Loitering munitions integrate multiple functions such as reconnaissance, attack, and assessment. They have low usage costs, flexible combat deployment, and are difficult to detect and intercept. They have significantly enhanced the strike accuracy and ability against targets and have become important weapons in the Russia-Ukraine conflict, attracting great attention from major military powers. This paper summarizes the characteristics of loitering munitions, introduces the development status of major loitering munitions at home and abroad, including the “Switchblade” series and “BattleHawk” series of the United States, the “KUB” series and “Lancet” series of Russia, and the “Hero” series and “Harpy” series of Israel, which have practical applications. It comparatively analyzes the performance parameters of relevant loitering munitions, combs the key technologies in the development of loitering munitions, including power technology, anti-interference technology, modular technology, and intelligent cooperative combat technology, summarizes the development trends of loitering munitions, and on this basis, analyzes the combat applications of loitering munitions and the defense means against them.

In order to improve the detection ability of infrared small targets in complex backgrounds, a small infrared target detection algorithm based on joint gradient discrimination and adaptive matching was proposed. The suspected target area is screened out for the first time in the image through multi-directional gradient features, and the adaptive model is generated by using the grayscale information in the region for re-judgment. Quantitative evaluation was established for gradient judgment and adaptive model matching, and confidence functions were introduced to evaluate different suspected target areas and screen out suspected targets. In order to enable the algorithm to be applied in dynamic platforms such as UAVs, an embedded system was built to realize the detection of small targets in real scenes by the detection system through infrared camera framing. By testing different public datasets and comparing with WSLCM (weighted strengthened local contrast measure) and TLLCM (tri-layer local contrast measure) algorithms in different complex scenarios, the proposed algorithm has good adaptability, and the recognition rate under different samples is more than 92%. The algorithm is processed by hardware acceleration through the customization of the IP core of the embedded platform and the co-design of software and hardware, and the real-time video frame rate is greater than 30 frame/s, which verifies its effectiveness.

In modern warfare, the damage efficacy of high-speed penetration weapons against high-strength targets such as underground bunkers and reinforced structures has become a focal point of research. This paper systematically reviews the research progress on the dynamic behavior of projectile materials, constitutive models, and structural responses under high-speed penetration. It analyzes the mechanisms of strain hardening, thermal softening, and adiabatic shear deformation under the coupled effects of high temperature and high strain rate, and compares the applicability of typical constitutive models such as the Johnson-Cook model. Key factors influencing mass erosion, critical instability velocity, and structural failure during high-speed penetration are emphasized, along with their underlying mechanisms. Additionally, technical approaches to enhance penetration capabilities through material optimization and structural design are explored, providing valuable references for researchers in related fields.

　　航天发射技术是研究运载火箭、各类飞行器等的发射原理、发射方式和发射设施设计、制造、试验和使用的工程科学技术，是航天技术的一个分支和重要组成部分，是一门综合性、系统性极强的学科方向。航天发射技术的发展水平决定了一个国家航天活动和国防保障区域的范围，反映了一个国家工程科学和基础工业的综合水平。
　　近年来，为适应航天运载工具的快速发展，全球航天发射技术正朝着规模化、复杂化、自动化与国际化的方向迈进。我国航天发射技术也获得了长足的进步，取得了丰硕的成果，如海南发射场的建成有力保障了我国未来各类航天任务的开展；以“长征-11”运载火箭为代表的机动航天发射技术的出现为我国应急航天发射任务提供了多元化发射手段。这些进展不仅彰显了我国航天技术的创新实力，也为未来深空探测、商业发射和国防需求奠定了坚实基础。
　　当前，各航天大国均把发展先进的发射和运载技术作为保持其在航天领域领先地位的战略部署之一。无论是空间应用、科学探测、载人航天、国际商业发射与国际合作，还是国防建设，都对发射技术提出了新的要求，亟待发展规模更大、成本更低、兼容性更强、可靠性更高、发射周期更短、更加智能的大型航天发射场；随着航天发射任务的大幅增加和各类应急发射需求的大量涌现，提出了发展快速机动的低成本航天发射方式；面向低碳、绿色、高效、低成本发射的需求，提出了地面超高速发射一级直接入轨等新型发射概念；此外，深空探测任务的推进，使得地外天体发射技术成为前沿研究热点。
　　在此背景下，《弹箭与制导学报》编辑部联合北京理工大学姜毅教授团队，策划推出本期“火箭导弹发射技术”专刊，旨在汇聚领域内最新研究成果，推动学术交流与技术突破。本期专刊特邀姜毅教授(北京理工大学)、姚建勇教授(南京理工大学)担任编审，共收录18篇高质量论文，内容涵盖发射技术理论创新、工程实践及未来发展趋势，以期为国家航天事业发展提供有益参考。
　　我们谨向所有参与本专刊撰写的专家学者致以诚挚谢意，也期待本期专刊能为航天发射技术的进步注入新动力，助力我国迈向航天强国的宏伟目标。

