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.

Flight control for a flexible air-breathing hypersonic vehicle based on quasi-continuous high-order sliding mode

The focus of this paper is on control design and simulation for the longitudinal model of a flexible air-breathing hypersonic vehicle（FAHV）.The model of interest includes flexibility effects and intricate couplings between the engine dynamics and flight dynamics.To overcome the analytical intractability of this model,a nominal control-oriented model is constructed for the purpose of feedback control design in the first place.Secondly,the multi-input multi-output（MIMO） quasi-continuous high-order sliding mode（HOSM） controller is proposed to track step changes in velocity and altitude,which is based on full state feedback.The simulation results are presented to verify the effectiveness of the proposed control strategy.

Reference tracking control for flexible air-breathing hypersonic vehicle with actuator delay and uncertainty

This paper considers the problem of reference tracking control for the flexible air-breathing hypersonic flight vehicle with actuator delay and uncertainty.By constructing the Lyapunov functional including the lower and upper bounds of the time-varying delay,the non-fragile controller is designed such that the resulting closed-loop system is asymptotically stable and satisfies a prescribed performance cost index.The simulation results are given to show the effectiveness of the proposed control method,which is validated by excellent output reference altitude and velocity tracking performance.

Back-Stepping Control for Flexible Air-Breathing Hypersonic Vehicles Based on Uncertainty and Disturbance Estimator

A theoretical framework of nonlinear flight control for a flexible air-breathing hypersonic vehicle(FAHV)is proposed in this paper.In order to suppress the system uncertainty and external disturbance,an uncertainty and disturbance estimator(UDE)based back-stepping control strategy is designed for a dynamic state-feedback controller to provide stable velocity and altitude tracking.Firstly,the longitudinal dynamics of FAHV is simplified into a closure loop form with lumped uncertainty and disturbance.Then the UDE is applied to estimate the lumped uncertainty and disturbance for the purpose of control input compensation.While a nonlinear tracking differentiator is introduced to solve the problem of“explosion of term”in the back-stepping control.The stability of the UDE-based control strategy is proved by using Lyapunov stability theorem.Finally,simulation results are presented to demonstrate the capacity of the proposed control strategy.

Effect of Short-Term Exposure to Mercuric Chloride on the Air-Breathing Catfish,Heteropneustes fossilis──Ⅱ.Scanning Electron Microscopic Study of the Gill

Alterations in the respiratory epithelium of gills of a catfish, Heteropneustes fossilis by the sublethal and lethal concentrations of HgCl2 have been observcd using scanning electron microscope. Fish exposed to 0.1 ppm shows a thin coat ot mucus, but the continuous layer of mucus is not retained in 0.3 ppm after 24 h of exposure. Epithelial cells of the gill filament are hyperplastic. Other changcs in the respiratory epithelium are contraction of epithelium, formation of interlamellar bridge and diminishing number of microridges. Massive destruction of the respiratory epithelium by deep lesions and sloughing of epithelial layer in 0.5 and 1.0 ppm solutions after 6-12 h of exposure apparently causes death to the fish. SEM observations are well in conformity with the findings made at light microscopic level

HOSVD-based LPV modeling and mixed robust H_2/H_∞ control design for air-breathing hypersonic vehicle

This paper focuses on synthesizing a mixed robust H_2/H_∞ linear parameter varying（LPV） controller for the longitudinal motion of an air-breathing hypersonic vehicle via a high order singular value decomposition（HOSVD） approach.The design of hypersonic flight control systems is highly challenging due to the enormous complexity of the vehicle dynamics and the presence of significant uncertainties.Motivated by recent results on both LPV control and tensor-product（TP） model transformation approach,the velocity and altitude tracking control problems for the air-breathing hypersonic vehicle is reduced to that of a state feedback stabilizing controller design for a polytopic LPV system with guaranteed performances.The controller implementation is converted into a convex optimization problem with parameterdependent linear matrix inequalities（LMIs） constraints,which is intuitively tractable using LMI control toolbox.Finally,numerical simulation results demonstrate the effectiveness of the proposed approach.

Experimental Investigation and Numerical Simulation of Air-Breathing Mode for Laser Propulsion

Laser propulsion is a new concept technique of propulsion and will have important application in future space technology. There are two main driving types: the air-breathing mode and the rocket ablation mode. Vertical flight experiments have been carried out with a simple paraboloid type lightcraft in the air-breathing mode by TEA-CO_2 laser. In simulation a new model is used for LSD/LSC wave, the result shows that the momentum coupling coefficient increases with the increase of the pulse energy.

Diagnostics of Air-Breathing Electric Thrusters

Air-Breathing Electric Thrusters (ABET) diagnostics is addressed, with on-ground prototypes in mind. It is based on detailed volume averaged Global Models and focuses on emission spectroscopy. Notably, the obtained optical emission spectroscopy diagnostics tools give important information about the thruster propellant constitution and also the ionization degree of each constituent. This allows for trade-off between various prototypes on the basis of their characterization and optimization.

Effect of Short-Term Exposure to Mercuric Chloride on the Air-Breathing Catfish,Heteropneustes fossilis──Ⅰ.Light Microscopic Study of the Gill

Histopathological effects of sublethal (0.1, 0.2, 0.3 ppm) and lethal (0.5, 1.0 ppm)solutions of mercuric chloride (HgCl2) have been studied on the gills of the air breathing catfish, Heteropneustes fossilis at the intervals ranging from 6h to 15 days. Thickening of the epithelium, swelling and hyperplasia of the mucous cells, fusion of secondary lamellae, formation of interlamellar bridge and deposition of mucous over the entire surface are some noteworthy features of mercury poisoning in sublethal concentrations. Acute pathological manifestations are formation of subepithelial space, sloughing of the epithelial layer, hemorrhage and hypertrophy of the mucous cells. Causes and impact of these cellular alterations affecting survival of the fish have been discussed

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.

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.

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。

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。