检索结果-南通市图书馆

Numerical and experimental research on axial injection end-burning hybrid rocket motors with polyethylene fuel

在线全文

学校读者我要写书评

暂无评论

Chinese Journal of Aeronautics 2024年第8期37卷 91-105页

作者： Xiangyu MENG Hui TIAN Lingfei HE Jingfei GAO Xiaoting NIU guobiao cai School of Astronautics Beihang UniversityBeijing 100191China Key Laboratory of Spacecraft Design Optimization&Dynamic Simulation Technologies Ministry of EducationBeijing 100816China

This study investigates the end-burning hybrid rocket motors with polyethylene fuel by the numerical simulation and *** on computational fluid dynamics,a numerical model is *** model is validated by two firing tests in this hybrid rocket motor,which uses oxygen and polyethylene as *** results show that the numerical and experimental data are in good agreement,and the error of the chamber pressure is less than 2.63%.Based on the simulation mode,the blowoff limit of the end-burning hybrid rocket motors is *** the nozzle throat diameter and the inner diameter of grain are large,it is more difficult for the hybrid rocket motor to achieve end-burning mode,i.e.,the flame spreading is prevented in the narrow *** main reason is that when the nozzle throat and the grain port are large,chamber pressure and oxidizer flow velocity are ***,the friction velocity considering the pressure and flow velocity is *** critical friction velocity is about 4.054–4.890 m/s in the hybrid rocket *** the friction velocity exceeds the critical friction velocity,the combustion mode in hybrid rocket motors changes from the flame spreading mode to the end-burning ***,the regression rate formula is obtained by fitting,which shows that the regression rate has a good correlation with combustion chamber *** critical friction velocity and regression rate formula can provide an important reference for end-burning hybrid rocket motors.

关键词： Hybrid rocket motor End burning Blowoff limit Flame propagation Regression rate

Geometric effects of fuel regression rate in hybrid rocket motors

在线全文

学校读者我要写书评

暂无评论

Science China(Technological Sciences) 2016年第5期59卷 807-813页

作者： cai guobiao ZHANG YuanJun WANG PengFei HUI Tian ZHAO Sheng YU NanJia School of Astronautics Beihang University Beijing 100191 China Beijing Institute of Aerospace Systems Engineering Beijing 100076 China

The geometric configuration of the solid fuel is a key parameter affecting the fuel regression rate in hybrid rocket motors. In this paper, a semi-empirical regression rate model is developed to investigate the geometric effect on the fuel regression rate by incorporating the hydraulic diameter into the classical model. The semi-empirical model indicates that the fuel regression rate decreases with increasing hydraulic diameter and is proportional to dh?0.2 when convective heat transfer is dominant. Then a numerical model considering turbulence, combustion, solid fuel pyrolysis, and a solid–gas coupling model is established to further investigate the geometric effect. Eight motors with different solid fuel grains are simulated, and four methods of scaling the regression rate between different solid fuel grains are compared. The results indicate that the solid fuel regression rates are approximate the same when the hydraulic diameters are equal. The numerical results verify the accuracy of the semi-empirical model.

关键词： geometric effect regression rate hydraulic diameter hybrid rocket motor

Three-dimensional particle simulation of ion thruster plume flows with EX-PWS

在线全文

学校读者我要写书评

暂无评论

Plasma Science and Technology 2018年第10期20卷 118-126页

作者： Hongru ZHENG guobiao cai Hongyue WANG Lihui LIU Bijiao HE School of Astronautics Beihang UniversityBeijing 100191People＇s Republic of China Beijing Institute of Aerospace Testing Technology Beijing 100074People＇s Republic of China

Ion thruster plumes from a multi-thruster array of different working configurations are simulated by a hybrid fluid-particle software. The particle in cell method is employed to model the transports of ions. The direct simulation Monte Carlo method is used to model momentum and charge exchange （CEX） collisions. The software is based on unstructured grids which make it easy to handle with complex geometry. The results of chamber simulation are compared with experimental data in ion current density and number density, which show good agreements. The maximum difference of current density along the thruster centerline is less than 9.30%. The interaction effects of plumes when multiple thrusters are operating in vacuum are predicted. Distributions of single charged xenon ions are significantly different in the near-field plume flow, however, merge into one in the far downstream region. Moreover, the interaction effect on the spatial distribution of CEX xenon ions is displayed as well.

