作　　者： ; ; ; ;

机构地区： 西北工业大学动力与能源学院

出　　处： 《西北工业大学学报》 2005年第5期549-552,共4页

摘　　要： 阐述了脉冲爆震火箭发动机的工作原理及其特点。设计并建立了整套脉冲爆震火箭发动机实验模型。以液体燃料航空煤油为燃料、氧气为氧化剂、压缩氮气为隔离气体,在内径为25 mm,长度为0.8 m的爆震管内产生了充分发展的爆震波。测量了不同工作频率下的爆震波压力,并对其进行了分析。实验结果表明,在设计的实验模型中,采用低的点火能量(50 m J)能够在较短的距离内产生充分发展的爆震波。 Our research has been done under a research project supported by NNSFC （National Natural Science Foundation of China）. In the research plan we submitted for the project, we explained the possible ways and means of making using liquid fuel in PDRE model practicable. Unlike conventional steady-state rocket engines, which use constant pressure combustion, PDREs utilize the high-energy release rate and thermodynamic characteristics of detonation waves to produce thrust. This enables PDREs to operate at higher thermodynamic efficiencies. Furthermore, since the reactants are injected into PDREs at relatively low pressures, the need for massive turbo-machinery, as used in conventional steady-state liquid rocket engines, is eliminated. The utilization and performance of liquid hydrocarbon fuels used in a PDRE are being investigated in this paper, since these tuels are attractive for volume-limited aerospace systems. In our experiments, the PDRE test model utilized kerosene as the fuel, oxygen as oxidizer, and nitrogen as purge gas. Solenoid valves were employed to control intermittent supplies of kerosene, oxygen, and purge gas. The spark plug igniter used in these experiments had initiation energy of only around 50 mJ. The solenoid valves and spark plug were controlled by the control and ignition system. PDRE test model was 25mm in inner diameter by 800 mm long. DDT （deflagration to detonation transition） enhancement device Shchelkin spiral was used in the test model. Three piezoelectric pressure transducers placed on the detonation tube at different locations from the thrust wall were used to measure pressure trace. The proofof-principle experiments of liquid kerosene-oxygen were achieved successfully within an aceeptable length （less than 0. 7 m from the thrust wall） with only 50 mJ ignition energy. The obtained pressure ratio of detonation wave was close to that of C-J （Chapman-Jouguet） detonation. Under various frequencies, the detonation pressures were measured and analyzed. The experimental results

关 键 词： 脉冲爆震火箭发动机 爆震 模型 实验研究

领　　域： [航空宇航科学与技术] [航空宇航科学技术]