In order to investigate the blast resistance of a 316L stainless steel honeycomb sandwich structure, a honeycomb sandwich structure was designed and fabricated using 316L stainless steel powder by selective laser melting (SLM). Concurrently, solid panels of equivalent surface density were produced by this method and constituted the control group. The mechanical behavior of the structure under near-field static explosion load is obtained through static explosion experiments and LS-DYNA simulation experiments, and the propagation mode of the stress wave within it is investigated in order to elucidate the underlying anti-explosion principle. Moreover, optistruct is utilized to optimize the topology and structure of the structure, with the objective of enhancing its blast resistance. The findings indicate that the backplate deflection of the porous sandwich structure is diminished by 13.2% in comparison to that of the plate with isoplanar density, thereby enhancing blast resistance. The established numerical model of fluid-solid coupling is capable of describing the three phases of the static explosion experiment, namely the shock wave propagation phase, the fluid-solid coupling phase, and the inertia phase. The explosion experiment yielded definitive results at the center of the target plate, thereby demonstrating that the "川" crack is caused by residual core layer extrusion. Moreover, the core layer deformation failure mechanism for the honeycomb panel was observed to manifest as in-plane stretching and tearing. The optistruct optimization results demonstrate the formation of a triangular skeleton and circular holes, alternating with corrugated plates. The structure, optimized for a corrugated core target plate, displays enhanced resilience in comparison to the optimization of a traditional honeycomb sandwich panel. The explosion load backboard deflection exhibited a 25.4% reduction, the peak pressure behind the plate demonstrated a 17.6% reduction, and the blast resistance was significantly enhanced. In comparison to honeycomb panels, the circular hole structure has been demonstrated to reduce the backplane deflection by 38.1%, while the triangular hole structure has been shown to reduce the peak pressure behind the plate by 22.4%.

The active sidestick is a crucial control device that can adjust the flight attitude of an aircraft. In the automatic flight mode, the active sidestick follow-up function tracks the control commands from the flight controller in real-time to enhance the pilot's perception of the aircraft's flight status, thereby effectively improving flight safety. However, unknown disturbance torques in the system can degrade the tracking accuracy of the active sidestick. To address this challenge, a control method based on composite nonlinear feedback and adaptive integral sliding mode is proposed for the follow-up control of the active sidestick. This method integrates an integral sliding mode control algorithm and the composite nonlinear feedback control algorithm, where the composite nonlinear feedback control effectively improves the steady-state performance of the active sidestick system, and the adaptive integral sliding mode control effectively restrains unknown system disturbances. Extensive experimental results show that the proposed improved active sidestick follow-up control algorithm can reduce the adjustment time by approximately 41% and the steady-state error by about 93.6% compared to the previous composite nonlinear feedback control algorithm. Furthermore, under disturbance conditions, the improved follow-up control algorithm can reduce the adjustment time by about 79% compared to the traditional PID algorithm. This demonstrates that the designed improved active sidestick follow-up control algorithm can significantly enhance the system's control accuracy and response performance, and possess excellent disturbance rejection capabilities.

The quality of explosive charge is a key factor affecting the performance of artillery weapons, and the explosive charge is often a mixture of powders. To improve the quality of cylinders made by a mixture of metal and non-metal powders, the mechanical behavior of the cylinders was described using continuum plasticity theory, the Shima-Oyane model and the Ogden model were employed as the material constitutive models for the cylinders and the rubber sleeve, respectively. And a simulation model for the isostatic pressing of cylinders was developed utilizing the nonlinear finite element software MSC.Marc. Based on the simulation model, the forming mechanism of cylinders was explored, and a comparative study was conducted on the influence of isostatic pressing process parameters on the forming quality of the cylinders. The results indicated that the simulation model could effectively reflect the forming characteristics of the cylinders. The maximum pressure and it’s holding time of isostatic pressing were the key factors that influenced the quality of the cylinders. When the pressure was set at 240 MPa and the holding time exceeded 400 s, the overall relative density of the cylinders reached above 97%, with the density distribution difference was below 0.5%. The results of the isostatic pressing experiment verified the accuracy of the simulation analysis results. The cylinders with approximate length-to-diameter ratios of 5∶1 achieved a higher process standard and satisfied the process requirements.