关键词： electric propulsion multi-thruster array PIC-DSMC method ion thruster plume

The effect of anode axial position on the performance of a miniaturized cylindrical Hall thruster with a cusp-type magnetic field

在线全文

学校读者我要写书评

暂无评论

Plasma Science and Technology 2022年第7期24卷 15-22页

作者： Yuanyuan GAO Weizong WANG Yifei LI Shuwen XUE guobiao cai Advanced Space Propulsion and Energy Laboratory(ASPEL) School of AstronauticsBeihang UniversityBeijing 100191People's Republic of China Aircraft and Propulsion Laboratory Ningbo Institute of TechnologyBeihang UniversityNingbo 315832People's Republic of China

A 200 W cylindrical Hall thruster with a cusp-type magnetic field was proposed, manifesting convergent plume and high specific impulse. In this paper, a series of ring-shaped anodes are designed and the influence of anode axial position on the performance of CHT with a cusp-type magnetic field is studied. The experimental results indicate that the thruster keeps stable operation at the condition of 140–270 W discharge power. When the anode moves axially towards the upstream cusp field, the thrust enhances from 6.5 mN to 7.6 mN and specific impulse enhances from 1658 s to 1939 s significantly. These improvements of thruster performance should be attributed to the enhancement of current utilization, propellant utilization and acceleration efficiency. According to the analyses on the discharge characteristics, it is revealed that as the anode moves upstream, the electron transport path could be extended, the magnetic field in this extended path could impede electron cross-field transport and facilitate the ionization intensity, yielding to the enhancement of current utilization and propellant utilization ***, along with this enhancement of upstream ionization at the given anode flow rate, the main ionization region is thought to move upstream and then separate more apparently from the acceleration region, which has been demonstrated by the narrowing of ion energy distribution function shape. This change in acceleration region could decrease the ion energy loss and enhance acceleration efficiency. This work is beneficial for optimizing the electrode structure of thruster and recognize the ionization and acceleration process under the cusp magnetic field.

关键词： performance ionization characteristics acceleration characteristics cylindrical Hall thruster anode axial position

在线全文

学校读者我要写书评

暂无评论

不同材料发汗冷却结构下氢冷却性能

航空动力学报 2022年第4期37卷 781-790页

作者：孙纪国蔡国飙北京航空航天大学宇航学院北京100191

为了实现航空航天等领域高温大热流燃烧装置的有效冷却,研究了不同材料和工艺制成的发汗冷却结构在高温高热流密度下,氢的发汗冷却性能。模拟高压推力室的结构特点和高热流设计发汗冷却试验件,用电弧加热主流空气模拟高温燃气、以氢气... 详细信息

为了实现航空航天等领域高温大热流燃烧装置的有效冷却,研究了不同材料和工艺制成的发汗冷却结构在高温高热流密度下,氢的发汗冷却性能。模拟高压推力室的结构特点和高热流设计发汗冷却试验件,用电弧加热主流空气模拟高温燃气、以氢气为发汗冷却剂对多孔陶瓷、烧结多孔不锈钢和多孔层板材料进行了33次172 s热试验研究。试验的材料设计孔隙率为10%~40%,燃烧室压力为2.7~8.4 MPa,主流燃气温度约为3600 K,主流空气流量为220~1490 g/s,冷却氢气流量为9.6~57 g/s,注入率为0.005~0.029。试验结果表明:当冷却剂氢注入率为1%时,主流与多孔陶瓷材料壁面和粉末冶金多孔结构壁面之间的换热分别减少了30%和70%以上;当注入率为3%时,主流与光刻多孔层间结构壁面之间的换热也能降低60%。证明氢发汗冷却可以有效减小壁面与燃气之间的对流热流。最后还总结得出了常温氢气对高压大热流环境进行发汗冷却的性能关联式。

关键词：发汗冷却氢气多孔材料热防护粉末冶金陶瓷层板

同方期刊数据库

在线全文

学校读者我要写书评

暂无评论

基于压敏漆的多羽流气动力效应试验研究

北京航空航天大学学报 2020年第6期46卷 1080-1088页

作者：吴靖蔡国飙福州大学机械工程及自动化学院福州350116 北京航空航天大学宇航学院北京100083

多股羽流相互作用会形成复杂流场形态的干扰羽流,为了对干扰羽流开展气动力效应试验研究,采用压敏漆(PSP)表面压力光学测量技术,对以常温空气为工质的单喷嘴羽流和双喷嘴干扰羽流撞击平板模型的气动力进行了高分辨率的全场测量,分析了... 详细信息

