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

2015 ◽  
Author(s):  
Jia Yu ◽  
Lucheng Ji ◽  
Weiwei Li ◽  
Weilin Yi
Keyword(s):  
Optimization Design ◽  
Adjoint Method ◽  
Thickness Distribution ◽  
Axial Compressor ◽  
Aerodynamic Performance ◽  
Aerodynamic Optimization ◽  
Suction Surface ◽  
Compressor Blades ◽  
Blade Thickness ◽  
Shape Parameterization

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.

Continuous adjoint aerodynamic optimization design for multi-stage gas turbine with cooling air

2017 ◽  
Vol 89 (3) ◽  
pp. 444-456
Author(s):  
Lei Chen ◽  
Jiang Chen
Keyword(s):  
Gas Turbine ◽  
Optimization Design ◽  
Adjoint Method ◽  
Air Injection ◽  
Aerodynamic Optimization ◽  
Content Type ◽  
Multi Stage ◽  
Surface Code ◽  
Cooling Air ◽  
Continuous Adjoint

Purpose This paper aims to conduct the optimization of the multi-stage gas turbine with the effect of the cooling air injection based on the adjoint method. Design/methodology/approach Continuous adjoint method is combined with the S2 surface code. Findings The optimization of the stagger angles, stacking lines and the passage can improve the attack angles and restrain the development of the boundary, reducing the secondary flow loss caused by the cooling air injection. Practical implications The aerodynamic performance of the gas turbine can be improved via the optimization of blade and passage based on the adjoint method. Originality/value The results of the first study on the adjoint method applied to the S2 surface through flow calculation including the cooling air effect are presented.

Aerodynamic Optimization Design of Multi-stage Turbine Using the Continuous Adjoint Method

2015 ◽  
Vol 32 (2) ◽  
Author(s):  
Lei Chen ◽  
Jiang Chen
Keyword(s):  
Optimization Design ◽  
Adjoint Method ◽  
Shape Design ◽  
Aerodynamic Optimization ◽  
Aerodynamic Shape ◽  
Adjoint Formulation ◽  
Multi Stage ◽  
Continuous Adjoint Method ◽  
Continuous Adjoint

AbstractThis paper develops a continuous adjoint formulation for the aerodynamic shape design of a turbine in a multi-stage environment based on S

Inverse Design of 3D Compressor Blades With Adjoint Method

2003 ◽  
Author(s):  
June Chung ◽  
Jeonghwan Shim ◽  
Ki D. Lee
Keyword(s):  
High Speed ◽  
Adjoint Method ◽  
Design Method ◽  
Computational Cost ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Internal Flow ◽  
Adjoint System ◽  
Inverse Design ◽  
Compressor Blades

A three-dimensional (3D) CFD-based design method for high-speed axial compressor blades is being developed based on the discrete adjoint method. An adjoint code is built corresponding to RVC3D, a 3D turbomachinery Navier-Stokes analysis code developed at NASA Glenn. A validation study with the Euler equations indicates that the adjoint sensitivities are sensitive to the choice of boundary conditions for the adjoint variables in internal flow problems and constraints may be needed on internal boundaries to capture proper physics of the adjoint system. The design method is demonstrated with inverse design based on Euler physics, and the results indicate that the adjoint design method produces efficient 3D designs by drastically reducing the computational cost.

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

2016 ◽  
Vol 33 (2) ◽  
Author(s):  
Jia Yu ◽  
Lucheng Ji ◽  
Weiwei Li ◽  
Weilin Yi
Keyword(s):  
Pressure Distribution ◽  
Adjoint Method ◽  
Static Pressure ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Optimal Solution ◽  
Optimization Procedure ◽  
Compressor Blades ◽  
Static Pressure Distribution ◽  
Practical Engineering

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

2015 ◽  
Author(s):  
Jia Yu ◽  
Lucheng Ji ◽  
Weiwei Li ◽  
Weilin Yi
Keyword(s):  
Pressure Distribution ◽  
Adjoint Method ◽  
Static Pressure ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Optimal Solution ◽  
Optimization Procedure ◽  
Compressor Blades ◽  
Static Pressure Distribution ◽  
Multi Stage

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.

scholarly journals Multi-objective aerodynamic optimization of flying-wing configuration based on adjoint method