The launch of the Mars ascent vehicle(MAV) is the first step for ascent from the Martian surface to orbit around Mars, a critical step in the Mars sample return mission. In the MAV's inclined hot launch process, the force-thermal impacts are significant and gas flow structure is complex, so when designing its launch system, it is necessary to consider the complex force-thermal effect in launching process. This study employs computational fluid dynamics to conduct a numerical simulation of the force-thermal impacts of the MAV's inclined thermal launch on the Martian surface, with comparisons to similar launch conditions on earth. The findings reveal that during the inclined hot launch from Mars, the MAV's maximum surface temperature reaches 2 868 K, while the launch platform attains a peak temperature of 2 908 K. Furthermore, during ejection, the platform's pitch-up moment progressively increases, the launch apparatus may topple under Mars' low-gravity conditions. Notable differences are observed between inclined hot launch processes on Mars and Earth; the MAV launch on Mars exhibits greater stability yet endures more severe force-thermal impacts on both the platform and the MAV.

In recent years, the development of large-scale low-earth-orbit (LEO) satellite constellations has been progressing at an unprecedented pace, and their potential applications in the military domain have become increasingly prominent. These constellations possess unique technical characteristics that enable them to restructure the traditional kill chain, effectively addressing core challenges such as reconnaissance delays and insufficient cross-domain coordination in precision strikes. This paper provides a comprehensive introduction to the development of large-scale LEO satellite constellations and delves into their application directions in the military field. By examining typical combat scenarios from the perspectives of the kill chain and kill web, this paper analyzes how large-scale LEO satellite constellations can enhance the construction of the kill web and facilitate the closure of the kill chain in the field of precision strike. The findings of this study offer valuable insights and references for the construction of a global and systematic kill web based on a large-scale low-Earth-orbit constellation. Moreover, this research holds significant reference value for the development of China’s anti-access/area denial (A2/AD) strategic system architecture. As related projects in our country continue to advance steadily, the exploration of the potential of large-scale LEO satellite constellations in enhancing military capabilities becomes even more crucial.

The visible light image of the rocket emitted at night provides clear texture and color information of the tail flame, while the infrared image offers a complete outline of the arrow body and saturated tail flame. Due to the high complementarity between the two, fusion of visible and infrared images has been a focal point in research aimed at enhancing rocket imaging observation capability at launch sites. This paper introduces a fusion method for nighttime visible and infrared images based on rocket tail recognition and region reconstruction. In this method, the invariant features of the arrow body are used to search for the tail secant line, which is then utilized to divide the region. Finally, a fusion strategy for region reconstruction is employed to merge the infrared arrow body with the visible tail flame. This approach effectively prevents loss of feature information from both types of images and maximizes semantic information fusion. Experimental results demonstrate that compared with current main image fusion methods, our proposed method significantly improves fusion effectiveness while better preserving feature details from both visible and infrared source images. Furthermore, it achieves quasi-real-time effects within a short time frame.

Liquid propellant sloshing alters the rocket's center of mass and generates dynamic loads on storage tanks, adversely affecting launch stability and safety. Aiming at the coupling problem between attitude deviation and liquid sloshing during the multi-stage piston eccentric ejection process of liquid rockets, a fluid-structure interaction model of the liquid rocket and its launch system was established by using the finite element method and smooth particle hydrodynamics method (FEM-SPH). The entire eccentric ejection process of the liquid rocket was simulated and analyzed, and the influence of the number and spatial distribution of the liquid rocket adapter on the initial disturbance of the rocket, the force on the rocket tank and the force characteristics of the adapter itself was investigated. The results indicate that the eccentric ejection of a liquid rocket induces deviations in the rocket's yaw angle. During the ejection process, the lateral sloshing loads on the oxidizer tank exceed those on the fuel tank at the same stage, and the force variations on the first-stage adapter located in the upper section of the rocket are more pronounced. When the adapters are distributed in a sparse upper and dense lower configuration, the rocket's yaw angle and the force variation on the adapter are the largest. When the adapters are arranged in a dense upper and sparse lower configuration, the peak sloshing forces on all tanks are the highest. Increasing the adapter from four to six circles reduces the rocket ejection yaw angle by 26.9%, the peak value of lateral slosh force on each tank by an average of 24.1%, and the change in force on the first adapter by 34.6%.

Infrared tracking of small and weak targets is an indispensable technology in anti-unmanned aerial vehicle systems. The small size, low contrast, strong maneuverability, and complex background of unmanned aerial vehicles pose lots of difficulties and challenges to tracking algorithms. To address these issues, an improved visual tracking Siamese fully convolutional classification and regression network (ISIAMCAR) is proposed. Initially, a side window filter is incorporated into the model to enhance the feature extraction capabilities of the deep convolutional neural network for small targets within the original network. Additionally, a location-aware module is integrated into the network to maintain the efficient propagation of deep convolutional features for infrared weak and small targets detection and tracking, thereby improving the classification and regression capabilities for infrared small and weak targets. The visual and numerical experimental results demonstrate that our approach outperforms existing state-of-the-art methods, achieving the highest tracking success rate and precision rate. Furthermore, the typical computational time of our proposed model has been thoroughly tested, and the ISIAMCAR network we designed meets the real-time requirements of the target tracking task in the anti-unmanned aerial vehicle system. Finally, ablation experiments have been designed and conducted to verify the effectiveness of the side window filtering and location-aware module in the proposed model.

