智能弹药用新型电控固体推进剂的研究进展

针对当前固体火箭发动机无法重复启停、难以实现推力可控等需求,结合固体火箭和液体火箭的优点,国内外研究机构探究了一种具有重复点火与熄火、燃烧速度可调、输出能量可控的电控固体推进剂。系统介绍了新型电控固体推进剂领域的国内外研究现状,首先,总结了国内外各类电控固体推进剂(硝酸铵基、硝酸羟胺基、高氯酸锂基)的发展,以及在微推、炮弹点火器和固体发动机方面的应用情况;其次,概述了配方组分、电压、电流、外界环境温度和压强等因素对电控固体推进剂的燃速、点火延迟时间以及力学性能的影响;然后,分析了电控固体推进剂的启停机理,介绍了硝酸铵羟胺和高氯酸锂基电控固体推进剂的反应机理;最后,对电控固体推进剂的未来发展进行了展望,指出未来仍需重点研究的方向：加强电控固体推进剂制备工艺放大研究;设计新型电敏氧化剂,解决高压下无法可控燃烧的问题;开展大尺寸电控固体发动机试验研究;通过数值模拟进一步揭示电能与推进剂燃烧耦合效应。附参考文献62篇。     

环氧涂料涂装火箭发动机

欧洲“织女星”火箭的第一级发动机P80日前在法属圭亚那库鲁进行首次静态点火试验。其最显著的技术优势是采用了由长纤维缠绕的石墨环氧树脂所制成的发动机助推器涂层。据悉,此涂层技术广泛用于民用火箭小型发动机与弹道导弹,涂层材料比目前“阿里安”火箭使用的不锈钢材料质量大大减轻,可极大提高载荷能力。同时新设计的点火器,也使用炭纤维石墨环氧树脂涂层。据了解,P80发动机是迄今为止进行试验的最大一段式固体发动机。[摘自《中国化工报》]环氧涂料涂装火箭发动机     

金属/AP颗粒混合物燃烧特性的光谱分析

为了使Al/AP双组元粉末火箭发动机密度比冲最大化,将燃烧室特征长度由2.31 m增至 12.62 m进行了Al/AP粉末火箭发动机点火测试.采用光谱仪、CCD 相机、CO2 激光点火器等对 Al/AP 混合物在 1.0132 5 × 105 Pa的氮气环境中的点火延迟、燃烧时间、燃烧平稳性等燃烧性能进行了研究.测量了Al颗粒的表观堆积密度.作为一种替代燃料,对镁颗粒也进行了研究.结果表明,增加燃烧室特征长度至 12.62 m 时,可以得到最大燃烧室压强振荡幅度±2 .43%的平稳燃烧性能.含粒径 1μm 铝粉的 Al/AP 混合物其燃烧过程的光强远大于含粒径10μm铝粉的样品,并且其在波长 568 nm 发射光谱的光子数强度超过了光谱仪检测上限(65 000 数).而含粒径10μm铝粉样品燃烧过程的568 nm发射光谱信号出现间断且其全程强度低于 19 036 数.粒径 10μm 铝粉点火延迟时间为粒径1μm铝粉点火延迟时间的3.65 倍,燃烧时间为3.03 倍以上,最大RAlO却比 1μm铝粉少 14.3%,密度低21 .3%,说明粒度小的铝粉具有更好的燃烧性能,但是其堆积密度也更低.虽然Mg/AP的理论比冲为Al/AP的95.6%,但是其堆积密度比粒径1μm铝粉高8%,其点火延迟时间比粒径10μm铝粉短 90.3%.火焰照片也表明镁粉可在很大程度上减少凝相沉积.     

边界条件对尾部点火性能的影响

通过模拟点火空间、改变点火药量及喷管堵片厚度的技术措施，解决了某发动机尾部点火装置研制过程中出现的点火延迟等技术问题，试验研究表明，对尾部点火设计而言，单靠增加点火药量并不是解决发动机点火延迟问题的有效途径。通过改变点火边界条件，使得点火药量与喷管堵片设计厚度协调、匹配，在解决发动机点火延迟等技术问题中起着重要作用。     

火驱电点火及其与其他点火方式的对比分析

火烧油层作为提高稠油资源采收率的重要有效的方法,具有明显技术优势和潜力,具有油藏适应范围广、物源充足、采收率高、成本低的优势。而火烧油层开发成功的关键是油层点火。而利用井底电加热点火器点燃地层原油是一种最常见的点火方法,即电点火工艺,其核心是电加热器,经试验成功率达100%,获得了良好的增有效果,较其它点火方式更好。     

影响固体火箭发动机初始压强峰的因素分析

针对某型号固体火箭发动机初始压强峰峰值超差的现象,分别对固体推进剂的能量特性、初始燃烧表面、发动机所用的点火具、测试系统以及发动机的装配过程等可能引起初始压强峰峰值超差的因素,进行了理论分析和实验验证。结果表明,发动机内筒涂层过厚,使得装配完的发动机的初始通气参量变大,同时发动机装配过程不合理,产生从推进剂药柱上刮下来的药屑,这些均导致初始压强峰超标。     