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.

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.

During the transfer alignment under moving base, the flexible deformation between the master and slave inertial navigation systems (INSs) is the primary error factor affecting the alignment accuracy of the slave INS. In traditional transfer alignment methods, flexible deformation is usually equated to empirical Markov models. However, a low match between actual flexible deformation and the empirical model can lead to a decrease in transfer alignment accuracy. Therefore, a high-accuracy transfer alignment method without relying on empirical model of flexible deformation is proposed. Firstly, the intrinsic relationship between the flexible deformation and the angular velocities measured by master/slave gyroscopes is derived and established. Thus, the rough value of flexible deformation is directly calculated using the measured angular velocities. Then, the coupling relationship between the calculation error of flexible deformation and the gyroscope error is derived, and a novel transfer alignment system model that does not rely on empirical models of flexible deformation is established. Furthermore, the optimal estimation algorithm is used to accurately estimate and correct the calculation error of flexible deformation. Simulation results show that the proposed method can accurately compensate the complex flexible deformation, thereby achieving high-accuracy transfer alignment under moving base.

This study commences with an examination of the operational sequence of the secondary fuze in a fuel-air explosive (FAE) munition, analyzing the potential interference factors such as shock waves and electromagnetic radiation, as well as their characteristics, that the secondary fuze may encounter at the terminal phase of the trajectory. Counter-interference design methods for the secondary fuze are proposed from various perspectives, including trajectory design, circuit design, and structural design. Static detonation experiments were conducted to compare the interference-resistance capabilities of the secondary fuze under two different configurations: independent design and cable-connected design. The experimental results indicate that the secondary fuze with an independent structural design exhibits superior interference-resistance, capable of emitting a normal ignition signal at the preset timing following the explosion of the dispersed charge. In contrast, the secondary fuze with a cable-connected design failed to issue an ignition command at the designated moment, resulting in electrical failure of the components within the fuze. It is postulated that the electromagnetic radiation generated by the detonation of the explosive charge enters the fuze through field-line coupling, thereby causing damage to the fuze. Furthermore, the experimental results demonstrate that under the conditions of this study, the pulse interference current induced has a duration reaching the order of hundreds of nanoseconds, with a peak current reaching tens of amperes, which is sufficient to directly inflict damage on the microcontroller and interface circuits of the fuze.

In order to provide reference for the analysis on exterior ballistic environment of fuze, FLUENT software is applied to numerically simulate the aerodynamic characteristics of a large caliber dynamic imbalance spin-stabilized projectile with base cavity, and corresponding aerodynamic characteristics of the whole trajectory are obtained. The function relationship of aerodynamic characteristics, dynamic imbalance angle with Mach number of the projectile is fitted 1stOpt software. The influence exerted by dynamic imbalance angle on drag coefficient is small and within 2% lift coefficient, overturning moment coefficient and polar damping moment coefficient are positively linearly correlated with dynamic imbalance angle of the projectile while Magnus force coefficient and Magnus moment coefficient are negatively linearly correlated with dynamic imbalance angle of the projectile. When dynamic imbalance angle is large (such as 1°), polar damping moment coefficient will increase 2~4 times compared with that dynamic imbalance angle is zero.

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 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.

