A new method for evaluating the firing precision of multiple launch rocket system based on Bayesian theory

How to effectively evaluate the firing precision of weapon equipment at low cost is one of the core contents of improving the test level of weapon system.A new method to evaluate the firing precision of the MLRS considering the credibility of simulation system based on Bayesian theory is proposed in this paper.First of all,a comprehensive index system for the credibility of the simulation system of the firing precision of the MLRS is constructed combined with the group analytic hierarchy process.A modified method for determining the comprehensive weight of the index is established to improve the rationality of the index weight coefficients.The Bayesian posterior estimation formula of firing precision considering prior information is derived in the form of mixed prior distribution,and the rationality of prior information used in estimation model is discussed quantitatively.With the simulation tests,the different evaluation methods are compared to validate the effectiveness of the proposed method.Finally,the experimental results show that the effectiveness of estimation method for firing precision is improved by more than 25%.

Derivation of a Revised Tsiolkovsky Rocket Equation That Predicts Combustion Oscillations

Our study identifies a subtle deviation from Newton’s third law in the derivation of the ideal rocket equation, also known as the Tsiolkovsky Rocket Equation (TRE). TRE can be derived using a 1D elastic collision model of the momentum exchange between the differential propellant mass element (dm) and the rocket final mass (m1), in which dm initially travels forward to collide with m1 and rebounds to exit through the exhaust nozzle with a velocity that is known as the effective exhaust velocity ve. We observe that such a model does not explain how dm was able to acquire its initial forward velocity without the support of a reactive mass traveling in the opposite direction. We show instead that the initial kinetic energy of dm is generated from dm itself by a process of self-combustion and expansion. In our ideal rocket with a single particle dm confined inside a hollow tube with one closed end, we show that the process of self-combustion and expansion of dm will result in a pair of differential particles each with a mass dm/2, and each traveling away from one another along the tube axis, from the center of combustion. These two identical particles represent the active and reactive sub-components of dm, co-generated in compliance with Newton’s third law of equal action and reaction. Building on this model, we derive a linear momentum ODE of the system, the solution of which yields what we call the Revised Tsiolkovsky Rocket Equation (RTRE). We show that RTRE has a mathematical form that is similar to TRE, with the exception of the effective exhaust velocity (ve) term. The ve term in TRE is replaced in RTRE by the average of two distinct exhaust velocities that we refer to as fast-jet, vx1, and slow-jet, vx2. These two velocities correspond, respectively, to the velocities of the detonation pressure wave that is vectored directly towards the exhaust nozzle, and the retonation wave that is initially vectored in the direction of rocket propagation, but subsequently becomes reflected from the thrust surface of the combustion chamber to exit through the exhaust nozzle with a time lag behind the detonation wave. The detonation-retonation phenomenon is supported by experimental evidence in the published literature. Finally, we use a convolution model to simulate the composite exhaust pressure wave, highlighting the frequency spectrum of the pressure perturbations that are generated by the mutual interference between the fast-jet and slow-jet components. Our analysis offers insights into the origin of combustion oscillations in rocket engines, with possible extensions beyond rocket engineering into other fields of combustion engineering.

China Accelerating Development of Two Large Reusable Rocket Models

BEIJING,March 12(Xinhua)——China′s development of both 4⁃meter⁃diameter and 5⁃meter⁃diameter reusable rockets is being accelerated,with their inaugural flights scheduled for 2025 and 2026,respectively.The manufacturing of these two new large reusable launching vehicle models is a response to growing demand in the commercial space market,said the China Aerospace Science and Technology Corporation.

火箭风暴:2024 Triumph Rocket 3 Storm R

2020年,凯旋推出了巨无霸“火箭三缸”Rocket 3 R,以排量高达2458mL的水平并列三缸发动机震惊世人,当仁不让地成为最大的量产摩托车。4年过去了,“火箭三缸”依然是寂寞的王者。能超越“火箭三缸”的,只有“火箭三缸”的自我进化。2024年,凯旋再次重磅出击,推出了新一代“火箭风暴”Rocket 3 Strom R,功率更大,扭矩更强,技术更先进,安全性能更好,掀起声势浩大的公路风暴!

