摘要
随着火箭发射任务轨道高度和载荷质量的提高,发动机推力和燃烧室直径增大,使燃烧室固有声学振型越发复杂。燃烧室收敛段、抗脉动隔板及其结构型式会显著影响燃烧室的声学特性,进而改变发动机的燃烧不稳定性裕度。为了研究燃烧室结构和隔板型式对声学特性的影响,建立了燃烧室声学有限元模型,并通过单喷嘴声学实验验证了仿真模型的准确性。研究了燃烧室收敛段和一周六径隔板对燃烧室声学特性的影响,重点分析了RD-170和F-1发动机不同隔板型式下燃烧室的声学特性,从声压分布的角度分析了其隔板设计的合理性。结果表明:添加收敛段后,燃烧室的1L和1T1L振型的频率分别提高了14%和17%。RD-170发动机的周向隔板位于2R振型速度波腹位置;F-1发动机所采用的两周八径13分区隔板不仅减小了2R振型速度波腹的半径,而且使切向振型的声压极值面积最小。双十字隔板使F-1发动机燃烧室中出现径向振型切向化的趋势。
Due to the increase of orbit height and load quality of launch mission,the engine’s thrust and di-ameter of combustion chamber increased,which makes its acoustic oscillation modes more and more complicated.Convergent section of combustion chamber,anti-pulsating baffle and its structural patterns could not only signifi-cantly affect the chamber’s acoustic characteristics,but also indirectly impact the combustion instability margin.In order to study the effects of structural parameters and baffle patterns on acoustic characteristics,a finite ele-ment simulation model of the combustion chamber is established. Firstly,the accuracy of the finite element modelis verified by a single nozzle acoustic experiment. On this basis,the influence rules of convergent section of com-bustion chamber and one hub and six radial baffle are analyzed. The design rationality of baffle patterns for RD-170 and F-1 engine is emphatically investigated in terms of sound pressure distribution. The results show the fre-quency of 1 L and 1 T1 L mode increases by 14% and 17%,respectively,when convergent section of combustionchamber is considered. For RD-170 engine,the hub baffle locates at the position of 2 R mode acoustic velocity anti-node. For F-1 engine,the 13-compartment baffle could not only reduce the radius of 2 R mode acoustic velocity an-ti-node,but also minimize the area of pressure amplitude in tangential modes. The trend of radial vibration modechanging to tangential vibration mode is observed in F-1 engine’s combustion chamber with double cross baffle.
作者
曹晨
谭永华
陈建华
李龙飞
CAO Chen;TAN Yong-hua;CHEN Jian-hua;LI Long-fei(Science and Technology on Liquid Rocket Engine Laboratory,Xi’an Aerospace Propulsion Institute,Xi’an 710100,China;Academy of Aerospace Propulsion Technology,Xi’an 710100,China)
出处
《推进技术》
EI
CAS
CSCD
北大核心
2021年第7期1581-1592,共12页
Journal of Propulsion Technology
关键词
大直径燃烧室
液氧煤油发动机
收敛段
声学特性
隔板型式
Large diameter combustion chamber
LOX/kerosene rocket engine
Convergent section
Acoustic characteristics
Baffle pattern
作者简介
曹晨,博士生,研究领域为液体火箭发动机技术。E-mail:18681840502@163.com;通讯作者:谭永华,博士,研究员,研究领域为液体火箭发动机总体技术。E-mail:tanyhcasc@163.com。
