矩形模型火箭发动机横向不稳定燃烧的数值模拟
2024,46(5):179-188
任永杰
航天工程大学 宇航科学与技术系, 北京 101416，spacedreamer@163.com,kkguo1003@126.com
郭康康
航天工程大学 宇航科学与技术系, 北京 101416，spacedreamer@163.com,kkguo1003@126.com
徐伯起
航天工程大学 宇航科学与技术系, 北京 101416
仝毅恒
航天工程大学 宇航科学与技术系, 北京 101416
聂万胜
航天工程大学 宇航科学与技术系, 北京 101416
航天工程大学 宇航科学与技术系, 北京 101416，spacedreamer@163.com,kkguo1003@126.com
郭康康
航天工程大学 宇航科学与技术系, 北京 101416，spacedreamer@163.com,kkguo1003@126.com
徐伯起
航天工程大学 宇航科学与技术系, 北京 101416
仝毅恒
航天工程大学 宇航科学与技术系, 北京 101416
聂万胜
航天工程大学 宇航科学与技术系, 北京 101416
摘要：
为研究火箭发动机横向不稳定燃烧特性,采用详细化学反应机理(GRI Mech 3.0)建表的小火焰生成流型,对模型火箭发动机中出现的横向不稳定燃烧进行数值模拟。通过与实验数据对比验证了模型的准确性；采用动态模态分解对压力场进行分析,研究了流场的动态特性；结合瑞利因子定量分析了不稳定燃烧的驱动特性。结果表明,数值模型能够有效捕捉横向不稳定燃烧,其主频与实验值相差不到1%；燃烧室横向压力振荡与喷嘴氧管纵向压力振荡相耦合,引起推进剂质量流量振荡；不稳定燃烧的驱动源主要位于燃烧室两侧,最边缘喷嘴对维持不稳定燃烧的贡献最大；推进剂与燃烧室侧壁面的相互作用极大增强了释热脉动,周期性释热为压力振荡提供能量,形成了不稳定燃烧极限环。
为研究火箭发动机横向不稳定燃烧特性,采用详细化学反应机理(GRI Mech 3.0)建表的小火焰生成流型,对模型火箭发动机中出现的横向不稳定燃烧进行数值模拟。通过与实验数据对比验证了模型的准确性；采用动态模态分解对压力场进行分析,研究了流场的动态特性；结合瑞利因子定量分析了不稳定燃烧的驱动特性。结果表明,数值模型能够有效捕捉横向不稳定燃烧,其主频与实验值相差不到1%；燃烧室横向压力振荡与喷嘴氧管纵向压力振荡相耦合,引起推进剂质量流量振荡；不稳定燃烧的驱动源主要位于燃烧室两侧,最边缘喷嘴对维持不稳定燃烧的贡献最大；推进剂与燃烧室侧壁面的相互作用极大增强了释热脉动,周期性释热为压力振荡提供能量,形成了不稳定燃烧极限环。
基金项目：
国家自然科学基金资助项目(51876219,12002386)
国家自然科学基金资助项目(51876219,12002386)
Numerical simulation of transverse combustion instability in a rectangle model rocket combustor
REN Yongjie
Department of Aerospace Science and Technology, Space Engineering University, Beijing 101416, China,spacedreamer@163.com,kkguo1003@126.com
GUO Kangkang
Department of Aerospace Science and Technology, Space Engineering University, Beijing 101416, China,spacedreamer@163.com,kkguo1003@126.com
XU Boqi
Department of Aerospace Science and Technology, Space Engineering University, Beijing 101416, China
TONG Yiheng
Department of Aerospace Science and Technology, Space Engineering University, Beijing 101416, China
NIE Wansheng
Department of Aerospace Science and Technology, Space Engineering University, Beijing 101416, China
Department of Aerospace Science and Technology, Space Engineering University, Beijing 101416, China,spacedreamer@163.com,kkguo1003@126.com
GUO Kangkang
Department of Aerospace Science and Technology, Space Engineering University, Beijing 101416, China,spacedreamer@163.com,kkguo1003@126.com
XU Boqi
Department of Aerospace Science and Technology, Space Engineering University, Beijing 101416, China
TONG Yiheng
Department of Aerospace Science and Technology, Space Engineering University, Beijing 101416, China
NIE Wansheng
Department of Aerospace Science and Technology, Space Engineering University, Beijing 101416, China
Abstract：
To study the transverse combustion instability characteristics of the rocket combustor, numerical simulations of transverse combustion instability in a model rocket combustor were conducted based on the detailed chemical reaction mechanism (GRI Mech 3.0) and the flamelet-generated manifolds method. Accuracy of the numerical model was verified by comparing it with the experimental data. Pressure field was analyzed by the dynamic mode decomposition method, and the dynamic characteristics of the flow fields were investigated. Driving characteristics of combustion instability were quantitatively estimated by Rayleigh index. Result shows that the transverse combustion instability that occurred in the experiment can be effectively captured by the numerical model. Dominant frequency identified by the numerical study differ from the experimental value by less than 1%. Transverse pressure oscillations in the combustion chamber are coupled with that the longitudinal mode in the oxidizer post, leading to the pulsated propellant mass flow rate. Driving regions of combustion instability are mainly located on both sides of the combustion chamber, and the most marginal injectors played a critical role in keeping combustion instability. Heat release pulsations which periodically provide the energy source for the pressure oscillations are highly enhanced by the interactions between the propellant and the sidewall of the combustion chamber. And the combustion instability limit-cycle is formed.
To study the transverse combustion instability characteristics of the rocket combustor, numerical simulations of transverse combustion instability in a model rocket combustor were conducted based on the detailed chemical reaction mechanism (GRI Mech 3.0) and the flamelet-generated manifolds method. Accuracy of the numerical model was verified by comparing it with the experimental data. Pressure field was analyzed by the dynamic mode decomposition method, and the dynamic characteristics of the flow fields were investigated. Driving characteristics of combustion instability were quantitatively estimated by Rayleigh index. Result shows that the transverse combustion instability that occurred in the experiment can be effectively captured by the numerical model. Dominant frequency identified by the numerical study differ from the experimental value by less than 1%. Transverse pressure oscillations in the combustion chamber are coupled with that the longitudinal mode in the oxidizer post, leading to the pulsated propellant mass flow rate. Driving regions of combustion instability are mainly located on both sides of the combustion chamber, and the most marginal injectors played a critical role in keeping combustion instability. Heat release pulsations which periodically provide the energy source for the pressure oscillations are highly enhanced by the interactions between the propellant and the sidewall of the combustion chamber. And the combustion instability limit-cycle is formed.
收稿日期：
2022-05-10
2022-05-10
