某型火箭发动机涡轮转子流热固耦合强度及疲劳寿命分析

黄朝晖; 袁奇; 张弘斌; 李浦; 王振; 黄道琼

doi:10.7652/xjtuxb202208008

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某型火箭发动机涡轮转子流热固耦合强度及疲劳寿命分析

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

- 某型火箭发动机涡轮转子流热固耦合强度及疲劳寿命分析
- Analysis of Fluid-Thermal-Solid Coupling Strength and Fatigue Life of a Certain Rocket Engine Turbine Rotor
- 西安交通大学学报 2022年56卷第8期页码：73-84
- 作者机构：
  
  1. 西安交通大学能源与动力工程学院,西安,710049
  2. 陕西省叶轮机械及动力装备工程实验室,西安,710049
  3. 中航科技集团航天六院西安航天动力研究所,西安,710100
- 作者简介：
  
  : 2021-12-06。作者简介: 黄朝晖(1998—),男,硕士生
  袁奇(通信作者),男,教授,博士生导师。基金项目: 国家自然科学基金资助项目(11872289,11902237)
- 基金信息：
  
  : 2021-12-06。作者简介: 黄朝晖(1998—),男,硕士生;袁奇(通信作者),男,教授,博士生导师。基金项目: 国家自然科学基金资助项目(11872289,11902237);“航空发动机与燃气轮机”国家科技重大专项资助(J2019-IV-0005-0072)
- DOI：10.7652/xjtuxb202208008
  中图分类号： TK14
- 网络首发：2022-08-10，
  
  纸质出版：2022
- 稿件说明：
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黄朝晖, 袁奇, 张弘斌, 等. 某型火箭发动机涡轮转子流热固耦合强度及疲劳寿命分析[J]. 西安交通大学学报, 2022,56(8):73-84.

HUANG Chaohui, YUAN Qi, ZHANG Hongbin, et al. Analysis of Fluid-Thermal-Solid Coupling Strength and Fatigue Life of a Certain Rocket Engine Turbine Rotor[J]. 2022, 56(8): 73-84.
黄朝晖, 袁奇, 张弘斌, 等. 某型火箭发动机涡轮转子流热固耦合强度及疲劳寿命分析[J]. 西安交通大学学报, 2022,56(8):73-84. DOI： 10.7652/xjtuxb202208008.

HUANG Chaohui, YUAN Qi, ZHANG Hongbin, et al. Analysis of Fluid-Thermal-Solid Coupling Strength and Fatigue Life of a Certain Rocket Engine Turbine Rotor[J]. 2022, 56(8): 73-84. DOI： 10.7652/xjtuxb202208008.

摘要

针对某国产型号液体火箭发动机涡轮泵在试车时转子叶片产生裂纹的问题

采用ANSYS流热固耦合方法

对叶片进行三维有限元仿真计算

提出了叶片顶部加围带的改型设计方案。对原始模型进行强度计算时考虑了启停工况各关键时间点的流动、传热、瞬态温度场、转速的影响

计算出叶片运行过程中的应力分布

确定产生裂纹的原因是启动过程中部分进气的气流冲击应力过高

且叶片型底前缘应力集中区域的高周疲劳安全系数过低。对改型设计方案进行仿真计算

结果表明:稳态综合加载时改型方案叶片最大等效应力比原模型下降了2 MPa

位置是叶片的型底背弧处; 启动0.35 s时改进模型叶片最大等效应力下降2.73%

停机过程中最大等效应力降幅为0.53%; 改型方案叶片高周疲劳安全系数均大于1.4的要求

叶片低周疲劳循环次数为22次

与原始模型相比降低了1次

仍满足要求。改型后发动机已安全飞行数次

证明改型方案符合设计要求。

Abstract

In view of the cracks in the rotor blades of the turbopump of a domestic model of liquid rocket engine during the test run

the ANSYS fluid-thermo-solid coupling method is used to carry out 3D finite element simulation calculation for the blades

and a modified design scheme of adding a shroud on the top of the blades is proposed. In the strength calculation of the original model

the influences of flow

heat transfer

transient temperature field and rotational speed at key time points under start-stop conditions are considered

and the stress distribution during blade operation is calculated. It is determined that the reason for the cracks is that the impact stress caused by some intake air is too high during the start-up process

and the high-cycle fatigue safety factor of the stress concentration area of the leading edge of the blade bottom is too low. The simulation calculation for the modified design scheme is carried out

and the results are as follows: At the back arc of the blade bottom

the maximum equivalent stress of the blades in the modified scheme is reduced by 2 MPa compared with the original model under steady comprehensive loading. After startup for 0.35 s

the maximum equivalent stress of the improved model blades decreased by 2.73%

and during the shutdown process

the maximum equivalent stress decreased by 0.53%; the high-cycle fatigue safety factor of the blades in the modified scheme is greater than the required value 1.4; the number of low-cycle fatigue cycles of the blade is 22

one less compared with the original model

but still meets the requirements. After the modification

the engine can support safe flying of the rocket several times

proving that the modification scheme meets the design requirements.

关键词

Keywords

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