Aiming at the path planning problem of UAVs in complex scenes,this paper proposes a coupling algorithm that uses the improved A-star algorithm and the improved dynamic window method for path planning,so that the UAV has the ability to avoid static and dynamic obstacles.In terms of global planning,by improving the evaluation function of the A^* algorithm,a path planning algorithm that does not rely on the obstacle expansion map is proposed,so that the UAV can plan a safe path against a priori static obstacles.In terms of local planning,an evaluation function for handling dynamic obstacles is added,so that the UAV has good obstacle avoidance capabilities when facing high-speed dynamic obstacles.Aiming at the problem of too many inflection points on the path resulting in frequent acceleration and deceleration after improving the global planning algorithm,a redundant inflection point deletion strategy was proposed.The simulation results show that compared with the traditional algorithm,the improved algorithm has better obstacle avoidance ability and shorter driving trajectory,which verifies the practicability of the algorithm.

As unmanned combat systems have increasingly emerged as a critical asymmetric warfare approach in naval operations, small unmanned surface vehicles (USVs) have become focal points of international military competition.While leading military powers are actively developing and equipping USVs while advancing countermeasure technologies, significant challenges remain. This study adopts the offensive-defensive dynamics of USVs as its analytical framework, examines the evolving nature of naval warfare in contemporary conflicts, and systematically analyzes the technical specifications, operational tactics, and defensive countermeasures associated with advanced small USVs deployed internationally.Building upon the concept of the USV closed-loop kill chain, we proposes an integrated anti-USV defense architecture that synergistically combines radar surveillance, electronic warfare (reconnaissance/jamming), infrared/electro-optical detection, multi-domain interception systems (air-to-surface/underwater), and asymmetric countermeasures. The findings establish a theoretical foundation for advancing USV-related technologies while providing actionable insights for the development and tactical deployment of next-generation counter-USV systems.

It is an important direction for the future development of UAV military field to coordinated attack of multi-UAV to accomplish specific strike tasks. Aiming at the problem of coordinated attack of multi-UAV, a typical confrontation scenario is constructed. The unmanned aerial vehicle cooperative attack problem is modeled as a decentralized partially observable Markov decision process (Dec-POMDP), and a unique reward function is designed. The multi-agent deep deterministic policy gradient (MADDPG) algorithm is used to train the attack strategy. Monte Carlo method is used to analyze the simulation experiment, and the results show that after the training of the multi-agent reinforcement learning algorithm, the completion rate of the UAV cooperative attack task reaches 82.9% in specific confrontation scenarios.

To ensure that the UAV (unmanned aerial vehicle) can accurately locate the reconnaissance target, image matching plays a critical role in the localization process, and its performance directly affects the localization accuracy of the targets. However, UAVs often encounter challenges when performing reconnaissance tasks in complex environments, especially when background information is similar, which hinders the effective elimination of a large number of outliers. This paper proposes an improved RANSAC (random sample consensus) algorithm. First, an initial dataset is constructed using triplet relationships. Based on these relationships, an improved strategy for selecting initial data is proposed, which reduces computational costs and improves matching accuracy through neighborhood-based geometric consistency and triangulation. Second, a data subset refinement strategy is introduced to further enhance the algorithm's sampling performance. Finally, the proposed algorithm is compared with other advanced algorithms on UAV reconnaissance image datasets featuring complex environments and similar background information. Experimental results demonstrate that the proposed algorithm achieves higher computational efficiency and correct matching rates, thus improving both the computational speed and localization accuracy, providing a novel method for UAV reconnaissance target localization.

In view of the problem that the infrared imaging for terminal guidance of rockets is affected by aerodynamic heating, the temperature response of the sapphire infraredwindow structure caused by aerodynamic heating is obtained through numerical simulation. On this basis, a wind tunnel test is designed to study the effect of aerodynamic heating on the infrared imaging of the seeker. In the numerical simulation, the aerodynamic heating of the infrared dome along the ballistic trajectory is obtained by combining computational fluid dynamics and engineering algorithms, and the temperature response of the dome structure is obtained by the finite element method. In the wind tunnel test, measuring points are installed on the inner surface of the dome to obtain the temperature response of the dome structure under aerodynamic heating, and the imaging effect of the infrared seeker under aerodynamic heating is tested. The research results show that, under aerodynamic heating, the highestinner wall temperature of the dome is 148 ℃, which is much higher than the target temperature of 27 ℃. Based on the sapphire window temperature, it needs to reduce the integration time of the infrared detector and adjust imaging algorithmto perform image correction.

In order to study the influence of the missile ejection condition in initial phaseon the deploying process of folding rudders, ADAMS is used to analyze the dynamic response of a torsion bar folding rudders without mechanical limit structure. The actual structural dynamic performance of the folding rudders is obtained through the observation of the folding rudders deploying test on the ground by mean of high-speed photography, and the model parameters of the dynamic simulation are modified according to the experimental results. Considering the impact of missile body attitude variations during aerial ejection on the folding rudders deploying process, the ejection force of the ejection device on the missile is measured by the ejection test on the ground, and CFD simulation is used to calculate the time-varying aerodynamic force experienced by the control surface of the rudders when they are deployed under the complex flow field of the plane belly, a dynamic deployment model for the rudders is established, which involving both ejection device force and aerodynamic resistance. The failure risks of folding rudders deployingt process are analyzed under multiple external folding rudder unlock time conditions and aerodynamic force deviations. Based on simulation results, a reasonable design range for external folding rudder unlock time is proposed. This work provides valuable reference for the engineering design and application of folding rudder.

