沿流向放样的轴流压气机叶道设计方法

赵奕博; 李浩; 琚亚平; 张楚华

doi:10.7652/xjtuxb202304005

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沿流向放样的轴流压气机叶道设计方法

—————————————————————— | 更新时间：2025-07-09

- 沿流向放样的轴流压气机叶道设计方法
- Design Method of Blade Passage for Axial Compressor by Lofting along Flow Direction
- 西安交通大学学报 2023年57卷第4期页码：39-48
- 作者机构：
  
  1. 西安交通大学能源与动力工程学院,西安,710049
  2. 上海交通大学机械与动力工程学院,上海,200240
  3. 西安交通大学机械结构强度与振动国家重点实验室,西安,710049
- 作者简介：
  
  : 2022-10-06。作者简介: 赵奕博(2000—),男,博士生
  琚亚平(通信作者),女,教授,博士生导师。基金项目: 国家重大科技专项资助项目(J2019-Ⅱ-0017-0038, 2017-Ⅱ-0006-0020)
- 基金信息：
  
  : 2022-10-06。作者简介: 赵奕博(2000—),男,博士生;琚亚平(通信作者),女,教授,博士生导师。基金项目: 国家重大科技专项资助项目(J2019-Ⅱ-0017-0038, 2017-Ⅱ-0006-0020);机械结构强度与振动国家重点实验室资助项目(SV2021-ZZ-23)。
- DOI：10.7652/xjtuxb202304005
  中图分类号： TK05
- 网络首发：2023-04-10，
  
  纸质出版：2023
- 稿件说明：
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赵奕博, 李浩, 琚亚平, 等. 沿流向放样的轴流压气机叶道设计方法[J]. 西安交通大学学报, 2023,57(4):39-48.

ZHAO Yibo, LI Hao, JU Yaping, et al. Design Method of Blade Passage for Axial Compressor by Lofting along Flow Direction[J]. 2023, 57(4): 39-48.
赵奕博, 李浩, 琚亚平, 等. 沿流向放样的轴流压气机叶道设计方法[J]. 西安交通大学学报, 2023,57(4):39-48. DOI： 10.7652/xjtuxb202304005.

ZHAO Yibo, LI Hao, JU Yaping, et al. Design Method of Blade Passage for Axial Compressor by Lofting along Flow Direction[J]. 2023, 57(4): 39-48. DOI： 10.7652/xjtuxb202304005.

摘要

针对高负荷压气机角区流动质量差、流动损失显著的问题

以NASA Rotor37为研究对象

提出了一种轴流压气机叶道几何定义与参数化建模新方法

实现了叶片、角区和端壁结构的协同设计及叶片到端壁的高阶光滑过渡

在此基础上

开展了前缘边条和叶身融合对压气机转子气动性能和内部流场影响的数值研究。结果表明

针对NASA Rotor37流道

采用该叶道设计方法对压气机转子角区改型后:绝热效率提高了0.3%

喘振裕度由11.85%增加至18.8%; 下凹端壁可以帮助稳定压气机转子通流能力并减少改型过程的流量漂移; 叶身融合结构在叶根处提供径向向上的压力梯度

将低能流体从叶片角区吹向主流区以减少角区流动损失

而近叶根处主流区流动损失略有增加。研究成果对发展轴流压气机先进设计方法应具有参考价值。

Abstract

Aiming at the problems of poor flow quality and significant flow loss at corner region of axial compressors

NASA Rotor37 axial compressor was taken as the research object to propose a new geometry definition and parameterization method for the blade passage of axial compressor

which enabled the cooperative design of blade

corner and endwall structure

and realized the high order smooth transition between blade and endwall. On this basis

the influence on the aerodynamic performance and internal flow field of hub fillet were studied. The results showed that after modifying the blade passage of NASA Rotor37

the compressor adiabatic efficiency increased by 0.3% at the design point

and surge margin increased from 11.85% to 18.8%. The concave on endwall can help stabilize the compressor flow capacity and avoid flow drift during modification. The blended blade and endwall structure can produce a transverse pressure difference

removing the low energy fluid from the corner region to the main flow to reduce corner flow separation

while the main flow losses slightly increased. The new geometry defining and parametric modeling method of the axial compressor proposed in this paper and the findings of flow improvement mechanism of the corner flow have reference value for the further development of axial compressor design system.

关键词

Keywords

references

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