多股羽流相互作用会形成复杂流场形态的干扰羽流,为了对干扰羽流开展气动力效应试验研究,采用压敏漆(PSP)表面压力光学测量技术,对以常温空气为工质的单喷嘴羽流和双喷嘴干扰羽流撞击平板模型的气动力进行了高分辨率的全场测量,分析了喷嘴和平板间不同入射距离和入射角对羽流气动力作用强度和范围的影响。试验结果表明,以高透氧聚合物为基层的压敏漆在羽流气动力测量中具有快速响应的特性,能分辨Pa级的微小压力变化。相对传统离散测压孔方法,压敏漆温度敏感度低,能捕捉大梯度的压力变化,准确描述羽流气动力效应。通过对比,发现干扰羽流对气动力具有明显增强作用,且增强作用随着入射距离的降低而减弱,干扰羽流的气动力不能直接用单股羽流的气动力进行线性叠加;羽流相互作用增强羽流返流,在航天器设计中需要考虑多股羽流相互作用后的羽流防护问题。

关键词：真空羽流干扰羽流气动力效应压敏漆(PSP) 压力增强

在线全文

学校读者我要写书评

暂无评论

旋转对固体火箭发动机的影响

推进技术 1999年第1期20卷 11-15页

作者：蔡国飙田辉北京航空航天大学宇航学院

对旋转对固体火箭发动机的影响进行了初步探讨。给出了旋转对推进剂应力应变的影响、产生的热力学方面的影响及旋转引起的气体动力学的影响等的分析计算模型。并在内弹道计算的基础上，对上述影响引起的发动机性能的变化进行了分析比较... 详细信息

对旋转对固体火箭发动机的影响进行了初步探讨。给出了旋转对推进剂应力应变的影响、产生的热力学方面的影响及旋转引起的气体动力学的影响等的分析计算模型。并在内弹道计算的基础上，对上述影响引起的发动机性能的变化进行了分析比较，得出了有意义的结论。

关键词：旋转固体推进剂火箭发动机内弹道计算

二维H_2/O_2混合和流动的直接数值模拟

在线全文

学校读者我要写书评

暂无评论

北京航空航天大学学报 1999年第6期25卷 680-683页

作者：蔡国飙庄逢甘北京航空航天大学宇航学院北京航空航天大学飞行器设计与应用力学系

用高阶精度的紧致差分方法，对低速的随时间发展的二维可压缩Ｈ２／Ｏ２混合层流动编制了直接数值模拟的计算机程序，对雷诺数Ｒｅ为４００、来流马赫数Ｍａ分别为０．４（上部的Ｈ２）和１．５７（下部的Ｏ２）的Ｈ２／Ｏ２... 详细信息

用高阶精度的紧致差分方法，对低速的随时间发展的二维可压缩Ｈ２／Ｏ２混合层流动编制了直接数值模拟的计算机程序，对雷诺数Ｒｅ为４００、来流马赫数Ｍａ分别为０．４（上部的Ｈ２）和１．５７（下部的Ｏ２）的Ｈ２／Ｏ２亚跨超声速混合层流动进行了计算，得到了相应的计算结果，并对结果进行了分析和研究．

关键词：混合层数值模拟流动 N-S方程氢/氧

在线全文

学校读者我要写书评

暂无评论

应用于DSMC方法的直角网格技术

北京航空航天大学学报 2010年第6期36卷 694-699页

作者：侯凤龙蔡国飙北京航空航天大学宇航学院北京100191

研究了应用于直接模拟蒙特卡洛DSMC(Direct Simulation Monte Carlo)方法的直角网格技术.通过对边界三角形表面微元进行统计,建立直角网格缺陷单元与表面微元之间的映射关系,并进一步将缺陷单元区分为一级缺陷单元以及二级缺陷单元,同... 详细信息

研究了应用于直接模拟蒙特卡洛DSMC(Direct Simulation Monte Carlo)方法的直角网格技术.通过对边界三角形表面微元进行统计,建立直角网格缺陷单元与表面微元之间的映射关系,并进一步将缺陷单元区分为一级缺陷单元以及二级缺陷单元,同时结合可变时间步长技术,从根本上解决了在DSMC方法中利用直角网格技术贴体光滑地匹配复杂几何型面的问题.给出了相应的自适应网格调整方法.基于以上策略,编制了通用DSMC计算程序RGD-Tool(Rarefied Gas Dynamic Tool).利用RGD-Tool程序对典型的圆球绕流以及无限薄斜挡板绕流两个算例进行了验证计算,结果表明所讨论直角网格技术是有效的.

关键词：直接模拟蒙特卡罗方法直角网格边界映射