Experimental study of Al agglomeration on solid propellant burning surface and condensed combustion products

Aluminum(Al) particles are commonly added to energetic materials including propellants,explosives and pyrotechnics to increase the overall energy density of the composite,but aluminum agglomeration on the combustion surface may lower the combustion efficiency of propellants,resulting in a loss in twophase flow.Therefore,it is necessary to understand the agglomeration mechanism of aluminum particles on the combustion surface.In this paper,a high-pressure sealed combustion chamber is constructed,and high-speed camera is used to capture the whole process of aluminum accumulation,aggregation and agglomeration on the combustion surface,and the secondary agglomeration process near the combustion surface.The microscopic morphology and chemical composition of the condensed combustion products(CCPs) are then studied by using scanning electron microscopy coupled with energy dispersive(SEM-EDS) method.Results show that there are three main types of condensed combustion products:small smoke oxide particles oxidized by aluminum vapor,usually less than 1 μm;typical agglomerates formed by the combustion of aluminum agglomerates;carbonized agglomerates that are widely distributed,usually formed by irregular movements of aluminum agglomerates.The particle size of condensed combustion products is measured by laser particle size meter.As the pressure increases from 0.5 MPa to 1.0 MPa in nitrogen,the mass average particle size of aluminum agglomerates decreases by 49.7%.As the ambient gas is changed from 0.5 MPa nitrogen to 0.5 MPa air,the mass average particle size of aluminum agglomerates decreases by 67.3%.Results show that as the ambient pressure increases,the higher oxygen content can improve combustion efficiency and reduce the average agglomeration size of aluminum particles.

Recent advances in catalytic combustion of AP-based composite solid propellants

Composite solid propellants(CSPs) have widely been used as main energy source for propelling the rockets in both space and military applications. Internal ballistic parameters of rockets like characteristic exhaust velocity, specific impulse, thrust, burning rate etc., are measured to assess and control the performance of rocket motors. The burn rate of solid propellants has been considered as most vital parameter for design of solid rocket motors to meet specific mission requirements. The burning rate of solid propellants can be tailored by using different constituents, extent of oxidizer loading and its particle size and more commonly by incorporating suitable combustion catalysts. Various metal oxides(MOs),complexes, metal powders and metal alloys have shown positive catalytic behaviour during the combustion of CSPs. These are usually solid-state catalysts that play multiple roles in combustion of CSPs such as reduction in activation energy, enhancement of rate of reaction, modification of sequences in reaction-phase, influence on condensed-phase combustion and participation in combustion process in gas-phase reactions. The application of nanoscale catalysts in CSPs has increased considerably in recent past due to their superior catalytic properties as compared to their bulk-sized counterparts. A large surface-to-volume ratio and quantum size effect of nanocatalysts are considered to be plausible reasons for improving the combustion characteristics of propellants. Several efforts have been made to produce nanoscale combustion catalysts for advanced propellant formulations to improve their energetics. The work done so far is largely scattered. In this review, an effort has been made to introduce various combustion catalysts having at least a metallic entity. Recent developments of nanoscale combustion catalysts with their specific merits are discussed. The combustion chemistry of a typical CSP is briefly discussed for providing a better understanding on role of combustion catalysts in burning rate enhancement. Available information on different types of combustion nanocatalysts is also presented with critical comments.