Accurately and quickly detecting military targets in complex scenarios has important military value in perceiving battlefield situations, conducting reconnaissance and early warning analysis, and providing precise missile guidance. A multi-scale object detection algorithm AEM-YOLOV5 (AFPN-EMA-MPDIoU-YOLOV5) based on improvements to YOLOV5s is proposed to address the issues of poor feature learning ability, low detection accuracy, and high computational complexity in existing algorithms. Firstly, the AFPN asymptotic feature pyramid network is introduced into the neck network to gradually fuse the detailed information at the bottom of the image and the high-level semantic features at the top, enhancing the network feature fusion effect. Secondly, an EMA attention mechanism module is added before each detection branch to aggregate pixel level features across spaces, improving the level of attention to multi-scale targets in complex scenes. Finally, MPDIoU is used to replace the original C_IoU bounding box loss function in YOLOV5, solving the problem of C_IoU degradation when the predicted box aspect ratio is the same but the absolute value is different, making the regression results more accurate. The experimental results show that the improved algorithm performs well on the RSOD dataset, P_mAP50 reaches 94.5%, FPS reaches 42 frame/s, and model size is 14.8 MB. Compared with existing algorithms, the improved algorithm significantly improves its performance, meets the real-time requirements of military target detection, and the model is lightweight.

The temperature of the rocket control cabin is predicted by aerodynamic heating and structural thermal response calculation, and verified by flight test. Firstly, the ballistic feature points are selected and transition is judged. The aerodynamic thermal environment of the rocket flight is obtained by the combination of computational fluid dynamics and engineering calculation methods. Then, based on the finite difference method, the temperature response of the heat protection structure at the rocket control cabin is obtained. Finally, the numerical calculation results are compared with the flight test data. The numerically predicted maximum temperature of the inner wall of the rocket control cabin is 5.6% higher than the flight test value, and the numerical prediction method in this paper can be used in the thermal protection design of the rocket.

To study the influence of the composite microburst wind field on the cruise phase of a light and small fixed-wing unmanned aerial vehicle (UAV), in this paper, a Dryden atmospheric turbulence model and a microburst model are constructed, and after fusing them, a composite microburst wind field model is created. Take a specific electric-powered UAV as the object of study and carry out a six-degree-of-freedom rigid body ballistic model simulation analysis, the simulation results show that the newly established composite wind field model has better randomness and typical wind shear characteristics and it can effectively depict the actual composite microburst wind field distribution and change. Under the influence of this composite wind field, the UAV has experienced a significant loss of height, and when the central induced velocity is between 10~25 m per second, the loss of height reached 168~537 m. Furthermore, the flight parameters of UAV, including flight duration, angle of attack, sideslip angle, and power margin, have changed in different degree.

In order to further enhance the damage capability of the explosively formed penetrator (EFP) warhead, a combination of theoretical analysis, numerical simulation, and experimental verification was employed to design a novel explosively formed arrow-shaped penetrator warhead. The formation effect of the explosively formed arrow-shaped penetrator was controlled by adjusting the number of linearly arranged explosive liners, thereby increasing the damage power of the damage element. Initially, theoretical analysis was conducted to establish a physical model, determining the number of liner arrangements. Subsequently, numerical simulations were performed to validate the feasibility of the designed scheme. A warhead with 12 columns of linearly arranged liners was then fabricated and subjected to static detonation experiments. The experimental results demonstrated that a single liner could form two explosively formed arrow-shaped penetrators, capable of penetrating a 30 mm-thick Q235 steel plate at a distance of 5 m. The research findings indicate that the principle of forming explosively formed arrow-shaped penetrators from linear liners is viable and holds significant implications for the design of EFP warheads.

In order to investigate the influence of liner structure parameters on the formation of jet penetration charge (JPC) as well as the penetration ability and post-effect action of the formed JPC on ceramic composite target plates, this study adopts orthogonal experimental optimization to design the structure parameters of an equal-wall-thickness conical liner. The LS-DYNA software is utilized to conduct research on how warhead structure parameters (cone angle, thickness, length-to-diameter ratio) affect the characteristic parameters of JPC damage elements. Through range analysis, the combination of structure parameters that results in better JPC formation performance is obtained. Based on the optimized structure, the damage situation of JPC damage elements on composite targets is studied. The results showed that when the liner cone angle is 90°, the wall thickness is 2.5 mm, and the charge length-to-diameter ratio is 0.9, the penetration effect is better. At this time, the funnel pit diameter increased by 8.93%, the funnel pit depth increased by 3.7%, the average diameter of the hole increased by 8.29%, and the total penetration thickness increased by 9.88% after removing the target plate spacing. This research can provide a theoretical and technical basis for the design of rod-shaped jet liner structure parameters and for studying the penetration ability on ceramic composite target plates.

