scholarly journals A Quasi-Three Dimensional Calculation System for the Flow Within Transonic Compressor Blade Rows

1985 ◽  
Author(s):  
W. J. Calvert ◽  
R. B. Ginder
Keyword(s):  
Data Transfer ◽  
Three Dimensional ◽  
Internal Flow ◽  
Compressor Blade ◽  
Meridional Plane ◽  
End Effects ◽  
Transonic Compressor ◽  
Calculation System ◽  
Set Up ◽  
Transonic Fan

A calculation system has been set up to predict both the internal flow field and the overall performance of a transonic compressor blade row. The system iterates between an inviscid-viscous time-marching blade-to-blade (S1) treatment and a streamline curvature throughflow calculation for the pitchwise-averaged flow in the meridional plane (S2). A blade geometry package and a data transfer/display program are used to link the S1 and S2 methods to give a semi-automatic convergence procedure. The only empirically-based correlation or correction required is an extra loss imposed near the blade hub and tip to allow for end effects. The system has been applied to a high bypass ratio transonic fan rotor near design point. The converged solution was in good agreement with the measured performance.

Detailed Flow Measurements and Predictions for a Three-Stage Transonic Fan

1994 ◽  
Vol 116 (2) ◽  
pp. 298-305 ◽  
Author(s):  
W. J. Calvert ◽  
A. W. Stapleton
Keyword(s):  
Shock Waves ◽  
Computational Efficiency ◽  
Three Dimensional ◽  
Internal Flow ◽  
Design Tool ◽  
Three Dimensional Flow ◽  
Flow Measurements ◽  
Overall Performance ◽  
Calculation System ◽  
Transonic Fan

Detailed flow measurements were taken at DRA Pyestock on a Rolls-Royce three-stage transonic research fan using advanced laser transit velocimetry and holography techniques to supplement the fixed pressure and temperature instrumentation. The results have been compared with predictions using the DRA S1-S2 quasi-three-dimensional flow calculation system at a range of speeds. The agreement was generally encouraging, both for the overall performance and for details of the internal flow such as positions of shock waves. Taken together with the computational efficiency of the calculations and previous experience on single-stage transonic fans and core compressors, this establishes the S1-S2 system as a viable design tool for future multistage transonic fans.

The Design of an Advanced Civil Fan Rotor

1987 ◽  
Vol 109 (3) ◽  
pp. 340-345 ◽  
Author(s):  
R. B. Ginder ◽  
W. J. Calvert
Keyword(s):  
High Efficiency ◽  
Three Dimensional ◽  
Internal Flow ◽  
Blade Profile ◽  
Transonic Compressor ◽  
Blade Row ◽  
New Type ◽  
Overall Performance ◽  
Calculation System ◽  
Asme Paper

A recent ASME paper by the authors described a quasi-three-dimensional calculation system for transonic compressor blade rows. The system predicts both the internal flow field and the overall performance of the blade row. It therefore enables the compressor engineer to optimize the blade shapes in order to improve the design point efficiency. This is explored in the present paper. A new type of blade profile has been developed to allow sufficient freedom for the optimization. Application to the design of a high-efficiency, transonic civil fan rotor is discussed.

Development of an Improved Streamline Curvature Approach for Transonic Axial Compressor

2016 ◽  
Author(s):  
Wu Xiaoxiong ◽  
Bo Liu ◽  
Shi Lei ◽  
Zhang Guochen ◽  
Mao Xiaochen
Keyword(s):  
Incidence Angle ◽  
Fluid Dynamic ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Internal Flow ◽  
Field Analysis ◽  
Design Stage ◽  
Axial Compressors ◽  
Transonic Compressor ◽  
Streamline Curvature

In this paper, an improved streamline curvature (SLC) approach is presented to obtain the internal flow fields and evaluate the performance of transonic axial compressors. The approach includes some semi-empirical correlations established based on previous literatures, such as minimum loss incidence angle model, deviation model and total pressure loss model. Several developments have been made in this paper for the purpose of considering the influences of three-dimensional (3D) flow in high-loaded multistage compressors with high accuracy. A revised deviation model is applied to predict the cascade with large deflection range. The method for predicting the shock loss is also discussed in detail. In order to validate the reliability of the approach, two test cases including a two-stage transonic fan and a three-stage transonic compressor are conducted. The overall performance and distribution of spanwise aerodynamic parameters are illustrated in this paper. Compared with both the experimental and computational fluid dynamic (CFD) data at design and a number of different off-design condition, the SLC results give reasonable characteristic curves. The validation demonstrates that this improved approach can serve as a fast and reliable tool for flow field analysis and performance prediction in preliminary design stage of axial compressors.