Controllable combustion behaviors of the laser-controlled solid propellant

Microsatellites have been widely applied in the fields of communication,remote sensing,navigation and science exploration due to its characteristics of low cost,flexible launch mode and short development period.However,conventional solid-propellant have difficulties in starting and interrupting combustion because combustion is autonomously sustained after ignition Herein,we proposed a new type of solid-propellant named laser-controlled solid propellant,which is sensitive to laser irradiation and can be started or interrupted by switching on/off the continuous wave laser.To demonstrate the feasibility and investigate the controllable combustion behaviors under different laser on/off conditions,the combus tion parameters including burning rate,ignition delay time and platform pressure were tested using pressure sensor,high-speed camera and thermographic camera.The results showed that the increase of laser-on or laser-off duration both will lead to the decrease of propellant combustion performance during re-ignition and re-combustion process.This is mainly attributed to the laser attenuation caused by the accumulation of combustion residue and the change of chamber ambient temperature.Simultaneously the multiple ignition tests revealed that the increased chamber ambient temperature after combustion can make up for the energy loss of laser attenuation and expansion of chamber cavity.However,the laser-controlled combustion performance of solid propellant displayed a decrease trend with the addi-tion of ignition times.Nevertheless,the results still exchibited good laser-controlled agility of laser-controlled solid propellant and manifested its inspiring potential in many aspects of space missions.

Experimental Investigation on Combustion Performance of Solid Propellant Subjected to Erosion of Particles with Different Concentrations

A test device with rectangular channel is developed to study the combustion performance of solid propellant in high temperature particles erosion.The flowfields in this newdevice and a test device with circular channels are simulated numerically.The particle erosion experiments in these two devices are carried out under different particle concentrations.The results showthat the test device with rectangular channel can effectively improve the clarity and precision of combustion diagnosis image and can be used for research on combustion performance of solid propellant under lowconcentration particle erosion;the circular channel device has good particle convergent effect,provides high concentration particle erosion,and can be used for research on the combustion performance of solid propellant under high concentration particle erosion.The experiment data indicates that the propellant burning rate does not change obviously in lower particle concentration;the propellant with lower static burning rate increases remarkably under particle erosion,while the propellant with high static burning rate is not sensitive to the particle erosion.

A New Simulation Method for 3D Propellant Grain Burn Analysis of Solid Rocket Motor

Describes a new computer program (Regress-3D) to simulate the regression of complex 3D grain cavity and calculate the burning surface area. It has a large region of applicability in solid rocket motor design and has made new improvements compared with other available codes. User can easily and rapidly build his initial grain shapes and then obtain geometric information of his design. Considering with the calclulting results, redesigning can be performed as desire until reaching at the satisfied result. Advantages and disadvantages of this method are also discussed.

Effect of carbon dioxide concentration on the combustion characteristics of boron agglomerates in oxygen-containing atmospheres

In ramjet combustion chambers,carbon dioxide(CO_(2))produced by the combustion of carbonaceous fuel enters the chamber together with boron agglomerates.In order to investigate the effect of CO_(2)concentration present in an oxygen-containing atmosphere on the combustion characteristics and oxidation mechanisms of boron agglomerates,we used a laser ignition system,an X-ray diffractometer(XRD),and a thermogravimetric-differential scanning calorimetry(TG-DSC)combined thermal analysis system.Single-particle boron was tested in the laser-ignition experiments as the control group.The ignition experiment results showed that with a fixed O2 concentration of 20%,when the particle temperature reaches the melting point of boron,increasing CO_(2)content causes the combustion process of boron agglomerates to transition from single-particle molten droplet combustion to porous-particle combustion.Furthermore,XRD analysis results indicated that the condensed-phase combustion products(CCPs)of boron particles in a mixed atmosphere of O2 and CO_(2)contained B4C,which is responsible for the porous structure of the particles.At temperatures below 1200℃,the addition of CO_(2)has no obvious promotion effect on boron exothermic reaction.However,in the laser-ignition experiment,when the oxygen concentration was fixed at 20%while the CO_(2)concentration increased from 0%to 80%,the maximum temperature of boron agglomerates rose from 2434 to 2573 K,the self-sustaining combustion time of single-particle boron decreased from 396 to 169 ms,and the self-sustaining combustion time of boron agglomerates decreased from 198 to 40 ms.This study conclusively showed that adding CO_(2)to an oxygen-containing atmosphere facilitates boron reaction and consumption pathways,which is beneficial to promoting exothermic reaction of boron agglomerates at relatively high temperatures.