铝镁贫氧推进剂的点火性能

为研究镁铝富燃料固体推进剂组分对点火性能的影响,采用改进的靶线法燃速测试系统对多种含镁铝富燃料固体推进剂在常压和加压下进行了通电金属丝点火性能的对比实验。被测试推进剂的镁铝合金含量为20%～40%,或者同时含镁铝合金及硼,氧化剂含量为30%～53%。实验表明,在固定外界输入热源的情况下,推进剂的点火性能主要与氧化剂含量和粒度有关;金属的含量和种类也有一定的影响;催化剂对点火延迟时间影响很小;压强对此种点火方式几乎无影响。该点火延迟测试方法简单易行,并具有一定的可靠度,适于配方调试。     

金属/AP颗粒混合物燃烧特性的光谱分析（英文）

为了使Al/AP双组元粉末火箭发动机密度比冲最大化,将燃烧室特征长度由2.31m增至12.62m进行了Al/AP粉末火箭发动机点火测试。采用光谱仪、CCD相机、CO2激光点火器等对Al/AP混合物在1.0132 5×105Pa的氮气环境中的点火延迟、燃烧时间、燃烧平稳性等燃烧性能进行了研究。测量了Al颗粒的表观堆积密度。作为一种替代燃料,对镁颗粒也进行了研究。结果表明,增加燃烧室特征长度至12.62m时,可以得到最大燃烧室压强振荡幅度±2.43%的平稳燃烧性能。含粒径1μm铝粉的Al/AP混合物其燃烧过程的光强远大于含粒径10μm铝粉的样品,并且其在波长568nm发射光谱的光子数强度超过了光谱仪检测上限(65 000数)。而含粒径10μm铝粉样品燃烧过程的568nm发射光谱信号出现间断且其全程强度低于19 036数。粒径10μm铝粉点火延迟时间为粒径1μm铝粉点火延迟时间的3.65倍,燃烧时间为3.03倍以上,最大RAlO却比1μm铝粉少14.3%,密度低21.3%,说明粒度小的铝粉具有更好的燃烧性能,但是其堆积密度也更低。虽然Mg/AP的理论比冲为Al/AP的95.6%,但是其堆积密度比粒径1μm铝粉高8%,其点火延迟时间比粒径10μm铝粉短90.3%。火焰照片也表明镁粉可在很大程度上减少凝相沉积。     

硼颗粒在固冲环境中点火过程影响因素的数值模拟

采用修正的Williams 硼颗粒燃烧模型对固体推进剂燃烧环境条件下的硼颗粒点火进行了数值模拟,计算了环境温度、气相氧化剂种类和分压、颗粒初始半径及氧化层厚度对硼颗粒点火的影响.结果表明,环境温度升高可以缩短硼颗粒点火延迟时间和点火时间;环境中氧分压过大会延长点火延迟时间;水蒸气分压越大,点火时间越短;硼颗粒半径增大会导致氧化层厚度增大进而延长点火时间和点火延迟时间.     

固体火箭发动机的热安全性研究

采用带源项的热传导方程,对固体火箭发动机在外界热源作用下的加热过程进行了数值模拟,分析了固体发动机内推进剂在外界热源作用下的燃烧特点,并确定了发动机产生热危险性的临界温度和起始燃烧时间。研究结果表明,在热传导方程中加入化学反应源项,可以有效地模拟发动机在外界热源作用下的加热过程;推进剂产生热危险性的临界温度为520～525K;在外界火焰作用下,发动机内的推进剂将点火燃烧,随着外界火焰温度的上升,推进剂起始燃烧的延迟时间减少。     

Numerical simulation of the effect of multiple factors on the ignition process of a solid rocket motor

With the continuous development of manned space technology, higher requirements have been proposed for solid rocket motors. The ignition process of solid rocket motors affects the reliability, stability and safety of their operation. The ignition powder, cover opening pressure and grain length-diameter ratio are the main factors affecting the ignition process. Therefore, the influence of different factors on the ignition process of solid rocket motors is studied with numerical simulations. Based on the finite volume method, the ignition process of a solid rocket motor is modelled and experimentally verified. Then, the pressure and temperature distribution characteristics during the ignition delay, flame propagation and gas filling times are analysed. Finally, the effects of different ignition powders, cover opening pressures and length-diameter ratios on the ignition process are compared and analysed. The results show that the model has high prediction accuracy. When the ignition powder is 6 g, the maximum combustion temperature of solid rocket motor increases from 2590 K to 2620 K between 0.1 ms and 0.44 ms. Between 0.44 ms and 3.18 ms, intermittent flame propagation and pressure oscillations occur. In the gas filling time, the flow field gradually stabilizes. Increasing the ignition powder mass is beneficial to the ignition process, but the disadvantages of pressure oscillations should be considered. Increasing the cover opening pressure enhances the ignition process, while increasing the length-diameter ratio increases the ignition pressure building time. The study results provide technical support for the structural design of solid rocket motors.     