In order to study the hot launch shock wave opening process of small launcher, a numerical simulation of the process is carried out by using CFD-FASTRAN aerodynamic analysis software. According to the calculation results analysis, it gives the changing process of surface temperature and pressure of main components affects by gas shock wave in the process of hot launch shock wave opening of small launch tube. The whole process of hot launch shock cap opening is analyzed by using CFD-FASTRAN pneumatic analysis software innovatively, and the influence of the distance between nozzle and back cap and the expansion angle of nozzle on the shock cap opening process is systematically analyzed for the first time. The results show that the pressure decreases from edge to center and the temperature increases from edge to center under the influence of gas jet. The shock wave reflected by the gas flow moves along the axis of the launch tube towards the warhead, and the pressure and temperature on the wall of the launch tube gradually decrease with the direction of the shock wave. The temperature and pressure of the front cover are weakened from edge to center by the gas shock wave influence. The increase of the distance between the nozzle outlet and the rear cover and the increase of the expansion angle of the nozzle can enhance the total energy reflected by the shock wave to the front cover, and shorten the opening time of the shock wave. The calculation and simulation method can provide a reference for the strength design and structure design of the front and back covers of the small launch tube with hot launch shock wave.

Aiming to address the issues of low lift-to-drag ratio during the gliding phase, poor maneuverability in the passive flight segment, and limited range inherent in traditional canard-layout field rockets, this study analyzes the aerodynamic characteristics of canard-layout rockets. Based on these findings, a novel aerodynamic layout design scheme is proposed for high-performance long-range rockets with a large aspect ratio. Simulation results indicate that at an 8° angle of attack, the aerodynamic efficiency of the canard control surfaces decreases sharply with increasing Mach number, while their contribution to the projectile's lift force increases significantly. When Ma ranges from 4 to 6, the proportion of lift generated by the control surfaces is less than the drag they produce, resulting in negative aerodynamic gains for the entire projectile. In contrast, the body lift accounts for 61%~67% of the total lift. Based on these insights, a tail-controlled high-performance rocket layout with a large aspect ratio is proposed. This design reduces the number of inefficient control/wing surfaces during high speed flight and enhances balance and maneuverability through optimized rear-positioned control surfaces. Numerical simulation results confirm that the tail-control layout substantially improves the rocket's lift-to-drag ratio, provides a larger usable angle of attack, and enhances overall maneuverability, making it a more suitable configuration for long-range rockets in hypersonic flight.

To address the issue of poor impact accuracy in elevated firing operations of helicopter-mounted aerial rockets, an improved method for elevated firing has been proposed through analyzing the combat mechanisms involved. A simulation model for the ballistic trajectory and the damage effectiveness has been established based on the publicly available parameters of a certain armed helicopter and its aerial rockets from abroad. Typical conditions of elevated and level flight firings are selected for comparative analysis of range, dispersion, and damage effectiveness. The simulation results indicate that the improved aerial rockets, when fired in an elevated position, can achieve an attack distance of up to 9 km, while also exhibiting a trend of reduced longitudinal dispersion. The effective damage area of the rocket salvo is 3 times that under level flight conditions. The improved elevated firing method and the suggestions for the fire control and the aiming stabilization optimization mentioned in this paper have certain guiding significance for the application of helicopter elevated firing operations.

It is difficult to accurately test overload of a projectile moving in a bore by conventional testing methods and the strong electromagnetic environment interferes with the testing devices in electromagnetic launching, a new method for testing the overload of the projectile in the bore is proposed. Firstly, based on the technical advantage of electromagnetic launching technology that the overload in the bore is adjustable and controllable, a testing principle for projectile overload test is put forward, that controls the current waveform and makes it into a flat wave, then establishes a constant overload loading interval to test projectile overload. Secondly, a new test method of the projectile overload, that is launched by electromagnetic rail, is established. By controlling the current loading waveform and changing the current amplitude, the corresponding relationship between the overload and the current is established, and through electromagnetic launching test, it can get the loading current and corresponding relationship and indirectly get overload of the projectile in-bore. Finally, the electromagnetic launching test is carried out, and the conventional test method and the new test method are compared. The results show that the new test method can complete the overload loading process test of the whole inner bore, and can more truly reflect the process of the projectile starting to move after overcoming the static friction. This test method can meet the requirements of in-bore overload test of electromagnetic projectile, and avoid the influence of electromagnetic interference on the test.

