Magnetoplasma rocket engine has a broad application prospect in the deep space exploration, manned space flight and other space missions. The ion energy is converted into the directed velocity in the magnetic nozzle of the engine. The investigation into the detachment process of the plasma with the magnetic field is of great significance for improving the engine propulsion efficiency. However, there are roughly five kinds of physical mechanisms which can all contribute to the detachment process and make the detachment in the magnetic nozzle quite complicated. Furthermore, the ion temperature is much higher than the electron temperature in the magnetic nozzle of the magnetoplasma rocket engine due to the heating effect of the ion cyclotron resonance stage. As a result, previous numerical model which were based on the assumption of cold ions are unapplicable for the simulation of the engine. In this work, a fluid simulation model is developed which is used for simulating the magnetic nozzle in the magnetoplasma rocket engine. The model includes the electron and the ion of single charge. For the characteristics of the magnetoplasma rocket engine, the ion energy equation is added into the governing equations. In order to analyze the effect of the inertial detachment, the static electric field due to the charge separation is also included. The simulations are performed under the conditions of different inlet ion temperatures and background magnetic fields. The results show that the ion axial velocity gradually increases in the magnetic nozzle and the ion stream lines detach from the magnetic field lines gradually. The loss of adiabaticity is the dominant mechanism in the detachment process. The ion axial velocity increases with the inlet ion temperature rising, and the ion streamlines detach earlier from the magnetic field lines. The resistive diffusion is unaffected by the inlet ion temperature while the detachment interfaces of other three mechanisms all move toward the upstream. With the increase of the background magnetic field, ion axial velocity decreases and the angle included between the streamline and the axis becomes smaller. The loss of adiabaticity is still the dominant physical mechanism when the magnetic field is changed.
为发展RBCC动力系统,同时进一步探索推进性能对入轨有效载荷的影响规律,对以DRB-CC为动力的两级入轨飞行器运载特性进行了研究.在给定飞行器构型和飞行剖面基础上,开展了该飞行器180 km近地轨道两级入轨设计.结果表明:以DRBCC飞行器作为第一级,配合独立火箭动力的第二级,150 t级飞行器180 km近地轨道的有效载荷为4.773 t;DRBCC的推力和比冲与飞行器飞行状态密切相关,DRBCC在2.5 Ma以下时一直工作在混合模态,而在2.5 Ma以上直接转入亚燃冲压模态;在亚燃和超燃冲压模态,DRBCC的比冲随马赫数变化较为平缓,而推力出现了波动,且在亚燃冲压模态波动较大;两级入轨过程中,DRBCC混合模态主要使飞行器完成爬高,亚燃冲压模态同时用来完成爬高和增速,超燃冲压模态主要用来增速.
在火箭发动机涡轮泵高速轴承试验中,轴承的可靠性对保证试验的成功至关重要.针对涡轮泵轴承故障特征难以从原始信号中提取的问题,基于EMD-Hilbert包络解调分析方法,对涡轮泵高速轴承故障特征进行识别.采用EMD方法对原始信号进行自适应分解,获得若干个IMF分量;基于相关性指标最大原则筛选IMF分量进行信号重构;对重构信号进行Hilbert包络解调分析,提取出故障轴承的特征.以某型号涡轮泵低温高速轴承试验的真实故障数据验证本方法的有效性,数据记录了试验装置在阶梯式升速全过程中保持架故障的振动加速度信号.分析结果表明,基于EMD-Hilbert包络解调分析方法能够提高信噪比,最大程度保留保持架故障信息的周期性冲击成分,并能有效提取保持架故障频率、故障倍频及各种调制频率成分,实现对涡轮泵高速轴承故障的有效识别.为深度解析轴承保持架故障情况,结合系统11种运行状态,提出了一种轴承渐进劣化全过程的解析方法,确定出轴承故障早期人为干预的具体时刻.
为了研究RBCC发动机火箭模态及火箭冲压模态的工作特性,基于发动机地面自由射流试验结果,利用一维气动理论构建了发动机火箭及火箭冲压模态的性能分析模型,对发动机6 Ma来流条件的试验数据进行了处理,获得了发动机轴向的沿程气流参数,分析了发动机的热力循环、工作效率、有效能分布以及部件和排气中的有效能分配比例.结果表明:一维计算得到的推力与试验结果误差在5％以内;火箭冲压模态下火箭燃气的引入可以有效提升发动机的热循环效率(约提升20％),火箭燃气的引入对有效能产生率和有效能的分配比例影响不大,火箭冲压和冲压模态的有效能产生率分别为0.45和0.48;火箭模态推力增益产生的主要原因是火箭燃气的能量添加至冲压流道中,形成了有效的热力循环,产生了机械能增量,最终表现出推力增益,约为29％.
硝酸羟胺(HAN)发动机如果采用传统肼类均匀分配喷注器,那么在靠近喷注器一段距离的催化床内,中心温度低于边缘温度,中心位置反应物转化率低于边缘位置,并且中心位置反应物质量分数高于边缘位置,该不均匀性将影响发动机工作寿命.针对HAN发动机喷注器采用传统设计方案容易使得催化床中心过载的问题,提出了一种"外密内疏"流量分配喷注器设计方法,并给出了该方法的设计理论.开展了基于传统喷注器结构和新型"外密内疏"喷注器结构的60 N HAN发动机热试车试验,对比结果表明,传统均匀分配喷注器的60 N HAN发动机工作680 s后发动机失效,而采用新型特种流量分配方案喷注器的发动机可较稳定地完成1 200 s长稳态工作.提出的新型喷注器设计方法为长寿命HAN发动机工程应用提供了参考.
