scholarly journals The Effects of Lean and Sweep on Transonic Fan Performance

2002 ◽  
Author(s):  
J. D. Denton ◽  
L. Xu
Keyword(s):  
High Efficiency ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Peak Efficiency ◽  
3D Design ◽  
Stall Margin ◽  
Fan Performance ◽  
Transonic Fan ◽  
Shock Loss ◽  
Design Techniques

The aerodynamics of transonic fans is discussed with emphasis on the use of three-dimensional design techniques, such as blade sweep and lean, to improve their performance. In order to study the interaction of these 3D features with the shock pattern a series of five different designs is produced and analysed by CFD. It is found that the 3D features have remarkably little effect on the shock pattern near the tip where the shock must remain perpendicular to the casing. Lower down the blade significant shock sweep, and hence reduced shock loss, can be induced by 3D design but this is usually at the expense of reduced stall margin and increased loss elsewhere along the blade span. Overall, very little change in peak efficiency or pressure ratio is produced by blade sweep or lean. However, there are significant effects on stall margin with forwards sweep producing a better stall margin and maintaining a high efficiency over a wider range.

Automatic Three-Dimensional Optimisation of a Modern Tandem Compressor Vane

2014 ◽  
Author(s):  
R. C. Schlaps ◽  
S. Shahpar ◽  
V. Gümmer
Keyword(s):  
Pressure Ratio ◽  
Three Dimensional ◽  
Trust Region ◽  
Automatic Generation ◽  
Navier Stokes ◽  
Design Parameters ◽  
High Fidelity ◽  
Stall Margin ◽  
Design Choice ◽  
Multipoint Approximation

In order to increase the performance of a modern gas turbine, compressors are required to provide higher pressure ratio and avoid incurring higher losses. The tandem aerofoil has the potential to achieve a higher blade loading in combination with lower losses compared to single vanes. The main reason for this is due to the fact that a new boundary layer is generated on the second blade surface and the turning can be achieved with smaller separation occurring. The lift split between the two vanes with respect to the overall turning is an important design choice. In this paper an automated three-dimensional optimisation of a highly loaded compressor stator is presented. For optimisation a novel methodology based on the Multipoint Approximation Method (MAM) is used. MAM makes use of an automatic design of experiments, response surface modelling and a trust region to represent the design space. The CFD solutions are obtained with the high-fidelity 3D Navier-Stokes solver HYDRA. In order to increase the stage performance the 3D shape of the tandem vane is modified changing both the front and rear aerofoils. Moreover the relative location of the two aerofoils is controlled modifying the axial and tangential relative positions. It is shown that the novel optimisation methodology is able to cope with a large number of design parameters and produce designs which performs better than its single vane counterpart in terms of efficiency and numerical stall margin. One of the key challenges in producing an automatic optimisation process has been the automatic generation of high-fidelity computational meshes. The multi block-structured, high-fidelity meshing tool PADRAM is enhanced to cope with the tandem blade topologies. The wakes of each aerofoil is properly resolved and the interaction and the mixing of the front aerofoil wake and the second tandem vane are adequately resolved.

scholarly journals A Three-Dimensional Shock Loss Model Applied to an Aft-Swept, Transonic Compressor Rotor

1996 ◽  
Author(s):  
Steven L. Puterbaugh ◽  
William W. Copenhaver ◽  
Chunill Hah ◽  
Arthur J. Wennerstrom
Keyword(s):  
Flow Rate ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Numerical Results ◽  
Design Flow ◽  
Design System ◽  
Loss Model ◽  
Transonic Compressor ◽  
Compressor Rotor ◽  
Shock Loss

An analysis of the effectiveness of a three-dimensional shock loss model used in transonic compressor rotor design is presented. The model was used during the design of an aft-swept, transonic compressor rotor. The demonstrated performance of the swept rotor, in combination with numerical results, is used to determine the strengths and weaknesses of the model. The numerical results were obtained from a fully three-dimensional Navier-Stokes solver. The shock loss model was developed to account for the benefit gained with three-dimensional shock sweep. Comparisons with the experimental and numerical results demonstrated that shock loss reductions predicted by the model due to the swept shock induced by the swept leading edge of the rotor were exceeded. However, near the tip the loss model under-predicts the loss because the shock geometry assumed by the model remains swept in this region while the numerical results show a more normal shock orientation. The design methods and the demonstrated performance of the swept rotor is also presented. Comparisons are made between the design intent and measured performance parameters. The aft-swept rotor was designed using an inviscid axisymmetric streamline curvature design system utilizing arbitrary airfoil blading geometry. The design goal specific flow rate was 214.7 kg/sec/m2 (43.98 lbm/sec/ft2), the design pressure ratio goal was 2.042, and the predicted design point efficiency was 94.0. The rotor tip sped was 457.2 m/sec (1500 ft/sec). The design flow rate was achieved while the pressure ratio fell short by 0.07. Efficiency was 3 points below prediction, though at a very high 91 percent. At this operating condition the stall margin was 11 percent.