固体火箭超燃冲压发动机点火燃烧过程实验研究

为解决硼基贫氧燃料固体火箭超燃冲压发动机补燃室内硼颗粒超声速点火燃烧难题,设计制造了在超声速燃气射流掺混区域开设观察窗的点火燃烧过程试验样机,开展了含硼贫氧固体燃料的超声速点火试验。试验模拟了26 km,Ma5.9的飞行工况并通过高速摄像获得了点火燃烧过程的火焰形态。试验结果表明:掺混增强装置可以显著改善补燃室内存在的分层流动和一次燃气气固两相分离的现象,为硼颗粒提供良好的点火条件从而提升其附近硼颗粒的点火燃烧性能。通过合理设计掺混增强装置位置,将硼颗粒在一次燃气喷注口附近的高温点火区点燃比在补燃室中段点燃具有更高的燃烧效率,本文设计的燃烧组织结构在试验中实现了硼贫氧固体燃料0.812的燃烧效率。

硝酸羟胺基固体推进剂电控燃烧特性实验研究

硝酸羟胺基固体推进剂以硝酸羟胺(HAN)为氧化剂,聚乙烯醇(PVA)为粘合剂,具有可电控燃烧、产气无毒害、能量特性好的特点,然而工作性能不稳定、点火机理不明朗的问题是限制其工程实际应用的主要因素之一。针对以上问题,制备了不同成分配比的推进剂,利用扫描电镜能谱仪和电化学工作站对推进剂的微观形貌、元素分布和电导率进行分析,然后对推进剂的点火功率、推进剂温度、火焰结构、熄灭特性进行实验研究并优选了最佳成分配比。实验结果表明,推进剂中PVA的含量越高,PVA分子链交联形成的骨架结构越紧密,造成推进剂的电导率更低,所需的点火功率更高,燃烧过程中推进剂表面温度也会更高,熄灭后推进剂的自分解会更剧烈,不利于推进剂的高效燃烧和重复点火。PVA的最佳质量分数为18%。另外,电化学阻抗是影响推进剂电能注入的重要因素,PVA含量过低会导致电能难以注入,使得推进剂表面无法达到点火温度。

硝基氰胺乙烯基咪唑过渡金属配合物的合成及结构与性能分析

为探索固体推进剂含能燃烧催化剂,以硝基氰胺根(NCA)为阴离子、1-乙烯基咪唑(VIM)为配体,过渡金属Mn、Co、Ni、Cu为中心离子,制备出4种新型含能配合物,结构式分别为[Mn(VIM)_(4)](NCA)_(2)(配合物1)、[Co(VIM)_(4)](NCA)_(2)(配合物2)、[Ni(VIM)_(4)](NCA)_(2)(配合物3)、[Cu(VIM)_(4)](NCA)_(2)(配合物4);通过元素分析、红外光谱(FT-IR)、X射线晶体衍射(XRD)、感度测试和氧弹量热法对其结构和物理化学性质进行了表征;用差示扫描量热法(DSC)和热重分析(TGA)研究了配合物1~配合物4的热分解过程,并利用DSC研究了配合物1~配合物4对高氯酸铵(AP)热分解的催化作用。结果表明,配合物1~配合物4均具有较高的能量密度,其中配合物2质量能量密度(E_(g))为19.00 kJ/g,体积能量密度(E_(v))为27.18 kJ/cm^(3);4种配合物的摩擦感度值均不小于288 N,撞击感度值均大于40 J。加入质量分数5%催化剂的AP样品高温分解温度均得到提前,放热量均显著提高;其中配合物2表现出最好的催化性能,加入配合物2的AP放热量升高至1564 J/g,分解温度提前了62.0℃,活化能降至142.9 kJ/mol,表明配合物2具有作为固体推进剂含能燃烧催化剂的潜力。