In response to the limitations of traditional prescribed performance in controller design, this paper studies a non singular terminal sliding mode fault-tolerant control method that improves prescribed performance. Aiming at the problem of control singularity caused by the convergence speed of the performance boundary of traditional prescribed performance being too fast and the tracking error exceeding the performance boundary, an improved performance function and an improved error conversion function are proposed. The converted error is introduced into the design of a non singular terminal sliding fault-tolerant controller. Simultaneously utilizing an extended state observer to accurately estimate unknown external disturbances in the system, and compensating them to the controller to solve the problem of partial failure of the actuator and suppress elastic effects, in order to obtain preset transient and asymptotic steady-state performance. The stability is proved through Lyapunov theory analysis. Finally, the effectiveness and superiority of the proposed fault-tolerant control method are verified through simulation and comparative experiments of the attitude system of air breathing hypersonic vehicle.

During the hot launch process of a carrier rocket, the high-temperature and high-pressure gas generated by the rocket engine will cause significant high-temperature erosion on flame division trough. In order to achieve the thermal protection of the division trough during launch of a rocket, the variation of the flow field during water injection onto the surface of a single-sided division trough was analyzed based on the computational fluid dynamics (CFD) method with coupled mixture multiphase flow model and Lee model for different water injection velocities for a rocket with two booster stages. The results show that spraying water into the division trough effectively suppresses the phenomenon of gas splashing, significantly reduces the distribution area of the high-temperature zone on the surface of the division trough, and provides good protection for the division trough. When the water spraying speed is low, it will cause oscillation of the maximum temperature on the surface of the division trough. As the water spraying speed increases, the degree of oscillation of the maximum temperature on the surface of the division trough gradually weakens until it disappears. As the water spraying speed increases, the maximum temperature on the surface of the division trough decreases, and the maximum vaporization rate also increases. Spraying water onto the gas jet will change the shape of the flow field. Different spraying positions will result in different distributions of physical parameters along the axis of the gas jet. Directly impacting the gas jet with the water jet will have a better cooling effect. This conclusion can provide some reference for the launch process of carrier rockets.

Rolling linear guides, known for their precise guidance and dynamic stability, are widely used in rocket and missile weapon launch transmission systems. Due to differences in functional adaptability, multiple sets of rolling linear guides are involved in the same launch system. To achieve efficient modeling, simplify the design process, and accelerate the product design iteration cycle for different purposes and models of rolling linear guides, this paper proposes a dual-drive parametric design system that integrates dynamic and static characteristic simulation analysis results with the engineering design experience of the engineers. The system extracts key structural parameters that affect the performance of the components, redevelops the Abaqus finite element software using the Python language and take multiple models of rolling linear guides as examples to establish a parametric modeling method for the rail pair. The results of typical test cases show that, based on the dual-drive parametric design system and using the Python-Abaqus parametric modeling method, a parametric model with adjustable geometric parameters can be established. This method creates a three-dimensional model of the rail pair with a geometric error of less than 0.3% in under 4 seconds.

The gas jet generated during rocket launching has a strong impact on the launching platform and may cause great damage to the launching platform. Therefore, it is necessary to study the impact effect of gas jet, researching the the wake field of rocket under different elevation angles. Based on computational fluid dynamics method, a finite volume method is used to discrete the rocket wake fields, and mathematical physical models of wake fields at different elevation angles are established for numerical simulation. Within the range of high and low angle adjustment, five angles of 0°, 5°, 15°, 28° and 38° are selected for simulation. By analyzing the distribution of jet velocity and pressure with the change of elevation angles and time, the development law of jet under different elevation angles can be obtained. Based on the analysis of the wall pressure distribution and maximum pressure variation, the impact effect of jet on the launching platform under different elevation angles can be obtained. The research results show that the wall pressure tends to increase with the increase of the elevation angle. And when the elevation angle is greater than 15°, the wall pressure will increase greatly. Each position of the launch platform is affected by jet impact to different degrees, and different protection strategies need to be adopted. The research provides theoretical support for the launch platform protection design.

The rapid separation characteristic for a sabot from its integrated launch projectile has an important impact on the projectile firing accuracy, and is one of the main design requirements about sabot. In order to improve the separation performance of the sabot, an unsteady CFD method based on six degrees of freedom equation coupled with URANS equation is created for the separation calculation of the sabot, and a surrogate optimization framework for the shape of the windward nest of the sabot is constructed based on Kriging model and genetic algorithm. A 16 degree compression corner example is used to verify the reliability of the CFD method by comparing with the experimental values. The shape of the optimized windward nest is obtained by surrogate optimization method, and the flow field and separation characteristics of the optimized sabot are compared with the initial sabot, which reveals the aerodynamic mechanism of the improved separation performance of the optimized sabot. The results show that the decrease of the pressure on the inner surface of the optimized sabot is significantly less than that of the initial sabot during the separation process, while the pressure on windward nest of the optimized sabot is basically the same as that of the initial sabot. Therefore, the overall separation force of the optimized sabot is significantly increased, with the transverse separation displacement increases by 14.86% and separation pitching angle increases by 13.75%, and the separation performance is improved.