Multiobjective Optimization Approach to Multidisciplinary Design of a Three-Dimensional Transonic Compressor Blade

2009 ◽  
Author(s):  
Chang Luo ◽  
Liming Song ◽  
Jun Li ◽  
Zhenping Feng
Keyword(s):  
Multiobjective Optimization ◽  
Optimization Algorithm ◽  
Performance Optimization ◽  
Three Dimensional ◽  
Compressor Blade ◽  
Multidisciplinary Design ◽  
Optimization Approach ◽  
Pareto Solution ◽  
Transonic Compressor ◽  
Turbomachinery Blades

An automatic multiobjective optimization approach to multidisciplinary design of turbomachinery blades is proposed in this paper. Based on this approach, an algorithm named Multiobjective Differential Evolution (MDE) is introduced as an optimizer to find the Pareto solution sets of the multidisciplinary design problem. A typical multiobjective function has been applied to demonstrate the performance of the presented multiobjective optimization algorithm. The Non-uniform B-Spline method is adopted to parameterize the turbomachinery blade profiles. The aerodynamic performance of design blade candidates is predicted by using a three–dimensional Reynolds-Averaged Navier-Stokes (RANS) solution. The blade stresses and vibration frequencies are evaluated by means of a finite element analysis coupled with the surface pressure of blades obtained from CFD calculation. To validate the optimization capability of the multiobjective optimization algorithm, the multidisciplinary design of a typical transonic compressor blade, NASA Rotor 37, is conducted. The blade is optimized for the maximization of the isentropic efficiency and the minimization of the maximum stresses with constraints on mass flow rate, total pressure ratio, and dynamic frequencies. The Pareto solutions are obtained from the multiobjective optimization. Based on the analysis of the design objectives between the Pareto designs and reference design, it is indicated that the overall performance of the optimized designs is improved. The results demonstrate that the presented multiobjective optimization algorithm has a potential in blade performance optimization and it is a promising method for the multidisciplinary design of turbomachinery blades.

Investigation of an innovative non-free vortex aerodynamic procedure to design a single-stage transonic compressor

2017 ◽  
Vol 232 (11) ◽  
pp. 2132-2143 ◽  
Author(s):  
A Shahsavari ◽  
M Nili-Ahmadabadi
Keyword(s):  
Fluid Dynamics ◽  
Computational Fluid Dynamics ◽  
Stokes Equations ◽  
Design Method ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Rotor Blades ◽  
Meridional Plane ◽  
Transonic Compressor ◽  
Performance Map

This paper presents an innovative design method for a transonic compressor based on the radial equilibrium theory by means of increasing blade loading. Firstly, the rotor blade of a transonic compressor is redesigned based on the constant spanwise de-Haller number and diffusion. The design method leads to an unconventional increased axial velocity distribution in tip section, which originates from non-uniform enthalpy distribution assumption. A code is applied to extract the compressor meridional plane and blade-to-blade geometry containing rotor and stator in order to design the blade three-dimensional view. A structured grid is generated for the numerical domain of fluid. Finer grids are used for the regions near walls to capture the boundary layer effects and behavior. Reynolds-averaged Navier–Stokes equations are solved by finite volume method for rotating zones (rotor) and stationary zones (stator). The experimental data, available for the performance map of NASA Rotor67, is used to validate the results of the current simulations. Then, the capability of the design method is validated by computational fluid dynamics that is capable of predicting the performance map. The numerical results of the new geometry by representing 11% improvement in efficiency and 19% in total pressure ratio verify the new method advantages. The computational fluid dynamics results also show that the newly designed rotor blades due to a higher velocity in the tip section have a special capacity to increase the loading without any separation. The mass flow reduction is observed in the new geometry, which could be easily improved by changing stagger angle.

scholarly journals Analysis of Flow Within Pump Impeller of Torque Converter

1996 ◽  
Author(s):  
Hoshio Tsujita ◽  
Shimpei Mizuki ◽  
Eiji Ejiri
Keyword(s):  
Coordinate System ◽  
Secondary Flow ◽  
Three Dimensional ◽  
Internal Flow ◽  
Torque Converter ◽  
Flow Patterns ◽  
Speed Ratio ◽  
Meridional Plane ◽  
Pump Impeller ◽  
Simple Method