scholarly journals Design and Optimization of Multiple Circumferential Casing Grooves Distribution Considering Sweep and Lean Variations on the Blade Tip

Energies ◽  
2018 ◽  
Vol 11 (9) ◽  
pp. 2401
Author(s):  
Weimin Song ◽  
Yufei Zhang ◽  
Haixin Chen
Keyword(s):  
Pressure Ratio ◽  
Three Dimensional ◽  
Axial Distribution ◽  
Stall Margin ◽  
Design Point ◽  
Casing Treatment ◽  
Design And Optimization ◽  
Practical Engineering ◽  
Engineering Strategy ◽  
Blade Tip

This paper focuses on the design and optimization of the axial distribution of the circumferential groove casing treatment (CGCT). Effects of the axial location of multiple casing grooves on the flow structures are numerically studied. Sweep and lean variations are then introduced to the blade tip, and their influences on the grooves are discussed. The results show that the ability of the CGCT to relieve the blockage varies with the distribution of grooves, and the three-dimensional blading affects the performance of both the blade and the CGCT. Accordingly, a multi-objective optimization combining the CGCT design with the sweep and lean design is conducted. Objectives, including the total pressure ratio and the adiabatic efficiency, are set at the design point; meanwhile, the choking mass flow and the near-stall performance are constrained. The coupling between the CGCT and the blade is improved, which contributes to an optimal design point performance and a sufficient stall margin. The sweep and lean in the tip redistribute the spanwise and chordwise loading, which enhances the ability of the CGCT to improve the blade’s performance. This work shows that the present CGCT-blade integrated optimization is a practical engineering strategy to develop the working capacity and efficiency of a compressor blade while achieving the stall margin extension.

scholarly journals Optimization Design of a 2.5 Stage Highly Loaded Axial Compressor with a Bezier Surface Modeling Method

2020 ◽  
Vol 10 (11) ◽  
pp. 3860
Author(s):  
Song Huang ◽  
Jinxin Cheng ◽  
Chengwu Yang ◽  
Chuangxin Zhou ◽  
Shengfeng Zhao ◽  
...  
Keyword(s):  
Flow Field ◽  
Dimensional Space ◽  
Three Dimensional ◽  
Axial Compressor ◽  
High Dimensions ◽  
Peak Efficiency ◽  
Stall Margin ◽  
Second Stage ◽  
Highly Loaded ◽  
Three Dimensional Space

Due to the complexity of the internal flow field of compressors, the aerodynamic design and optimization of a highly loaded axial compressor with high performance still have three problems, which are rich engineering design experience, high dimensions, and time-consuming calculations. To overcome these three problems, this paper takes an engineering-designed 2.5-stage highly loaded axial flow compressor as an example to introduce the design process and the adopted design philosophies. Then, this paper verifies the numerical method of computational fluid dynamics. A new Bezier surface modeling method for the entire suction surface and pressure surface of blades is developed, and the multi-island genetic algorithm is directly used for further optimization. Only 32 optimization variables are used to optimize the rotors and stators of the compressor, which greatly overcome the problem of high dimensions, time-consuming calculations, and smooth blade surfaces. After optimization, compared with the original compressor, the peak efficiency is still improved by 0.12%, and the stall margin is increased by 2.69%. The increase in peak efficiency is mainly due to the rotors. Compared with the original compressor, for the second-stage rotor, the adiabatic efficiency is improved by about 0.4%, which is mainly due to the decreases of total pressure losses in the range of above 30% of the span height and 10%–30% of the chord length. Besides, for the original compressor, due to deterioration of the flow field near the tip region of the second-stage stator, the large low-speed region eventually evolves from corner separation into corner stall with three-dimensional space spiral backflow. For the optimized compressor, the main reason for the increased stall margin is that the flow field of the second-stage stator with a span height above 50% is improved, and the separation area and three-dimensional space spiral backflow are reduced.

