile was parameterized using 19 control parameters.Non-dominated sorting Genetic algorithm(NSGA-II)was used to drive the optimization evolution,with the computational fluid dynamics(CFD)-based cascade performances correction added for each generation.Inside the automatic optimization system,a pressure boundary condition iterative algorithm was developed for simulating the cascade performance with a constant supersonic inlet Mach number.The optimization results of the cascade showed that the deflection of the subsonic blade changed evidently.The shock wave intensity of the first blade row was weakened because of the reduced curvatures of the optimized pressure and suction sides of the front blade part and the downstream moved maximum thickness position.The total pressure losses decreased by 15.6%,20.9%and 19.9%with a corresponding increase in cascade static pressure ratio by 1.3%,1.8%and 1.7%,for the three cascade shapes in the Pareto solution sets under the near choke,the design and near stall conditions,respectively.

Aerodynamic configuration integration design of hypersonic cruise aircraft with inward-turning inlets

In this work, a novel airframe/propulsion integration design method of the wing-body configuration for hypersonic cruise aircraft is proposed, where the configuration is integrated with inward-turning inlets. With the help of this method, the major design concern of balancing the aerodynamic performance against the requirements for efficient propulsion can be well addressed. A novel geometric parametrically modelling method based on a combination of patched class and shape transition（CST） and COONs surface is proposed to represent the configuration, especially a complex configuration with an irregular inlet lip shape. The modelling method enlarges the design space of components on the premise of guaranteeing the configuration integrity via special constraints imposed on the interface across adjacent surfaces. A basic flow inside a cone shaped by a dual-inflection-point generatrix is optimized to generate the inward-turning inlet with improvements of both compression efficiency and flow uniformity. The performance improvement mechanism of this basic flow is the compression velocity variation induced by the variation of the generatrix slope along the flow path. At the design point, numerical simulation results show that the lift-to-drag ratio of the configuration is as high as 5.2 and the inlet works well with a high level of compression efficiency and flow uniformity. The design result also has a good performance on off-design conditions. The achievement of all the design targets turns out that the integration design method proposed in this paper is efficient and practical.

Aerodynamic optimization and mechanism design of flexible variable camber trailing-edge flap

Trailing-edge flap is traditionally used to improve the takeoff and landing aerodynamic performance of aircraft.In order to improve flight efficiency during takeoff,cruise and landing states,the flexible variable camber trailing-edge flap is introduced,capable of changing its shape smoothly from 50% flap chord to the rear of the flap.Using a numerical simulation method for the case of the GA（W）-2 airfoil,the multi-objective optimization of the overlap,gap,deflection angle,and bending angle of the flap under takeoff and landing configurations is studied.The optimization results show that under takeoff configuration,the variable camber trailing-edge flap can increase lift coefficient by about 8% and lift-to-drag ratio by about 7% compared with the traditional flap at a takeoff angle of 8°.Under landing configuration,the flap can improve the lift coefficient at a stall angle of attack about 1.3%.Under cruise state,the flap helps to improve the lift-todrag ratio over a wide range of lift coefficients,and the maximum increment is about 30%.Finally,a corrugated structure–eccentric beam combination bending mechanism is introduced in this paper to bend the flap by rotating the eccentric beam.

气动优化设计和计算流体力学在无人机设计中的应用

In the past two decades,the world’s unmanned aerial vehicle(UAV)industry has developed rapidly.Various kinds of UAVs have been used in military and civilian fields.Based on the characteristics of UAVs and the development of aerodynamics,this article analyzes the development of aerodynamic optimization design and dynamic numerical simulation technology,then lists engineering applications.Both aerodynamic optimization design and dynamic numerical simulation have greatly shortened the UAV design period and reduced the research and design cost.These two methods gradually replace traditional methods such as wind tunnel test.

A new non-linear vortex lattice method:Applications to wing aerodynamic optimizations

This paper presents a new non-linear formulation of the classical Vortex Lattice Method （VLM） approach for calculating the aerodynamic properties of lifting surfaces. The method accounts for the effects of viscosity, and due to its low computational cost, it represents a very good tool to perform rapid and accurate wing design and optimization procedures. The mathematical model is constructed by using two-dimensional viscous analyses of the wing span-wise sections, according to strip theory, and then coupling the strip viscous forces with the forces generated by the vortex rings distributed on the wing camber surface, calculated with a fully three-dimensional vortex lifting law. The numerical results obtained with the proposed method are validated with experimental data and show good agreement in predicting both the lift and pitching moment, as well as in predicting the wing drag. The method is applied to modifying the wing of an Unmanned Aerial System to increase its aerodynamic efficiency and to calculate the drag reductions obtained by an upper surface morphing technique for an adaptable regional aircraft wing.