It is difficult to measure flow patterns within rotating elements of a torque converter due to the complicated construction. Therefore, the numerical calculation is considered to be an effective tool to know the internal flow. Three-dimensional incompressible turbulent flow within a pump impeller of an automotive torque converter was analyzed numerically at three different speed ratios, 0.02, 0.4 and 0.8 under the same inlet boundary condition. The speed ratio was defined as the ratio of rotating speed of the turbine impeller to that of the pump. The governing equations using the k-ε model in the physical component tensor form were solved with a boundary-fitted coordinate system fixed on a rotating impeller. The solution algorithm was the SIMPLE method applied to the curvilinear coordinate system. The computed results were compared with those obtained experimentally by an oil film flow visualization technique for the pressure, suction, core and shell surfaces. Moreover, the results at three different speed ratios were examined in detail in order to clarify the behavior of secondary flow patterns. The computed results showed good agreement with the experimental results and clarified the behavior of the complicated flow patterns. The secondary flow patterns were strongly influenced by the correlation between the intensities of the Corinlis force (COF) and the centrifugal force due to the passage curvature in the meridional plane (CMF).

Development of Hub Corner Stall and Its Influence on the Performance of Axial Compressor Blade Rows

1999 ◽  
Vol 121 (1) ◽  
pp. 67-77 ◽  
Author(s):  
C. Hah ◽  
J. Loellbach
Keyword(s):  
Numerical Solutions ◽  
Stokes Equations ◽  
Flow Direction ◽  
Three Dimensional ◽  
Suction Side ◽  
Compressor Blade ◽  
Suction Surface ◽  
Simple Diffusion ◽  
Transonic Compressor ◽  
Compressor Rotor

A detailed investigation has been performed to study hub corner stall phenomena in compressor blade rows. Three-dimensional flows in a subsonic annular compressor stator and in a transonic compressor rotor have been analyzed numerically by solving the Reynolds-averaged Navier–Stokes equations. The numerical results and the existing experimental data are interrogated to understand the mechanism of compressor hub corner stall. Both the measurements and the numerical solutions for the stator indicate that a strong twisterlike vortex is formed near the rear part of the blade suction surface. Low-momentum fluid inside the hub boundary layer is transported toward the suction side of the blade by this vortex. On the blade suction surface near the hub, this vortex forces fluid to move against the main flow direction and a limiting stream surface is formed near the hub. The formation of this vortex is the main mechanism of hub corner stall. When the aerodynamic loading is increased, the vortex initiates further upstream, which results in a larger corner stall region. For the transonic compressor rotor studied in this paper, the numerical solution indicates that a mild hub corner stall exists at 100 percent rotor speed. The hub corner stall, however, disappears at the reduced blade loading, which occurs at 60 percent rotor design speed. The present study demonstrates that hub corner stall is caused by a three-dimensional vortex system and that it does not seem to be correlated with a simple diffusion factor for the blade row.

Development of Hub Corner Stall and its Influence on the Performance of Axial Compressor Blade Rows

1997 ◽  
Author(s):  
Chunill Hah ◽  
James Loellbach
Keyword(s):  
Numerical Solutions ◽  
Stokes Equations ◽  
Flow Direction ◽  
Three Dimensional ◽  
Suction Side ◽  
Compressor Blade ◽  
Suction Surface ◽  
Simple Diffusion ◽  
Transonic Compressor ◽  
Compressor Rotor

A detailed investigation has been performed to study hub corner stall phenomena in compressor blade rows. Three-dimensional flows in a subsonic annular compressor stator and in a transonic compressor rotor have been analyzed numerically by solving the Reynolds-averaged Navier-Stokes equations. The numerical results and the existing experimental data are interrogated to understand the mechanism of compressor hub corner stall. Both the measurements and the numerical solutions indicate that a strong twister-like vortex is formed near the rear part of the blade suction surface. Low momentum fluid inside the hub boundary layer is transported toward the suction side of the blade by this vortex. On the blade suction surface near the hub, this vortex forces fluid to move against the main flow direction and a limiting stream surface is formed near the hub. The formation of this vortex is the main mechanism of hub corner stall. When the aerodynamic loading is increased, the vortex initiates further upstream, which results in a larger corner stall region. For the transonic compressor rotor studied in this paper, the numerical solution and the measured data indicate that a mild hub corner stall exists at 100 percent rotor speed. The hub corner stall, however, disappears at the reduced blade loading which occurs at 60 percent rotor design speed. The present study demonstrates that hub corner stall is caused by a three-dimensional vortex system and that it does not seem to be correlated with a simple diffusion factor for the blade row.

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