Multi-Objective and Multi-Point Aerodynamic Optimization of Transonic Fan Blades

2014 ◽  
Author(s):  
Maryam Khelghatibana ◽  
Jean-Yves Trépanier ◽  
Christophe Tribes ◽  
Jason Nichols
Keyword(s):  
High Speed ◽  
Stokes Equations ◽  
Pressure Ratio ◽  
Operating Conditions ◽  
Optimized Design ◽  
Aerodynamic Optimization ◽  
Stall Margin ◽  
Choked Flow ◽  
Multi Objective ◽  
Transonic Fan

A multi-objective and multi-point optimization methodology is developed for aerodynamic design of transonic fan blades. The optimization method aims to increase design efficiency, near stall efficiency and stall margin while maintaining the required design pressure ratio and high speed choke margin. Numerical analyses are performed by solving three-dimensional Reynolds-Averaged Navier-Stokes equations combined with shear stress turbulence model. A multi-level blade parameterization is employed to modify the blade geometry. The proposed method is applied to redesign NASA rotor 67. First, an optimization case with considering two operating conditions at peak efficiency and near stall is performed to demonstrate the relation between near stall efficiency and stall margin. An investigation on Pareto optimal solutions of this optimization shows that the stall margin is increased with improving near stall efficiency. Then, in order to maintain the required choke margin, an operating point at high speed choked flow is added to the optimization process. A final optimized design is selected by considering the interaction of design requirements at all three operating points. The new design presents higher efficiency and stall margin without any reduction in the chocking mass flow rate.

Numerical Investigation on the Effect of Bleed Port With Self-Recirculating Casing Treatment on the Stability of a 1.5-Stage Transonic Compressor

2019 ◽  
Author(s):  
Yiming Zhong ◽  
WuLi Chu ◽  
HaoGuang Zhang
Keyword(s):  
Pressure Ratio ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Stability Margin ◽  
Axial Position ◽  
Stable Operation ◽  
Stall Margin ◽  
Casing Treatment ◽  
Transonic Compressor ◽  
Compressor Efficiency

Abstract Compared to the traditional casing treatment, the self-recirculating casing treatment (SCT) can improve or not decrease the compressor efficiency while achieving the stall margin improvement. For the bleed port, the main design indicator is to reduce the flow loss caused by suction, while providing sufficient jet flow and jet pressure to the injector. In order to gain a better study of the bleed port stabilization mechanisms, the bleed configuration was parameterized with the bleed port inlet width and the bleed port axial position. Five kinds of recirculating casing treatments were applied to a 1.5-stage transonic axial compressor with the method of three-dimensional unsteady numerical simulation. Fifteen identical self-recirculating devices are uniformly mounted around the annulus. The numerical results show that the SCT can improve compressor total pressure ratio and stability, shift the stall margin towards lower mass flows. Furthermore, it has no impact on compressor efficiency. The optimal case presents that stability margin is improved by 6.7% employing 3.1% of the annulus mass flow. Expanding bleed port inlet width to an intermediate level can further enhance compressor stability, but excessive bleed port inlet width will reduce the stabilization effect. The optimal bleed port position is located in the blocked area of the low energy group at the top of the rotor. In the case of solid casing, stall inception was the tip blockage, which was mainly triggered by the interaction of the tip leakage vortex and passage shock. From radial distribution, the casing treatment predominantly affects the above 70% span. The reduction of tip reflux region by suction effect is the main reason for the extension of stable operation range. The SCT also has an obvious stability improvement in tip blockage stall, while delaying the occurrence of compressor stall.

Flow Field Analysis of Inlet Distortion in a Transonic Fan With Straight and Bowed Stator Blade

2007 ◽  
Author(s):  
Peng Sun ◽  
Jingjun Zhong ◽  
Guotai Feng
Keyword(s):  
Flow Field ◽  
Total Pressure ◽  
Three Dimensional ◽  
Field Analysis ◽  
Navier Stokes ◽  
Fan Performance ◽  
Inlet Distortion ◽  
Flow Loss ◽  
Stator Blade ◽  
Transonic Fan

The performance and stability of a fan in clean and distorted inlet flow can be improved through the use of bowed stator blades. Measurements between the blade rows in transonic and supersonic flow are too complex to provide any useful insights, so 3D flow simulations are required. In this paper, a time-accurate three-dimensional Navier-Stokes solver of the unsteady flow field in a transonic fan is carried out using “Fluent-parallel” in a parallel supercomputer. Two sets of simulations are performed. The first simulation focuses on a better understanding of inlet total pressure distortion effects on a transonic fan. The second set of numerical simulation aims at studying the improvements of fan performance made by bowed stator blades. Three aspects are contained in this paper. The first is about the distortion effects on characteristics of the fan stage with straight stator. The effects of bowed stator on fan performance with inlet distortion are demonstrated secondly. One hand bowed stator increases the loss in rotor. On the other hand, it reduces the flow loss in stator. Finally, the patterns of flow loss caused by total pressure distortion with straight/bowed stator are compared. The scale of vortex in stator induced by inlet total pressure distortion is weakened by bowed blades, which decreases the stator loss.