降速剂及其作用位点对三组元推进剂燃烧性能的影响规律研究

针对两种典型季铵盐和金刚烷衍生物降速剂,研究了其作用位点对推进剂点火燃烧性能的影响。基于两种不同结构的Al@AP和AP@Al复合颗粒,将降速剂分布在不同作用位置:即AP颗粒内部和表面。利用扫描电子显微镜(SEM-EDS)和计算机断层扫描(CT)等技术对所制备复合颗粒微观形貌和结构进行表征,研究了降速剂对推进剂爆热、点火延迟和燃烧性能的影响。惰性降速剂的加入会使得推进剂爆热值降低,但Al@AP复合颗粒的高反应活性能部分抵消这一现象。与机械混合方式相比,当降速剂作用于AP表面时能明显延缓推进剂点火过程,其中金刚烷衍生物能使推进剂点火延迟时间增加至420 ms,比基础配方增加了192%。在基体中直接加入降速剂可以有效降低推进剂燃速,但其降压强指数效果有限。铝基复合颗粒和降速剂的共同作用能显著降低推进剂的高压燃速和压强指数:当作用位置在AP内时,两种降速剂使推进剂1~20 MPa内压强指数从基础配方的0.45分别降低至0.20和0.33,同时可使10~20 MPa压强指数分别减小了48%和49%。

降速剂及其作用位点对丁羟四组元推进剂燃烧性能的影响规律

为研究降速剂对丁羟四组元推进剂燃烧性能的影响规律,将典型季铵盐和金刚烷衍生物两种高效降速剂引入核壳结构铝基复合颗粒Al@HMX和AP@Al,使其分别作用于HMX颗粒内部和AP颗粒表面,制备了4种含降速剂的铝基复合颗粒(Al/A@HMX,Al/B@HMX,AP@Al/A和AP@Al/B);采用扫描电子显微镜对样品形貌进行了表征;采用高速红外相机拍摄推进剂燃烧过程的火焰红外照片,并对推进剂的爆热、密度、点火延迟时间和燃速进行了测试。结果表明,加入惰性降速剂会导致推进剂爆热降低,而Al@HMX复合颗粒能部分抵消这一现象,使推进剂爆热值增加了338 J/g;降速剂能够抑制AP和HMX的热分解过程,使达到AlO辐射峰值前维持低强度的“平台段”;而引入Al@HMX后,推进剂的点火延迟时间比基础配方减小49.4%;在10~20 MPa范围内两种降速剂均能有效降低推进剂燃速,在此基础上采用Al@HMX可使含季铵盐丁羟四组元推进剂20 MPa下的燃速降低7.1 mm/s(38.4%),压强指数降至0.25;当降速剂作用于AP表面时,含质量分数1%季铵盐的推进剂在20 MPa下燃速可降低5.0 mm/s(27.3%)。

丁羟三组元推进剂的增材制造及性能研究

针对传统浇注成型与直写式3D打印对固体推进剂药浆工艺性能要求相冲突的问题,为实现小型药柱的3D打印,采用添加少量定型助剂(YJ)的方法对丁羟三组元推进剂配方进行改性,对改性前后推进剂的工艺性能、力学性能、燃烧性能和能量性能进行对比分析,并探究了YJ对推进剂性能的影响。结果表明,改性后的推进剂药浆具备可控挤出和室温堆积的流变特性;YJ的加入使得推进剂在20、70℃下的最大抗拉强度分别降低0.1和0.15 MPa,断裂伸长率分别增加了12.7%和9.9%,表明YJ对其力学性能影响显著;此外,实验及理论计算表明,YJ对推进剂的燃烧性能和能量性能影响甚微,燃速最大降低0.24 mm/s,能量变化幅度均在1%以内;表明定型助剂(YJ)的加入不仅使药浆满足3D打印要求,而且对原始推进剂的整体性能没有显著负面影响。

