离子液体电喷雾推进与单组元化学推进组合形成的双模式推进系统兼具电推进模式的高比冲和化学推进模式的大推力，而且两种模式共用同一种推进剂，使得卫星具备在轨任务调整的能力，具有良好的发展前景。双模式推进系统使用的推进剂由咪唑类离子液体和氧化性离子液体以及其他助剂混合而成，其组分分子量大，含碳量高，应用于化学推进时存在点火困难、易积碳导致催化床活性下降等问题。本文聚焦于一种由硝酸羟胺（HAN）、1-乙基-3-甲基咪唑硫酸乙酯盐（[Emim][EtSO4]）、水（H2O）和其他微量助剂组成的新型高含能离子液体推进剂，开展了其点火技术的初步研究。首先，采用热重-傅里叶变换红外光谱（TG-FTIR）实验对比研究原始推进剂以及额外添加一定量H2O的推进剂样品的分解反应特性和产物。结果表明，推进剂的分解与反应可分为四个阶段，依次为H2O蒸发、HAN分解、[Emim][EtSO4]快速氧化还原反应和残余物质的缓慢反应，推进剂的反应速率受分解反应的控制。推进剂分解的主要气相产物为N2O和CO2。随后，通过催化床反应器实验，发现了催化床长度对于推进剂点火的关键性影响，推进剂流量为2 mL/min时，催化床长度增长约13 mm，可以使点火预热温度降低70℃以上。进而，对推进剂在催化床内的点火过程进行建模分析。提出了描述推进剂在催化床内流动与反应过程的简化偏微分方程组和定解条件，使用梯度下降法确定了方程组中各物性参数，并使用有限差分方法对方程组进行了离散求解。基于此模型开发了预测催化床尺寸结构的软件，软件可以考虑催化剂活性下降带来的影响，依据不同工况给出准确的点火位置。软件还支持用户通过实测温度变化数据和催化剂可持续点火的最长时间数据对物性参数和催化剂失活速率常数进行校准。最后，通过直流电、直流放电等离子体和微波等离子体点火装置对离子液体推进剂直接点火性能进行探究。结果表明，由于离子液体推进剂具有较好的导电性，使用直流电点火方式可以实现离子液体推进剂的成功点火，但实现大流量推进剂稳定点火较困难。直流放电等离子体点火方式容易因电源的自脉冲放电效应造成推进剂点火燃烧不连续，因而对等离子体电源有较高要求。微波等离子体点火方式可在输入功率约100 W时对0.5~3 mL/min的推进剂实现成功点火，且停止等离子体输入后，推进剂可以自维持燃烧，是一种有良好发展前景的直接点火技术。
The dual-mode propulsion system combines the high specific impulse of the electric propulsion mode and the large thrust of the chemical propulsion mode. It shows a great application prospect since the dual modes can share the same propellant that enables the satellite to have flexible in-situ mission adjustment capabilities. The propellant used here is a mixture of imidazole-based ionic liquid and oxidizing ionic liquid as well as other additives, whose components have large molecular weight and high carbon content. When the chemical propulsion mode is applied, it results in a series of problems such as hard ignition and serious carbon deposits leading to the deactivation of catalytic bed. This paper aims to conduct a preliminary study on the ignition technology of a high-energy ionic liquid propellant consisting of hydroxylamine nitrate (HAN), 1-ethyl-3-methylimidazolium ethylsulfate ([Emim][EtSO4]) and H2O and small amounts of other additives.First, the decomposition characteristics and the products of propellant samples with or without additional amounts of H2O were studied comparatively with thermogravimetric-fourier transform infrared spectroscopy (TG-FTIR) experiments. The results indicates that the complete reactions of propellant can be divided into four stages, i.e. H2O evaporation, HAN decomposition, [Emim][EtSO4] rapid reaction and slow reaction of residuals. The global reaction rate of the propellant is dominated by these decomposition reactions. The main gas-phase products of propellant are N2O and CO2. Then, through catalytic bed reactor experiments, the critical effect of catalytic bed length on propellant ignition was found. At a propellant flow rate of 2 mL/min, increasing the catalytic bed length by about 13 mm could reduce the ignition preheat temperature by more than 70°C.Second, the ignition process of propellant in the catalytic bed was modeled and analyzed. A set of simplified partial differential equations describing the flow and reaction process of propellant in the catalytic bed, as well as the initial and boundary conditions were proposed. The physical parameters in the equations were determined using gradient descent method, and then the equations were discretized and solved using finite difference method. Based on this model, a software was developed to predict the catalytic bed size structure, which could give the exact ignition position according to different working conditions considering the effect of catalyst deactivation. The software also supports the user to calibrate the physical parameters and the rate constant of catalyst deactivation using measured temperature variation data and the maximum time that the catalyst can work continuously.Finally, the ionic liquid propellant ignition characteristics were explored by a DC ignition device, a DC discharge plasma ignition device and a single electrode microwave plasma ignition device. The results shows that successful ignition can be achieved using the DC ignition device since the ionic liquid propellant has good electrical conductivity. However, it is difficult to achieve stable ignition and combustion for the propellant of a large flow rate. The DC discharge plasma ignition device is prone to discontinuous propellant ignition and combustion due to the self-pulse discharge effect, and thus has high requirements for the discharge power supply. The microwave plasma ignition device can achieve successful ignition of the propellant at a flow rate of 0.5~3 mL/min at an input power of about 100 W, and the propellant combustion can be self-sustained after stopping the plasma input. It is demonstrated to be an efficient direct ignition technology with a good development prospect.