Cold launching of rocket is a kind of launching method by means of the gas generator in the launching tube to catapult the rocket out of the launching tube, and then ignite the main engine after the projectile body leaves the tube. In order to ensure that the rocket can get enough muzzle velocity at the end of the cold launching process, a structure of the cold launcher is designed and improved, and the internal ballistic characteristics during the launching process are studied, then a two-dimensional axisymmetric model of rocket cold launcher is established. It adopts layer spread dynamic grid technology and fluid control equation to numerically simulate the internal flow field of the launcher under the same propellant quantity and different working time, and obtain the pressure inside the launcher and the motion characteristic curves of the projectile body under different working conditions. The simulation results show that under the condition of the same mass of the propellant column, the increase of the working time of the gas generator will lead to the extension of the rocket discharge time and the decrease of the discharge speed, but it will also reduce the peak pressure and acceleration in the launching tube, the rate of change in the acceleration of the rocket is more stable, and the rocket overload will be reduced during the launching process.

To address the challenge of effectively analyzing the dynamic response of structures under high-speed impact, this paper proposes an adaptive sensitivity analysis method based on the Material point method (MPM) and sparse solynomial chaos expansion (sPCE) to address uncertainties in structural response. First, MPM is used to simulate the high-speed impact process of projectiles, revealing the evolution pattern of structural strain energy over time. Second, a surrogate model is developed using the sPCE method to analyze key parameters influencing the dynamic response of the structure and to assess the sensitivity of each parameter to structural behavior. Finally, the proposed method is validated through numerical experiments on projectile impact and penetration scenarios. The results indicate that the intrinsic material properties play a dominant role in structural response, and the coupling effects of impact angle and velocity exhibit strong nonlinear characteristics under specific conditions. The findings provide important theoretical support for optimizing structural design under high-speed impact and demonstrate the accuracy and practical value of sPCE in sensitivity analysis.

Deep debonding integrated propellant column is a high loading grain suitable for small opening & non-segmented shell single chamber dual thrust engines. In order to conduct a safety assessment of the engine in low-temperature ignition, this article conducted simulation analysis on the integrity of the low-temperature structure of the propellant column based on a finite element model, and studied the influence of the pressure difference between the inner hole and gap of the column on the integrity of the charge structure during low-temperature ignition process. The pressure distribution inside the gap during the ignition process was tested by a simulated engine. The results indicate that the pressure difference between the inside and outside of the propellant has a significant impact on the stress and strain distribution of the grain during low-temperature ignition process. Through experiments, it was found that under low-temperature ignition conditions, the pressure building rate at the tail of the gap is synchronized with the combustion chamber whose max pressure is only 4% lower, while the pressure building in the middle of the gap is delayed obviously, and the maximum pressure difference is about 64% of the combustion chamber. The calculated comprehensive safety factor of the column at low temperature reaches 1.76 indicating high safety and reliability of low-temperature ignition.

To optimize the aerodynamic shape of small-caliber super-high-speed armor-piercing projectiles, a research on the influence of the shape parameters of the armor-piercing projectiles on the aerodynamic characteristics and flight stability was conducted. Based on three key factors, namely the sweepback angle of the tail fins, the aspect ratio, and the overall length-to-diameter ratio of the entire projectile, multiple improvement schemes for the aerodynamic shape were put forward. A numerical calculation model of the hypersonic flow field was established by using Fluent to explore the variation patterns of the aerodynamic parameters and flight stability. Through comparative analysis, the superior aerodynamic shape scheme was acquired. The results suggest that the zero-lift drag coefficient is reduced by 5.3%~26.9% after the improvement, and the lift coefficient and static moment coefficient are better than those of the original projectile. The static stability reserve amounts to 11%~29%, and the logarithmic attenuation rate of the amplitude ε equals 34.39%, meeting the stability requirements of tail-fin-stabilized armor-piercing projectiles. Meanwhile, the firing altitude, velocity drop, and oblique range at each firing angle have witnessed significant improvements compared to the original projectile. This offers a certain reference for the shape optimization design of small-caliber super-high-speed sabot-armor-piercing projectiles.

For the tube-launched missiles carried on air based platforms, the launchers equipped with traditional flat sealing film lead to unavoidable issues that aerodynamic drag increases and broken fairing threatens the safety of aircraft. To address the problem, the influences of the aspect ratio of ellipsoidal fairing and aircraft cruising speed on its aerodynamic characteristics are studied. The relationship between fairing aerodynamic moment and increment of rotation angle is established. Based on the model of twins torsional spring, a fairing mechanism with tandem torsional springs is designed. The analytical model for the contact force of fairing adapter and its contact point parameters is established to obtain the optimal contact point and the minimum contact force. As a result, a multifunctional fairing mechanism is proposed which is capable of reducing aerodynamic drag, nondestructively opening, reducing contact force and automatically resetting itself. The analysis results of the virtual prototype indicate that the dynamic contact force of the adapter and the reset time are respectively 224.0 N and 17.7 ms, it verifies the rationality and feasibility of the mechanism. The proposed method has important reference significance for the design of launch tube fairing of tube-launched missiles equipped on aircraft platform.