基于高速摄像和激光吸收光谱的固体推进剂交变应变燃烧研究

针对固体推进剂燃烧特性容易受到机械应变的作用,设计开发了一套附带光学视窗的固体推进剂燃烧测量系统,并基于高速摄像以及激光吸收光谱方法,研究了机械应变对AP-HTPB固体推进剂燃速、CO_(2)浓度以及燃烧温度的影响.结果表明:对于静态应变条件,在-10%~10%的机械应变范围内,推进剂燃速、温度和CO_(2)浓度随应变量的增加而增加.其中,燃速变化范围为1.72~2.1 mm/s,推进剂燃烧温度变化范围为1490~1680 K,CO_(2)空间积分浓度变化范围为0.138~0.165 cm.对于交变应变条件,推进剂燃速变化呈现分阶段特征,即在燃烧前中期(0~1500 ms),平均燃速为1.926 ms左右,而在燃烧后半段(1500~2200 ms),测量燃速达到5 ms左右,造成该现象的原因可能是燃烧后期推进剂在交变应变的作用下发生了整体解体.另一方面,在燃烧前半段,推进剂燃烧温度和CO_(2)浓度呈现周期性变化,且变化频率与交变应变相同.

HAN基电控固体推进剂电化学性能对燃烧性能的影响

硝酸羟胺(HAN)基电控固体推进剂是一种以HAN为主要氧化剂的新型智能推进剂,通过外加电源控制其燃烧速率与启停。为探索电化学性能与燃烧性能间的关系,在研究推进剂的离子迁移率、过电位的变化规律的基础上,将其与推进剂的点火电压、点火延迟进行关联。研究发现,推进剂在外加低电压(1.6~2.8 V)刺激下,离子迁移率与电压大小正相关,推进剂基体的改良在提高离子迁移率方面的效果大于离子掺杂。在燃烧性能方面,离子迁移率影响点火电压,过电位决定点火延迟;离子迁移率越高,点火时,推进剂的电流密度与反应活性越高;当迁移率由0.09 mm^(2)·s^(-1)·V^(-1)升高至0.16 mm^(2)·s^(-1)·V^(-1)时,点火电压由210 V降至70 V,有利于拓宽电控固体推进剂的应用范围;点火电压一定时,推进剂过电位越小,电分解速率越大,点火延迟越小,当过电位由0.67 V降至0.51 V时,点火延迟由1.68 s降至0.31 s,有利于推进剂满足快速响应的战略需求。

金属复合氧化物基燃烧催化剂在固体推进剂中的研究进展

为了阐明不同结构类型的金属复合氧化物及其复合材料作为燃烧催化剂时对固体推进剂含能组分的热分解催化性能,分别对尖晶石型、钙钛矿型、类钙钛矿型、碳材料基金属复合氧化物及金属元素掺杂的金属复合氧化物的催化活性进行总结。根据其结构及负载材料的不同,分析其对固体推进剂燃烧速率、压力指数及含能组分热分解峰温、放热量及表观活化能的影响,发现金属复合氧化物及其复合材料表现出比单一金属氧化物更为优异的催化活性。最后,指出了今后金属复合氧化物基燃烧催化剂在固体推进剂中的研究方向。

Overview of Al-based nanoenergetic ingredients for solid rocket propulsion

The introduction of nano-sized energetic ingredients first occurred in Russia about 60 years ago and arose great expectations in the rocket propulsion community, thanks to the higher energy densities and faster energy release rates exhibited with respect to conventional ingredients. But, despite intense worldwide research programs, still today mostly laboratory level applications are reported and often for scientific purposes only. A number of practical reasons prevent the applications at industrial level: inert native coating of the energetic particles, nonuniform dispersion, aging, excessive viscosity of the slurry propellant, possible limitations in mechanical properties, more demanding safety issues, cost, and so on.This paper describes the main features in terms of performance of solid rocket propellants loaded with nanometals and intends to emphasize the unique properties or operating conditions made possible by the addition of the nano-sized energetic ingredients. Steady and unsteady combustion regimes are examined.