Formalism of rotating-wave approximation in high-spin system with quadrupole interaction

We investigate the rotating wave approximation applied in the high-spin quantum system driven by a linearly polarized alternating magnetic field in the presence of quadrupole interactions.The conventional way to apply the rotating wave approximation in a driven high-spin system is to assume the dynamics being restricted in the reduced Hilbert space.However,when the driving strength is relatively strong or the driving is off resonant,the leakage from the target resonance subspace cannot be neglected for a multi-level quantum system.We propose the correct formalism to apply the rotating wave approximation in the full Hilbert space by taking this leakage into account.By estimating the operator fidelity of the time propagator,our formalism applied in the full Hilbert space unambiguously manifests great advantages over the conventional method applied in the reduced Hilbert space.

Measurement of alumina film induced ablation of internal insulator in solid rocket environment

This study investigates the ablation of internal insulation induced by the deposition of an alumina film with different lateral film speeds.A sub-scale test solid rocket motor(SRM)was designed in an impinging jet configuration to form an alumina film on the sample and to encourage the lateral movement of the film by a high-speed wall jet.Fifteen static fire tests of the test SRM were conducted with six different jet velocities(V_(jet)=100 m/s,150 m/s,200 m/s,268 m/s,330 m/s,and 450 m/s)that indirectly affected the velocity of the wall jet and the deposition rate of alumina droplets.The ablation velocity was deduced from the difference in the sample thickness after a test using a coordinate measuring machine.The droplet deposition mass flux and wall jet velocity were obtained via two-phase flow simulation with the same jet velocity and effective pressure.As a result,the characteristics of alumina-induced ablation and the changes in ablation with jet velocities were obtained.The area within0.8×jet diameter was focused upon,where the ratio of ablation velocity to incoming alumina mass was constant for each jet velocity,and showed a similarity in jet structure.When the ablation velocity was increased from 2.05 to 9.98 mm/s with increasing jet velocity,the ratio of the ablation velocity and alumina mass flux decreased from 1.07×10^(-4)to 0.49×10^(-4)m^(3)/kg as Al_(2)O_(3)-C reactions became less efficient with a reduced residence time of the film.Because the decrease in residence time by the wall jet is more pronounced for slow reactions involved in Al_(2)O_(3)-C reactions,fast reactions in Al_(2)O_(3)-C reactions are less affected and result in a convergence of the volumetric rate of ablation per unit mass of alumina.

基于SmartRocket的箭载嵌入式软件集成测试

针对目前箭载嵌入式软件集成测试工具测试用例编写时间长、人力成本高、测试用例报告适用性低等方面的问题,提出一种基于Smart Rocket的箭载嵌入式软件集成测试方法。SmartRocket采用广度优先自顶向下集成测试方法确定集成测试组合,基于动态符号执行技术自动生成集成测试用例,在满足函数覆盖和调用覆盖指标下生成与实际要求相符的测试用例报告。通过实验发现,该集成测试方法能够很好地实现集成测试用例自动化,提高了测试用例报告适用性,一定程度地增加了箭载嵌入式软件集成测试人员的测试效率。

基于网侧资源协调的自储能柔性互联配电系统日前-日内优化

为优化储能型智能软开关(E-SOP)、传统网侧调控设备的运行策略,提升高渗透率光伏配电网运行水平,提出了基于网侧资源协调的自储能柔性互联配电系统日前-日内优化模型。首先,提出了“源网荷储”下多种网侧资源协调利用的多时间尺度优化控制架构;其次,在日前阶段构建计及有载调压变压器、无功补偿装置、联络开关、E-SOP和柔性负荷的长时间有功-无功联合优化模型,在日内阶段建立多目标滚动优化模型,并采用线性化方法将模型转换为混合整数线性规划模型;然后,基于KL散度构建源荷不确定性模糊集,给出日前两阶段分布鲁棒优化模型转化过程和基于列与约束生成算法的求解方法;最后,算例仿真验证所提方法的有效性。

A Space-Flight Ship Travelling by a Plasma Rocket Engine from the Earth Ground to the Moon

Since a thrust of an ion rocket engine is much weaker than the one of a chemical fuel engine, nowadays, ion engines are used mainly in spaces where gravities are very weak. Here, as a powerful plasma rocket to make a heavy ship get out from the gravity-sphere of the earth without relying on an atomic power rocket, an ion-velocity booster is investigated. It is a main challenge how to protect the engine wall from the melting due to collisions of ions which grow into high-energy particles.

