Weilin Yi

Multi-Stage Turbomachinery Blades Optimization Design Using Adjoint Method and Thin Shear-Layer N-S Equations

Traditionally, 3D aerodynamic shape design with the aid of optimization algorithm in an analysis mode has provided a rational and direct search through design space, but it is usually too time-consuming. Further improvement to reduce design cycle is probably a necessary concern in turbomachinery community. Due to less computational cost, adjoint method has received considerable attention in recent years. This paper focuses on continuous adjoint method, and couples with thin shear-layer N-S equations to formulate an efficient sensitivity analysis model for multi-stage turbomachinery blades in the specified objective function. This model includes adjoint equations/boundary conditions, and the sensitivity of objective function to design variable vector. Integrating a 3D blade perturbation parameterization and the simple steepest decent method, a frame of a gradient-based aerodynamic shape design system is constructed. Numerical implementation to solve flow equations and adjoint equations is very similar, and once they are converged respectively, the sensitivity can be calculated by complex method and mesh perturbation efficiently. Thus, a fast Automatic-CFD-Design tool is developed, including three sub-solvers to solve flow equations, adjoint equations and calculate sensitivity respectively. Flow surface design of a 1-1/2 compressor stage in the specified target pressure distribution is used to validate the present approach. Flow field design of NASA transonic compressor stage 35 aiming to increase efficiency and remain mass flow rate and pressure ratio unchanged is taken.© 2012 ASME

Numerical Studies on Improving Performance of Rotor-67 by Blended Blade and EndWall Technique

The paper describes an advanced blading concept for highly loaded turbomachinery, named as Blended Blade and EndWall (BBEW). It implies the following three separate aspects or any of their combinations, i.e. increasing the dihedral between blade surface and endwall as much as possible, increasing the minimum curvature radius of curved transition surface that connects the blade with the endwall, and decreasing the streamwise gradient of the dihedral as the dihedral reduces streamwise. To the knowledge of the authors this is the first approach reported in the open literature that try to control corner separations and horse-shoe vortex (HSV) systematically under the guidance of the Rules of Dihedral (RD). In this paper, BBEW is applied to NASA rotor 67 which is featured by hub corner separation at the suction side and intensive HSV originating from blade leading edge at the hub. Two successive BBEW modifications are made to promote the performance of rotor 67. First, only a curved transition surface is designed to cover the hub corner at the suction side. Emphasis is paid on the streamwise distribution of the minimum curvature radius. Then, the second curved transition surface is added at the leading edge in near hub part. The nearby streamwise gradient of the dihedral is changed from infinity to a finite. Numerical tests showed that the implementation of BBEW had not only nearly eliminated the hub corner separation, but also weakened the leading edge HSV. Performance improvements are observed throughout its operation range. The work presented makes a contribution to a fully 3-D blading methodology for higher aerodynamic loading, engine efficiency and specific thrust.Copyright

Study on aerodynamic optimal super/transonic turbine cascade and its geometry characteristics

The shock waves are important phenomena in transonic turbines, which cause lots of negative effects on the aerodynamic performance. Much of attention had been paid on reducing the strength of the shock waves via modifying turbine cascade geometry, and it is highly preferred to build experiences on the relationship between the cascade aerodynamic performance and the geometric parameters. The paper presents a numerical study on the aerodynamic optimal transonic turbine cascade and its geometry characteristics. Three typical Russia transonic turbine cascades with different design conditions are selected and optimized using adjoint method at three different back pressures, respectively. Thus, the best geometry parameters for optimum aerodynamic performance can be found. Then the key geometry parameters of optimized cascades are extracted and compared with the original ones. Results show that even the best designs by hands could be less efficient than ones by computer-aided optimizations. Some experiences on how to set the key geometry parameters for a best performance are obtained. The reduced shock profiling is applied to the thermal turbomachinery and machine dynamics transonic turbine by using the adjoint method. The performance of the thermal turbomachinery and machine dynamics transonic turbine was increased significantly.

Adjoint Optimization of Multistage Axial Compressor Blades with Static Pressure Constraint at Blade Row Interface

Abstract Adjoint method is an important tool for design refinement of multistage compressors. However, the radial static pressure distribution deviates during the optimization procedure and deteriorates the overall performance, producing final designs that are not well suited for realistic engineering applications. In previous development work on multistage turbomachinery blade optimization using adjoint method and thin shear-layer N-S equations, the entropy production is selected as the objective function with given mass flow rate and total pressure ratio as imposed constraints. The radial static pressure distribution at the interfaces between rows is introduced as a new constraint in the present paper. The approach is applied to the redesign of a five-stage axial compressor, and the results obtained with and without the constraint on the radial static pressure distribution at the interfaces between rows are discussed in detail. The results show that the redesign without the radial static pressure distribution constraint (RSPDC) gives an optimal solution that shows deviations on radial static pressure distribution, especially at rotor exit tip region. On the other hand, the redesign with the RSPDC successfully keeps the radial static pressure distribution at the interfaces between rows and make sure that the optimization results are applicable in a practical engineering design.