Innovative Metallized Formulations for Solid Rocket Propulsion

Several metallized solid rocket propellants,AP/Metal/HTPB in the ratio 68/18/14,were experimentally analyzed at the Space Propulsion Laboratory of Politecnico di Milano.Effects of the metals(micrometric and nanometric Al,B,Mg,and a variety of dual metals) on the performance of the propellant were studied and contrasted to a conventional micrometric aluminum(30 μm average grain size) taken as reference.It is shown that the propellant microstructure plays a fundamental role in controlling the critical aggregation/agglomeration phenomena occurring below and near the burning surface.Two specific effects of microstructure in terms of steady burning rate and average agglomerate size are illustrated.

基于高速离轴全息可视化方法的Al/AP/HTPB推进剂燃烧性能研究

为了探究金属颗粒燃烧的复杂变化过程,搭建了25kHz高速数字离轴全息成像系统,对Al/AP/HTPB推进剂燃烧过程中颗粒的典型现象进行了时间分辨三维可视化研究;通过三维可视化研究分析了推进剂燃烧中金属铝颗粒燃烧典型现象,其中包括燃烧过程中推进剂燃面颗粒的剥落现象、颗粒/颗粒团的微爆炸现象、以及燃烧颗粒存在的尾流与火焰面现象,并且通过数字离轴全息技术对观测到的铝颗粒进行了时间分辨的三维重建。结果表明,离轴全息技术消除了孪生像的干扰,可以获得清晰的颗粒重建图像,并且其独立配置的参考光在很大程度上消除了火焰对光传播的影响,有效地抑制了颗粒周围折射率急剧变化所带来的像差。表明高速数字离轴全息三维成像技术在固体推进剂燃烧可视化诊断方面具有较强的实用性,并具有广阔的应用前景。

一种高能固体推进剂定压燃烧温度的确定方法

根据推进剂爆热与燃烧温度的正相关性,提出高能固体推进剂实际定压燃烧温度的确定方法;阐述并明确了爆热末态产物为准稳态非绝对稳态(自由能最小)的事实;根据高温燃烧产物降温到298 K确定了爆热末态燃烧产物,计算了推进剂理论定压爆热;测试不同样品质量的推进剂爆热,得到最大爆热值作为推进剂的实际定容爆热;通过测试定容爆热实验末状态的实际温度和压强,建立了定容爆热末状态燃气摩尔数的测试方法,获得了实际最大定压爆热。结果表明,根据爆热效率得到某高能推进剂在6.86 MPa、10 MPa、15 MPa和20 MPa下的实际能达到的最大定压燃烧温度分别为3598.02 K、3636.40 K、3675.54 K和3702.82 K,20 MPa下的温度误差最大为252.28 K,该值低于传统测温技术的误差。因此通过爆热间接确定高能固体推进剂的实际定压燃烧温度是合理的。

一种评定固体推进剂燃烧效率的方法——爆热效率

针对当前理论爆热仅有概念没有计算实例的现象,通过测试不同样品量的推进剂定容爆热并对最佳质量的燃气进行气相色谱实验,分析了最小自由能法计算推进剂燃烧产物的应用界限。研究认为,最小自由能法只能用于计算发动机燃烧室高温下的绝热定压燃烧产物,不适用于计算常温下的爆热产物。提出并明确了推进剂理论定压和定容爆热的计算方法;建立并验证了降温到298K燃烧产物的确定方法。针对当前推进剂燃烧效率大多为定性评判的现象,将爆热效率定义为实际定容爆热与理论定容爆热之比,用来定量表征推进剂的燃烧效率,可以实现用少量推进剂对其燃烧效率进行综合评判,尤其适用于推进剂配方研制阶段。