助推段火箭发动机喷焰红外辐射特性数值计算

助推段火箭发动机尾喷焰具有显著的红外辐射特性,成为天基红外预警系统重点关注的对象。以三组元固体发动机为研究对象,建立连续流域和过渡流域的尾喷焰反应流场计算模型,采用两级助推段质心弹道法确定飞行弹道,通过反向寻优法完备发动机型谱参数,基于等效单喷管法简化流场计算,基于窄谱带法和视在光线法计算尾喷焰的红外辐射特性。研究结果表明:复燃效应使得尾喷焰红外辐射光谱强度升高约20倍,谱带积分强度升高约10倍;尾喷焰红外光谱存在2.7μm和4.3μm的特征峰值;低空流域的光谱强度较高,且不同谱带内积分强度随高度的升高呈现出波动规律和暗区现象;助推段尾喷焰辐射强度差值高达3个量级,尤其在4.3μm较为明显。所得成果可为火箭在助推段的探测和识别提供一定的理论参考。

重复使用运载火箭液体动力技术发展

重复使用是未来运载火箭更新换代的技术发展趋势,是降低航天发射成本、实现规模化航天发射的有效途径。重点概述了国内外垂直起降重复使用运载火箭动力技术的发展现状,分析了垂直起降重复使用运载火箭发射和回收全任务剖面,总结了垂直起降重复使用运载火箭动力技术的特点,包括宽范围入口压力多次启动技术、大范围快速高精度推力调节技术、故障诊断及健康管理技术、状态检测与维修维护技术等。

Rapid Launch of Medium-Sized Solid Carrier Rockets at Sea With High Reliability and Convenient Intelligence

Gravity-1 solid-propellant carrier rocket utilizes a three-vertical testing and launch mode, and adopts a sea-based launch method. As the demand for satellite launches continues to grow, the scarcity of launch site resources,and the consideration of cost savings, the need for rapid testing and launch of carrier rockets has become increasingly strong. The capability of rapid testing and launch has even become one important aspect of evaluating a rocket. This paper focuses on the characteristics of the Gravity-1 solid-propellant strap-on medium-sized carrier rocket and designs and implements a highly reliable, convenient, and intelligent low-cost rapid testing and launch solution. The main aspects include the design of a highly reliable dual-redundant ground architecture and the application of ground-based shelf products.

液体火箭发动机健康监控技术研究进展

液体火箭发动机健康监控技术作为保障运载火箭安全、可靠发射的核心关键技术,经过几十年的发展,有力推动了航天事业的进步。介绍了液体火箭发动机健康监控技术中故障检测与诊断、容错控制与健康监控系统研制等技术的研究现状与发展趋势;梳理了健康监控领域面临的重难点问题,并提出相应的解决方案。分析展望了液体火箭发动机健康监控技术未来发展趋势,为从事火箭发动机健康监控技术研究的科研人员提供参考。

火箭触发闪电的回击电流及电场特征分析

2019年夏季在广州开展了火箭引雷试验,成功地触发了14次闪电,总回击(RS)数达74次,平均回击数为5.3次,单次闪电的最大回击数为14次。该文对火箭引雷的回击电流及不同距离电场波形特征进行了分析。回击的间隔时间、峰值电流、10%~90%上升时间、半峰值宽度、1 ms转移电荷和1 ms作用积分的几何平均值分别为38.45 ms、12.38 kA、0.25μs、8.31μs、0.68C和2.12×10^(3)A^(2)·s,回击电流的转移电荷量和作用积分量均与回击峰值电流呈幂函数相关性。回击的先导电场强度峰值和回击电场强度峰值会随距离的增大而减小,而电场的10%~90%上升时间和半峰值宽度会随距离的增大而增大。回击电场强度峰值会随距离的增大而呈现出幂函数减小,且峰值电流与不同距离的先导电场强度峰值和回击电场峰值均呈现一定线性相关,距离越大,其线性拟合关系越好。

火箭煤油传递特性分子动力学模拟研究

针对火箭煤油中添加减阻剂后导致对流换热系数减小、传热效果减弱问题,从微观尺度研究纳米颗粒的加入对火箭煤油黏度和导热系数的影响关系。通过分子动力学(MD)模拟方法,采用Green-Kubo方程的平衡分子动力学(EMD)模拟对火箭煤油传热规律进行系统研究。模拟常温下(T=303K)加入传热增强剂纳米颗粒后的火箭煤油黏度与传热特性。以纳米颗粒质量分数与纳米颗粒粒径为变量,对纳米流体的导热强化规律进行研究,在此基础上将导热油与碳氢燃料分别加入减阻火箭煤油中并与纳米流体的黏度与导热系数进行比较,模拟结果表明增加纳米颗粒质量分数可使粒子间距离减小,加强了粒子间的相互作用以及热传递过程,可有效增加流体黏度和导热系数。