Shape Optimization of Axial Compressor Blades Using Adjoint Method With Emphasis on Thickness Distribution

Shape parameterization plays an important role in aerodynamic optimization design of axial compressor blades. Blade thickness is one of the most important parameters in blade design, which has strong influence on compressor aerodynamic performance. However, the previous adjoint-based optimization designs using the Hicks-Henne functions only parameterized the perturbations to the tangential coordinates of points on suction surface or meanline, and kept the tangential thickness of the blade constant during the optimization process. In previous development work of turbomachinery blade optimization using adjoint method and thin shear-layer N-S equations, a new shape parameterization is introduced, which uses Hicks-Henne functions to parameterize the perturbations to both the tangential coordinates of mesh points on suction blade surface and the tangential thickness of the blade. This new approach is applied to the redesign of NASA rotor 67 and the results obtained with and without the blade tangential thickness parameterization are discussed in detail. The results show the redesign with and without the blade tangential thickness parameterization can both improve the aerodynamic performance of the axial compressor. However, the redesign with the blade tangential thickness parameterization can produce a consistently better performance than that without it.Copyright

Shape Optimization of Multi-Stage Axial Compressor Blades Using Adjoint Method With Static Pressure Constraint at Blade Row Interface

Adjoint method is an important tool for design refinement of multistage compressors. However, the radial static pressure distribution deviates during the optimization procedure and deteriorates the overall performance, producing final designs that are not well suited for realistic engineering applications. In previous development work on multistage turbomachinery blade optimization using adjoint method and thin shear-layer N-S equations, the entropy production is selected as the objective function with given mass flow rate and total pressure ratio as imposed constraints. The radial static pressure distribution at the interfaces between rows is introduced as a new constraint in the present paper. The approach is applied to the redesign of a five-stage axial compressor, and the results obtained with and without the constraint on the radial static pressure distribution at the interfaces between rows are discussed in detail. The results show that the redesign without radial static pressure distribution constraint (RSPDC) gives an optimal solution that shows deviations on radial static pressure distribution, especially at rotor exit tip region. On the other hand, the redesign with the RSPDC successfully keeps the radial static pressure distribution at the interfaces between rows and make sure that the optimization results are applicable in a practical engineering design.© 2015 ASME

Coupled Fluid-Thermal-Solid Simulation of Axial Through-Flow Rotating Cavities

The rotating cavities of aero-engine compressors are the main part of secondary air flow system. It is known that there are typical multidisciplinary fluid-thermal-solid coupling characteristics in them. The high precision prediction of disc surface temperature is very important for structure designer to select materials, control blade clearances et al. The aim of this paper is to investigate the aerodynamic-thermal simulation model to obtain the method and tool for reliable temperature prediction.The paper firstly selected publicly available experimental data of two rotating cavity geometries with twin-discs to validate the precision of established fluid-thermal simulation model with the different grids, difference schemes and turbulence models. The results showed that the RNG-KE turbulence model with QUICK scheme has the better simulation precision for flow structure and Nusselt number distribution.Based on the above research, a fluid-thermal-solid coupling simulation of a twin-cavities model which is approaching to the real conditions of aero-engine has been carried out. The wall temperature distribution on inner surface has been obtained and its maximum error comparing with the experimental value is 8°C. Also the results further validated the reliabilities of the flow model, heat transfer model and fluid-thermal-solid coupling model. The paper also shows the flow field structure of rotating cavity for further understanding the internal flow characteristics.Copyright

A Model for Describing the Influences of SUC-EW Dihedral Angle on Corner Separation

This paper presents a model for describing the influences of SUC-EW dihedral angle on corner separation in turbomachinery, in which SUC-EW dihedral angle refers to the dihedral angle at the intersection line between blade ‘SUCtion’ and End-Wall surfaces. Based on the physical intuition of that the three-dimensional (3D) corner boundary layer is the conflux of both blade and end wall boundary layers, an equivalent two-dimensional(2D) corner boundary layer is put forward to predict the behavior of corner boundary layer. In this procedure, the cross flow effect in corner boundary layer and the three-dimensionality of the nearby main flow are ignored. The influence of the SUC-EW dihedral angle is included by another assumption. That is, the aero blockage and momentum loss of both blade and end wall boundary layers are conserved during the procedure of superimposing the two (both blade and end wall) 2D boundary layers to form the equivalent corner one. Then the corner separation is judged by combining the behaviors of the three boundary layers, i.e. the blade, the end wall and the equivalent 2D corner boundary layers. The present model reveals the influence of the SUC-EW dihedral angle and its streamwise gradient on the corner separation. Carefully monitoring and controlling this dihedral angle and its streamwise gradient are important ways to alleviate or even eliminate the corner separation. Simple numerical investigations show that the model is qualitatively correct.Copyright

Maximizing 3D Clocking Effect by the Guidance of Edge Matching

Clocking can produce benefit has become a common viewpoint in turbomachinery field. However, there still is a question, that is “why the benefit of two-dimensional (2D) clocking model is always larger than that of three-dimensional (3D) clocking model?”. A general way of maximizing 3D clocking benefit by edge matching (EM) is put forward. On the basis of experience and knowledge accumulated in investigations on clocking effects, a further simplified method is presented and emphasized. Aachen 1.5 stage low speed subsonic axial turbine is used as an example of maximizing 3D clocking benefits by applying EM. Results show that 3D clocking benefit can be maximized by routinely using EM with relatively low computational resource requirements, and the benefit is considerable enough to be paid more attention.Copyright