Design and analysis of annular combustion chamber of a low bypass turbofan engine in a jet trainer aircraft

The design of an annular combustion chamber in a gas turbine engine is thebackbone of this paper.It is specifically designed for a low bypass turbofan engine in a jettrainer aircraft.The combustion chamber is positioned in between the compressor and turbine.lt has to be designed based on the constant pressure,enthalpy addition process.The presentmethodology deals with the computation of the initial design parameters from benchmarking ofreal-time industry standards and arriving at optimized values.It is then studied for feasibilityand finalized.Then the various dimensions of the combustor are calculated based on differentempirical formulas.The air mass flow is then distributed across the zones of the combustor.The cooling requirement is met using the cooling holes.Finally the variations of parameters atdifferent points are calculated.The whole combustion chamber is modeled using Siemens NX8.0,a modeling software and presented.The model is then analyzed using various parametersat various stages and levels to determine the optimized design.The aerodynamic flowcharacteristics is simulated numerically by means of ANSYS 14.5 software suite.The air-fuelmixture,combustion-turbulence,thermal and cooling analysis is carried out.The analysis isperformed at various scenarios and compared.The results are then presented in image outputsand graphs.

Aerodynamic optimization of a high-lift system with adaptive dropped hinge flap

The Adaptive Dropped Hinge Flap(ADHF) is a novel trailing edge high-lift device characterized by the integration of downward deflection spoiler and simple hinge flap, with excellent aerodynamic and mechanism performance. In this paper, aerodynamic optimization design of an ADHF high-lift system is conducted considering the mechanism performance. Shape and settings of both takeoff and landing configurations are optimized and analyzed, with considering the kinematic constraints of ADHF mechanism, and the desired optimization results were obtained after optimization. Sensitivity analysis proves the robustness of the optimal design. Comparison shows that the ADHF design has better comprehensive performance of both mechanism and aerodynamics than the conventional Fowler flap and simple hinge flap design.

Numerical optimization of transonic natural laminar flow nacelles

Natural laminar flow nacelle is a promising technology for drag reduction.In this paper,an optimization platform is established for the design of transonic axisymmetric and threedimensional natural laminar flow nacelles for large civil aircraft.The platform adopts the class/shape transformation method for geometric parameterization,a four-equation transition model for transition prediction,and the differential evolution algorithm combined with the radial basis function surrogate model as the optimization algorithm.The optimized axisymmetric nacelle demonstrates approximately 31%chord length of laminar flow,with the drag reduction of 13.3%.The influence of the Reynolds number and inlet mass flow rate on the optimization result is also investigated.The axis-symmetric nacelle optimization method is further used for the section profile design of a non-axisymmetric nacelle.An equivalent method is used to simulate the different local flow angles at different sections in the circumferential direction of the non-axisymmetric nacelle by using different inlet mass flow rates of the axisymmetric nacelle.The optimized natural laminar flow nacelle maintains over 30%chord length of laminar flow with robustness to the change of the freestream angle of attack.The total drag of the non-axisymmetric nacelle is reduced by 5.4%under cruise conditions.

Recent development of a CFD-wind tunnel correlation study based on CAE-AVM investigation

This paper starts with brief introduction to the open topic of the CFD and wing tunnel correlation study, followed by a description of the Chinese Aeronautical Establishment（CAE） –Aerodynamic Validation Model（AVM） and its wind tunnel test in the German-Dutch Wind tunnels（DNW）. The features of the aerodynamic design, the CFD approach, the wind tunnel model fabrication and the experimental techniques are discussed along with the motivation of the CAEDNW workshop on CFD-wind tunnel correlation study. The workshop objective is focused on the interference from the aero-elastic deformation of the wind tunnel model and the model support system to the aerodynamic performance and CFD validations. The four study cases, geometry and mesh preparation of the workshop are introduced. A comprehensive summary of the CFD results from the organizer and the participants is provided. Major observations, both CFD to CFD and CFD to wind tunnel, are identified and summarized. The CFD results of the participants are in good agreement with each other, and with the wind tunnel test data when the wing deformation and a Z-sting system are included in the CFD, indicating the importance of considering such interference at high subsonic Mach number of 0.85.