火箭引雷至架空线路与地面电流对比分析

2018—2019年夏季在广州从化开展了火箭引雷试验,雷击对象分为两种,分别是架空线路和地面。对比了两种引雷情况下各阶段雷电流的电流波形参数的差异,发现引雷至地面情况下初始阶段的最大电流、平均电流、转移电荷、作用积分分别是引雷至架空线路情况下的2.8、2.4、2.0、5.3倍。引雷至地面情况下回击的上升时间几何均值为0.25μs,小于引雷至线路情况下的0.60μs,而对于电流峰值等其他回击波形参数而言,差异不明显。M分量与初始连续电流脉冲类似,引雷至线路情况下的电流峰值、转移电荷、连续电流水平均小于引雷至地面情况下的。基于诺顿电流源等效电路模型分析两种引雷情况下雷电流不同阶段存在差异的原因,可能是雷电在不同电流阶段时其通道等效阻抗不一样所导致的。

长时间工作固体火箭发动机长尾喷管外绝热装置设计

固体火箭发动机长尾喷管周围一般安装有电子仪器,故而对喷管壳体表面温度有严格的限制。对于某长时间工作的发动机,为满足喷管壳体表面温度要求,同时降低消极质量,设计了一种新型双腔外绝热装置,以实现在较小的消极质量下具有较好的绝热性能。对该绝热装置进行了内流场及传热计算分析,而后进行了绝热结构优化及试验验证。结果表明:该双腔外绝热装置可以将长时间工作固体火箭发动机长尾喷管外壁温度从483℃降低至103℃,降低约78.7%;而采用同等材料同等质量的常规外防护装置只能从483℃降低至227℃,降低约53.0%;而要达到相同的防护效果,同等材料常规外防护装置的质量将增加约160%。双腔外绝热装置中段及尾端处温度随时间变化趋势与实测结果吻合较好,有望应用于其他长时间工作发动机。

节流式燃/氧分离组合固体发动机推力性能调节数值研究

节流式燃/氧分离组合固体发动机将燃烧室分为富燃燃烧室和富氧燃烧室,采用富燃固体推进剂和富氧固体推进剂串联装填的形式,在两个燃烧室中间增加节流阀,通过调节节流阀开度实现推力调节。利用商业软件对节流式燃/氧分离组合固体发动机在节流阀流道半径不同时的工作状态进行模拟,研究节流式燃/氧分离组合固体发动机的工作特性,获得节流式燃/氧分离组合固体发动机燃气流动和掺混特征,揭示节流式燃/氧分离组合固体发动机推力调节机理以及节流阀流道半径对节流式燃/氧分离组合固体发动机燃烧效率的影响规律。研究发现:节流阀流道开度(即节流阀流道半径R)减小可增大发动机推力;随着节流阀开度的减小,节流阀出口燃气马赫数随之增大,而发动机燃烧效率却出现先增大后减小的规律,当节流阀出口燃气流动达到临界状态时发动机燃烧效率出现最大值。

基于Hessian局部线性嵌入和MLP-Mixer的液体火箭发动机涡轮泵轻量化故障诊断框架

作为液体火箭发动机推进剂输送系统的关键部件,涡轮泵的运行状态直接影响着整个运载系统的性能,然而,现有的故障诊断方法往往面临特性参数选择片面及计算复杂度高等问题。针对上述局限,提出了面向涡轮泵的轻量化故障诊断框架。所提方法利用Hessian局部线性嵌入算法对信号时域、频域及时频特征进行降维,并引入一种轻量化的深度学习模型MLP-Mixer作为分类器,进而实现不同故障状态的辨识。采用某型号涡轮泵试车数据验证了所提方法的有效性,结果表明,该方法能够在保障诊断精度的同时有效降低计算复杂度,提高诊断效率。

泵入口参数对全流量补燃循环发动机动态特性影响研究

全流量补燃循环发动机性能高,并具备便于重复使用的优势,是未来重复使用运载器的最佳选择之一。在液体火箭发动机启动过程中,泵入口推进剂的压力和温度变化会对全流量补燃循环发动机动力系统的启动特性造成一定的影响。为分析其影响规律,建立了涡轮泵、预燃室、推力室、两相流和混合气体组件等动力学模型并进行求解。结果表明,随着燃料泵入口压力升高或氧化剂泵入口压力降低,涡轮泵系统参数的波动将会逐渐增大,甲烷温度和液氧温度所产生的影响要强于甲烷压力和液氧压力,而液氧的影响则强于甲烷。为动力系统中泵入口推进剂状态参数控制提供了一定的指导。