In view of the technical difficulties such as unstable and target loss in the process of tracking targets of the image seeker, this paper combines the advantages of reliable and stable control of “people in the loop” in manual tracking mode, and the advantages of automated information processing and simple shooter operation in automatic tracking mode, and innovatively proposes an anti-interference stable tracking fusion design method for image seeker. In the automatic tracking mode of seeker, when occasional abnormalities such as target tracking failure or target loss occur, the tracking point is slightly corrected through manual adjustment, which not only reduces the operation pressure of the shooter, but also maximizes the reliability and anti-interference ability of automatic tracking under complex backgrounds, bad weather and other conditions. At the same time, based on the principles of intervention continuity, correction agility and automatic tracking stationary, the optimal solution for micro-revision interval time and correction cycle is designed using the analytic hierarchy process method. The semi-physical simulation results show that the micro-repair preferred solution can achieve a missile hit rate of 99% under different simulation conditions.

In order to study the spinning projectile stability of coning motion,provide the dynamic equation around the center of mass in quasi-body coordinate system and establish the short period dynamic model.Derive the equations of coning motion represented by Euler angles.Obtain the analytical solution by introducing complex angle of attack.Analyze the stability of coning motion with initial disturbance in two cases:One situation is ignoring the Magnus effect and damping effect,only analyze the coning motion influenced by the gyroscope effect and the aerodynamic static moment.Another situation is the coning motion adding in the Magnus effect and damping effect,analyze the stabilization and provide the conditions for coning stability.Finally,summarize the influence of rotating speed and statically stability acting on the dynamic stability of coning motion based on theoretical analysis and simulation results:For the low-speed rotating missiles,only under the influence of gyroscopic effect and aerodynamic static moment,a statically stable aerodynamic shape is necessary to achieve coning stability.Statically unstable missiles can only achieve dynamic stability by significantly increasing the rotating speed.Considering the Magnus effect and damping effect,the convergent coning motion of the statically stable missiles with low rotating speed is the easiest to achieve.The divergence of coning motion may occur result in the increasing of rotating speed.For the statically unstable missiles,it is necessary to select the moment of inertia reasonably and increase the rotating speed to achieve dynamic stability.

The three-axis seeker of loitering missile faces the technical problems of single optical axis pointing mapping multiple frame angle combinations and fast real-time calculation of optimal instructions when tracking the target. Based on the kinematics analysis of the three-axis seeker, an adaptive fast tracking strategy for the three-axis seeker is proposed. The influence of the combination of three-axis bandwidth and three-axis increment on the tracking speed of the seeker is analyzed, and the fitness function evaluation standard with frame adaptability is established. The simulation results show that the proposed strategy can solve the optimal tracking instruction under the framework constraint, solve the problem that the tracking instruction cannot be tracked due to exceeding the mechanical limit, and realize the tracking target of the control optical axis along the optimal path. The average iteration time is 3.54 ms, and the tracking speed is improved by 30.4%.

The super caliber canard-controlled projectiles have a large front diameter,a forward center of gravity,a variable diameter cross section in the middle of the projectile body,and more complex surface gas flow.In order to study its aerodynamic characteristics,the lifting resistance characteristics,pitching moment characteristics and the surface flow field distribution of the projectile were analyzed by numerical calculation.The results show that the lift-drag ratio increases first and then decreases with the increase of attack angle,and the lift-drag ratio reaches the maximum when the attack angle is 8°.The pitch moment provided by different parts of the projectile body is studied in different areas.It is found that the forward pitch moment provided by the projectile tail takes up more than 30% of the total pitch moment when the attack angle α=0°,elevator angle δ_z=5°,and the stability ratio of the δ_z=10° decreases by 46.8% compared with δ_z=5°.Vortices appear at the variable diameter of the projectile body,which interact with the vortices of the rudder surface and the vortices around the surface of the projectile body,making the gas flow on the surface of the projectile body more complicated and increasing the difficulty of control.

In order to address the problem of difficulty in determining parameters for the Weibull reliability identification test for complex electromechanical products, a Weibull distributed reliability qualification test method based on historical zero-failure data is proposed.The prior information of failure rate is calculated based on the two-parameter Weibull model and historical data, the posterior information of failure rate is obtained by Bayesian statistical theory, and the parameter estimation of Weibull distribution is obtained by least square method, and the reliability qualification test scheme is calculated. The reliability qualification test scheme is designed and compared with the scheme in GJB 899A—2009 through taking a viewing device as an example in this paper. The reliability assessment result of the Weibull distribution for reliability identification test is 9.2% more accurate than that of the exponential distribution. The established Weibull distribution reliability qualification test scheme based on zero-failure data is effective. The overall Weibull distribution test plan is stricter than the exponential distribution test plan, and this method is highly suitable for the design of reliability qualification test plans for electromechanical products. The established Weibull distribution reliability qualification test plan based on no-failure data is effective and can be directly applied to the formulation of reliability qualification test plans for complex electromechanical products.