An Optimum Aero-Design Process of Turbine Blades

2002 ◽  
Author(s):  
C. Xu ◽  
R. S. Amano
Keyword(s):  
Design Process ◽  
High Efficiency ◽  
Pressure Ratio ◽  
Turbine Blades ◽  
Three Dimensional ◽  
Optimization Procedure ◽  
Compressor Blade ◽  
Advanced Technology ◽  
Test Case ◽  
Blade Design

With the development of the advanced technology, the combustion temperature is raised for increased efficiencies. At the same time, the turbine and compressor pressure ratio and the mass flow rate rise; thus causing turbine and compressor blades turning and blade lengths increase. Moreover, the high efficiency requirements had made the turbine and compressor blade design difficult. A turbine airfoil has been custom designed for many years, but an optimization for the section design in a three-dimensional consideration is still a challenge. For a compressor blade design, standard section cannot meet the modern compressor requirements. Modern compressor design has not only needs a custom designed section according to flow situation, but also needs three-dimensional optimizations. Therefore, a good blade design process is critical to the turbines and compressors. A blade design of the turbomachines is one of the important steps for a good turbomachine design. A blade design process not only directly influences the overall machine efficiency but also dramatically impact the design time and cost. In this study, a blade design and optimization procedure was proposed for both turbine and compressor blade design. A compressor blade design was used as a test case. It was shown that the current design process had more advantages than conventional design methodology.

Design Optimization of Circumferential Casing Grooves for a Transonic Axial Compressor to Enhance Stall Margin

2010 ◽  
Author(s):  
Kwang-Jin Choi ◽  
Jin-Hyuk Kim ◽  
Kwang-Yong Kim
Keyword(s):  
Design Optimization ◽  
Numerical Solutions ◽  
Stokes Equations ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Latin Hypercube Sampling ◽  
Optimum Shape ◽  
Data Set ◽  
Stall Margin

This paper presents a design optimization of an axial compressor with NASA Rotor 37 and five circumferential casing grooves for enhancement of stall margin. Three-dimensional Reynolds-averaged Navier-Stokes equations with the shear stress transport turbulence model are discretized by finite volume approximations and solved on hexahedral grids for the flow analyses. The validation of the numerical results is performed in comparison with experimental data for pressure ratio and adiabatic efficiency. The Latin-hypercube sampling as design-of-experiments is used to generate the twelve design points within the design space. A stall margin parameter is considered as an objective function with two design variables defining the geometry of the circumferential casing grooves. The radial basis neural network method employed as a surrogate model for the design optimization of the circumferential casing grooves is trained on the numerical solutions by carrying out leave-one-out cross-validation for the data set. The results show that the stall margin of the optimum shape is enhanced considerably by the design optimization compared to the cases with smooth casing and the reference grooves.

scholarly journals The Effect of the Inlet Boundary Layer on the Tip Flow Field in a Transonic Fan Rotor

1998 ◽  
Author(s):  
Junji Takado ◽  
Toyotaka Sonoda ◽  
Satoshi Nakamura
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Interaction Region ◽  
Operating Conditions ◽  
Laser Doppler Velocimeter ◽  
Intensity Level ◽  
Pressure Surface ◽  
Transonic Fan

Experimental and numerical investigations have been carried out to understand the effects of the inlet boundary layer (IBL) on the tip flow field including the aerodynamic performance in a transonic fan rotor. Both the steady and the unsteady phenomena in the tip flow field have been investigated for operating conditions near peak efficiency and near stall with the two types of tip IBL. In order 10 study these phenomena, high response pressure data with Kulite transducers and laser doppler velocimeter (LDV) data have been acquired around the tip region. Furthermore, three-dimensional Navier-Stokes numerical simulations have been compared with the measured results. The results indicate that the tip IBL significantly influences the spanwise distribution of pressure ratio around the tip region and the stall characteristics including the passage shock / tip leakage vortex interaction, the blockage generation, the wake structure, and the unsteadiness of the tip flow field. In particular, at a near stall condition for the thick IBL with high turbulence intensity level, the tip diffusion level is increased due to a larger blockage, which is generated downstream of a much stronger interaction region. These phenomena are a consequence of the low momentum fluid in the tip IBL, and significantly reduce the stall margin. Furthermore, the unsteadiness drastically increases around the interaction region and around the pressure surface where the blockage migrates. These unsteady phenomena are distinctive features near stall. Downstream of the rotor, the larger and more unsteady blockage is discharged from the pressure surface side, and complicates the three-dimensional rotor exit flow field around the tip region.

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