2021 ◽  
Vol 39 (4) ◽  
pp. 753-760
Author(s):  
Xiaodong Liu ◽  
Peiliang Zhang ◽  
Guanghong He ◽  
Yongen Wang ◽  
Xudong Yang
Keyword(s):  
Optimization Design ◽  
Adjoint Method ◽  
Stokes Equations ◽  
Design Method ◽  
Adjoint Equation ◽  
Geometric Constraints ◽  
Navier Stokes ◽  
Aerodynamic Design ◽  
Aerodynamic Optimization ◽  
Multi Objective

In order to solve the multi-objective multi-constraint design in aerodynamic design of flying wing, the aerodynamic optimization design based on the adjoint method is studied. In terms of the principle of the adjoint equation, the boundary conditions and the gradient equations are derived. The Navier-Stokes equations and adjoint aerodynamic optimization design method are adopted, the optimization design of the transonic drag reduction for the two different aspect ratio of the flying wing configurations is carried out. The results of the optimization design are as follows: Under the condition of satisfying the aerodynamic and geometric constraints, the transonic shock resistance of the flying wing is weakened to a great extent, which proves that the developed method has high optimization efficiency and good optimization effect in the multi-objective multi-constraint aerodynamic design of the flying wing.

scholarly journals The Effect of Adding Roughness and Thickness to a Transonic Axial Compressor Rotor

1995 ◽  
Vol 117 (4) ◽  
pp. 491-505 ◽  
Author(s):  
K. L. Suder ◽  
R. V. Chima ◽  
A. J. Strazisar ◽  
W. B. Roberts
Keyword(s):  
Surface Roughness ◽  
Surface Finish ◽  
Axial Compressor ◽  
Leading Edge ◽  
Aerodynamic Performance ◽  
Blade Surface ◽  
Suction Surface ◽  
Compressor Rotor ◽  
Performance Deterioration ◽  
Design Speed

The performance deterioration of a high-speed axial compressor rotor due to surface roughness and airfoil thickness variations is reported. A 0.025 mm (0.001 in.) thick rough coating with a surface finish of 2.54–3.18 rms μm (100–125 rms μin.) is applied to the pressure and suction surface of the rotor blades. Coating both surfaces increases the leading edge thickness by 10 percent at the hub and 20 percent at the tip. Application of this coating results in a loss in efficiency of 6 points and a 9 percent reduction in the pressure ratio across the rotor at an operating condition near the design point. To separate the effects of thickness and roughness, a smooth coating of equal thickness is also applied to the blade. The smooth coating surface finish is 0.254–0.508 rms μm (10–20 rms μin.), compared to the bare metal blade surface finish of 0.508 rms pm (20 rms μin.). The smooth coating results in approximately half of the performance deterioration found from the rough coating. Both coatings are then applied to different portions of the blade surface to determine which portions of the airfoil are most sensitive to thickness/roughness variations. Aerodynamic performance measurements are presented for a number of coating configurations at 60, 80, and 100 percent of design speed. The results indicate that thickness/roughness over the first 2 percent of blade chord accounts for virtually all of the observed performance degradation for the smooth coating, compared to about 70 percent of the observed performance degradation for the rough coating. The performance deterioration is investigated in more detail at design speed using laser anemometer measurements as well as predictions generated by a quasi-three-dimensional Navier–Stokes flow solver, which includes a surface roughness model. Measurements and analysis are performed on the baseline blade and the full-coverage smooth and rough coatings. The results indicate that adding roughness at the blade leading edge causes a thickening of the blade boundary layers. The interaction between the rotor passage shock and the thickened suction surface boundary layer then results in an increase in blockage, which reduces the diffusion level in the rear half of the blade passage, thus reducing the aerodynamic performance of the rotor.