IGNITION OF AMMONIUM PERCHLORATE

A simplified theory for the ignition of ammonium perchlorate is proposed, which is derived from a unified theory that also explains the low-pressure deflagration limit as well as the steady deflagration. The theory provides an approximate method of calculating the ignition delay and the minimum external he»t flux for a successful ignition, as functions of pressure and initial solid temperature. The ignition calculations show that there exists a pressure limit due to the weakness of the igniter strength, in addition to the low-pressure deflagration limit which is an inherent property of the solid independent of the igniter strength. The theory can be extended to other monopropellants for which exothermic reaction occurs only in the gas phase.     

IGNITION OF AMMONIUM PERCHLORATE

A simplified theory for the ignition of ammonium perchlorate is proposed, which is derived from a unified theory that also explains the low-pressure deflagration limit as well as the steady deflagration. The theory provides an approximate method of calculating the ignition delay and the minimum external he»t flux for a successful ignition, as functions of pressure and initial solid temperature. The ignition calculations show that there exists a pressure limit due to the weakness of the igniter strength, in addition to the low-pressure deflagration limit which is an inherent property of the solid independent of the igniter strength. The theory can be extended to other monopropellants for which exothermic reaction occurs only in the gas phase.     

Shock Tube Effect Inside a Pyrotechnic Igniter

An output closure is a critical component of a pyrotechnic igniter. It controls the heat transfer duration of initiation train, stress loading on the propellant grain, and the pressure drop during closure deployment. Normally the pressure profiles calculated by a quasi‐static interior ballistics code are adequate for igniter design evaluation. But following a case of premature closure deployment in which the propellant failed to ignite, the authors discovered that the design geometry mimicked that of a shock tube. The shock tube effect occurred whenever the high‐temperature gases of the initiator were rapidly discharged into a long conduit. The shock resultant from the initiator opened the closure prior to ignition of the ignition aid. In this paper, we report results from both quasi‐static computations for static pressure and time‐dependent simulations for dynamic pressure. Designers need to consider both static and dynamic pressure when devices have a sudden high‐pressure gas released into a conduit.     

Development of MTV Compositions as Igniter for HTPB/AP Based Composite Propellants

MTV compositions were prepared by keeping the magnesium/Teflon ratio constant and increasing the Viton content of the mixture up to 14% by an increment of 2% to investigate the effect of binder content on the heat of explosion, which is found to increase with the increasing Viton percentage as the magnesium content concomitantly goes down toward the stoichiometric value. In the second part of the study, fuel-rich MTV compositions were prepared by changing the magnesium content and keeping the Viton fraction constant at a specific value to investigate the effect of magnesium content on the heat of explosion and combustion characteristics. The observed general trend is that the heat of explosion of MTV compositions decreases as the magnesium content increases. All the MTV compositions were tested in a closed vessel to measure the maximum pressure achieved and the rate of reaching this pressure. The ignition performance of three selected MTV compositions was examined in 2.75 inch rocket motor by using the same charge of igniter and the same HTPB/AP composite propellant of the equal amount in each test. Two of them have excellent ignition performance and, therefore, can be used as igniter for the HTPB/AP based composite rocket propellants.     

An experimental study on the hypergolic ignition of hydrogen peroxide and ethanolamine

This paper presents the hypergolic ignition test results of a potential environmentally friendly liquid propellant consisting of hydrogen peroxide oxidizer (with a concentration of 85%) and ethanolamine fuel for use in rocket engines. Open cup drop tests were conducted to study the effect of amount of metal salt catalyst in fuel and the initial temperatures of fuel and oxidizer on the ignition delay time. To test the hypergolic ignition of bipropellant formulation in a real rocket engine environment, a pressure-fed liquid propellant rocket engine (LPRE) was designed and developed. During the tests it was found that the amount of catalyst and the initial temperature of the fuel had a significant effect on the ignition delay of hypergolic bipropellant. However, the oxidizer temperature seemed to have almost no affect on the ignition delay. There was also significant difference between the ignition delay times from open cup tests and those from rocket engine static firings.     

Three‐Dimensional Simulation of Pressure Fluctuation in a Granular Solid Propellant Chamber within an Ignition Stage

The effects of the conditions of the ignition system in the propellant chamber of a gun system using a granular solid propellant are numerically investigated with respect to ignition performance criteria such as the differential pressure generation between the breech and the projectile base. Simulations, in which the length of the primer and the igniter mass are varied, are performed using a solid/gas two‐phase fluid dynamics code for three‐dimensional calculation of gas flow and discrete solid propellant particles. This code simulates the igniter combustion in the primer, the movement of burning solid propellant grains, and the formation of pressure gradients inside the chamber in the ignition process. The differential pressures between the breech and the projectile base measured in experiments are well predicted by the simulations for various igniter conditions. In the process of igniting the solid propellant, the propellant grains are accelerated toward the projectile base by the igniter gas flows from the primer vents. Fixed‐particle simulation is also carried out in order to examine the effects of the movement of the solid propellant grains on the chamber pressure profile. The simulated results reveal that the movement of solid propellant grains causes differential pressure fluctuations, which depend on the discharge from the primer vents and the locations of these vents.