Transonic shock oscillations and cavity flows are both primary contributors to the intense pulsating pressures and structural vibrations in laterally maneuverable rocket missiles. However, existing studies have predominantly focused on single flow characteristics impacting on the wall load environment, while neglecting the coupling effects between two unsteady flow phenomena. This oversight poses potential risks to the flight safety of aerospace vehicles under specific flight conditions. In this study, a rocket fairing and a sidewall cavity structure are employed as the research subjects. The Delayed Detached Eddy Simulation (DDES) method is utilized to conduct a numerical investigation into the flow field characteristics and the time-frequency properties of pulsating pressures in the presence of both transonic shock oscillations and cavity flows, aiming to explore the coupling mechanisms between shock oscillations and cavity flows. The research findings indicate that cavity flows reduce the effect of separated flows on the shock, thereby weakening the intensity of transonic shock oscillations, while shock oscillations stabilize the cavity flow field, significantly lowering the sound pressure level (SPL) on the cavity walls, maximally, which can reach up to 30 dB at Mach 0.95. Based on the flow dynamics cognition in coupled transonic flow fields, 3 flow control strategies for the coupled flow field are designed. Comparative results reveal that a steep rise in leading edge of the cavity promotes the mutual inhibition of shock oscillations and cavity flows, thereby reducing the pulsating pressure levels on the cavity walls. The analysis of the coupled flow field involving transonic shock oscillations and cavity flows provides valuable support for the development of next-generation aerospace vehicles.

A three-dimensional simulation model is established to investigate the impact of engine gas flow on the fragile front cover of surrounding launch tube during the process of a missile leaving its launch tube matrix. The dynamic grid method is used to simulate the motion of the missile body, and the transient flow field on the surface of the surrounding front cover is numerically simulated and calculated. By analyzing the gas streamline and flow field contours, the influence of the engine expansion and compression wave on the airflow direction and pressure distribution on the surface of the front cover during missile motion is obtained. Based on the changes in pressure difference between the inside and outside of the front cover at different times, combined with the safe range of pressure difference obtained from experiments, it assists in determining whether the front cover will rupture prematurely. By changing the distance between adjacent launch tubes, the magnitude of the change in pressure difference between the inside and outside of the front cover is calculated, and the safe distance between launch tubes is obtained. This provides simulation prediction and data support for the structural optimization design of the front cover and the arrangement scheme of multiple launch tube units.

A solid rocket motor-boosted uncontrolled target missile is launched through ground ignition method, which is implemented by a certain type of launch control equipment. In this paper, the fault tree analysis (FTA) is applied to diagnose and localize anomalies in time synchronization signal transmission occurred during the system testing. Firstly, the operational principles of the launch control equipment are introduced, which comprising an ignition controller, a battery box, a safety control box, a laptop computer, a target-side radio set and a command-side radio set. As the same time, the symptoms of time synchronization signal anomalies are also described. Subsequently, a fault tree model is constructed based on the signal transmission chain architecture in the system. Through the qualitative analysis of basic events and step-by-step verification procedures, the root cause is identified as a write conflict in time synchronization signal logic outputted from the CPU software of the ignition controller. And this conflict is found to reduce the pulse widths below 101 ms under specific triggering conditions, that do not meet the signal input requirements from the target-side radio set. The failure mechanism is rigorously analyzed, and the anomaly is successfully reoccurred by experiments. Design optimizations are implemented to resolve the software conflict, and test results confirm the effectiveness of these modifications, which demonstrating enhanced system reliability.

To address the practical challenges of nonlinear dynamic modeling in missile hydraulic erection systems, as well as the presence of mechanical and hydraulic dynamic uncertainties, this study proposes an incremental nonlinear dynamic inversion (INDI) control method. The proposed approach reduces the reliance on complex nonlinear hydraulic models, offers strong robustness, and enables a concise and efficient controller design with a clear structure. First, a dynamic model of the valve-controlled cylinder-driven system is established, and a virtual control law for the hydraulic channel is constructed using the backstepping method. Then, a first-order Taylor expansion is applied to decouple the nonlinear hydraulic dynamics, based on which an INDI controller is designed to track the virtual control law and achieve accurate missile erection trajectory tracking. The stability of the closed-loop system is proven using the Lyapunov theory. Comparative simulations provide further evidence supporting the effectiveness of the proposed controller.

In view of the fact that the existing research has not systematically revealed the scientific issue of the dynamic response of the unanchored launch system to road damage, this study based on the dynamic model of the underground cavity-damaged road and the unanchored mobile launch system, analyzes the influence of the spatial distribution of cavities on the coupled system. The results show that the cavity damage causes the road to undergo brittle fracture, and the maximum deflection of the road increases by 23.5%. However, the initial disturbance of the missile decreases by 21.5% due to the road damage. When the cavity shifts longitudinally, the maximum deflection of the road first increases by 3.36% and then gradually decreases. When the cavity shifts towards the rear of the vehicle, the initial disturbance of the missile first increases by 42.1% and then gradually decreases. When the cavity shifts towards the front of the vehicle, the initial disturbance of the missile monotonically decreases by 28.95%. When the cavity shifts laterally, the maximum deflection of the road first increases by 14.33% and then decreases by 22.76%, and the yaw rate of the missile increases from 0.001°/s to -0.091°/s and then decreases to -0.049°/s. The depth parameter of the cavity has a relatively small influence on the coupled system. With the increase in the size of the cavity, the maximum deflection of the road increases by 47.65%, while the initial disturbance of the missile first decreases by 21.8% and then gradually increases due to the road damage.