Flow Characteristics of an Optimized Axial Compressor Rotor Using Smooth Design Parameters

2016 ◽  
Author(s):  
Syed Moez Hussain Mahmood ◽  
Mark G. Turner ◽  
Kiran Siddappaji
Keyword(s):  
Thickness Distribution ◽  
Axial Compressor ◽  
Leading Edge ◽  
Design Tool ◽  
Smooth Transition ◽  
Design Parameters ◽  
Spanwise Direction ◽  
Aerodynamic Optimization ◽  
Blade Shape ◽  
Smooth Design

Blade designs have evolved from NACA series and free vortex assumptions to detailed meanline and forced vortex definitions. A design process is presented with numerous parametric options to explore a large design space. Smoothness in turbomachinery blade shapes is critical to an effective design. A cubic B-spline is used to control spanwise variations in the curvature definition of airfoil camber, thickness distribution, leading edge definition, inlet angles and outlet angles as parameters with a small number of control points. Varying parameters of individual blade sections requires more control variables that increases the parameter space and adds kinks in the 3D blade shape. Benefits of this smooth spanwise capability are demonstrated by linking the blade design tool with an aerodynamic optimization system. A single subsonic rotor (rotor 6 of a 10 stage axial compressor derived from the GE EEE design) has been considered as the baseline for the optimization process. Optimization is performed by varying curvature of the airfoil camberline as well as inlet and outlet angles in the spanwise direction. A single objective optimization was performed to optimize isentropic efficiency. An improvement in efficiency of 0.83% from 91.87% to 92.63% was obtained. The optimized blade geometry has a smooth transition from a traditional airfoil shape at the hub section to an S-shaped airfoil at the mid and tip sections. This unique blade shape was obtained because the airfoil camber curvature definition was allowed to vary smoothly spanwise. An S-shaped blade near the mid and tip section promotes flow to move radially downwards which allows for a reduction in entropy generation due to tip leakage flows. Entropy is used to quantify losses and improvement in efficiency.

Aerodynamic Optimization Design of Exhaust Hood Diffuser for Steam Turbine With Three-Dimensional Reynolds-Averaged Navier-Stokes Solutions

2014 ◽  
Author(s):  
Jiandao Yang ◽  
Taowen Chen ◽  
Jun Li ◽  
Zhenping Feng
Keyword(s):  
Optimization Design ◽  
Static Pressure ◽  
Three Dimensional ◽  
Aerodynamic Performance ◽  
Navier Stokes ◽  
Exhaust Hood ◽  
Aerodynamic Optimization ◽  
Recovery Coefficient ◽  
Last Stage ◽  
Pressure Recovery Coefficient

Combined with three-dimensional parameterization method of exhaust diffuser profile, aerodynamic performance evaluation method, response surface approximation evaluation model and Hooke-Jeeves direct search approach, aerodynamic optimization design of exhaust hood diffuser for steam turbine is presented. The aerodynamic performance of exhaust hood design candidate is evaluated using three-dimensional Reynolds-Averaged Navier-Stokes (RANS) solutions. Aerodynamic optimization design of exhaust hood is conducted for the maximum of the static pressure recovery coefficient of exhaust hood. The design variables are specified by the exhaust diffuser profile parameterization method. The aerodynamic performance of the optimized exhaust hood and referenced design is numerically calibrated with consideration of the full last stage and rotor tip clearance. The static pressure recovery coefficient of the optimized exhaust hood is higher than that of the referenced design with consideration of the upstream last stage influence. Furthermore, the detailed flow pattern of the optimized exhaust hood and referenced design is also analyzed and compared.

Aerodynamic Optimization of Industrial Fan Blade Cascades

2005 ◽  
Author(s):  
O. Lotfi ◽  
J. A. Teixeira ◽  
P. C. Ivey ◽  
G. Sheard ◽  
I. R. Kinghorn
Keyword(s):  
Genetic Algorithm ◽  
Flow Rate ◽  
Thickness Distribution ◽  
Automated Design ◽  
Axial Fan ◽  
Aerodynamic Optimization ◽  
Flow Solver ◽  
Blade Thickness ◽  
Fitness Value ◽  
Fan Blade

The paper describes the application of a Genetic Algorithm (GA) to the aerodynamic shape optimization of a low speed fan cascade. This task is accomplished by modifying the blade camber line while keeping the same blade thickness distribution and mass flow rate. A number of design examples have been studied and the behaviour of the genetic algorithm has been tested. The fitness value of the objective function is evaluated using the commercial turbomachinery CFD code CFX-TASCflow. In this work Bezier curves were used for the description of the blade camber line. Specific interfaces were developed in order to link the optimization code, the grid generator TASCgrid utilized to define the computational meshes and the commercial flow solver TASCflow employed within an automated design loop. The obtained results show that the genetic algorithm is capable of automatically improving the design of axial fan blade profiles.

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