This article proposes a heading calculation and integrated navigation algorithm based on continuous matching, which realizes the use of small field of view guided cameras to calculate the heading and provides a low-cost heading measurement and integrated navigation solution for unmanned aerial vehicles. The algorithm first calculates the attitude increment based on the essential matrix obtained from continuous matching, and further calculates the heading angle of the current frame through the first frame attitude matrix and joint calibration matrix. Based on this, a fusion algorithm for MEMS navigation solution and visual continuous matching is designed based on inertial navigation attitude filtering. To verify the accuracy and computational efficiency of the algorithm proposed in this paper, an inertial/visual combination device and a captive flight experimental system are constructed. The flight experiment results show that for cameras with a pixel count of 1 920×1 080, the matching success rate reaches 99.6% during flight at an altitude of 80~200 m. The heading calculation accuracy is better than 0.21° and the update frequency is better than 20 Hz during half hour navigation. Compared with traditional methods, it has a higher matching success rate, heading calculation accuracy, and lower time consumption.

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.

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.

To investigate the damage effectiveness of low, slow and small Unmanned Aerial Vehicles(UAVs) under fragmentation strikes, an evaluation model is established to comprehensively assess the damage effectiveness of fragments against UAVs. Numerical simulations are conducted to analyze the damage probability of a single fragment impacting various UAV compartments under different initial velocities, diameters, and material conditions, followed by extensive shooting simulations using the Monte Carlo method, with thousands of iterations to ensure statistical robustness. The results indicate that the initial velocity, diameter, and material of the fragment significantly influence the damage probability. An increase in initial velocity and diameter notably enhances the damage probability; however, the effectiveness tends to saturate within the velocity range of 900~1 500 m/s and diameter range of 4~8 mm, showing diminishing returns beyond these thresholds. The use of multiple fragments significantly increases the damage probability, suggesting that an increase in the number of fragments leads to a substantial improvement in damage effectiveness, making the quantity of fragments a crucial factor in determining the overall damage outcomes. Based on these findings, selecting fragment diameters between 2 mm and 4 mm, initial velocities between 600 m/s and 900 m/s, and appropriate materials can optimize the design of fragmentation warheads, thereby significantly enhancing their damage effectiveness against UAVs. This study provides a scientific basis for the precise evaluation of UAV damage effectiveness, offering valuable insights for the development of more effective countermeasures and supporting informed decision-making in combat scenarios to ensure enhanced operational effectiveness and strategic advantages.

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.

A sliding mode robust control strategy based on an adaptive first-order sliding mode disturbance estimator and a two-phase combined function reaching law is proposed for the control problem of uncertain quadrotor aircraft system. The quadrotor aircraft system is divided into two fully-actuated subsystems and two under-actuated subsystems, and the sliding surface of each subsystem is constructed by adopting proper sliding surface design method. A continuous adaptive first-order sliding mode disturbance estimator is adopted to online approximate the uncertainties of each subsystem, and a two-phase combined function reaching law that can dynamically adapt to the variation of the sliding surface is used to sequentially design the control amount of the fully-actuated subsystems and the under-actuated subsystems. The stability of the closed-loop control system and the convergence time of the sliding surface of the control system are theoretically analyzed. The simulation test results verify the robust control performance and controller chattering reduction ability of the proposed sliding mode control strategy.

To address the issue of hit accuracy of terminal guidance phase of a rotor loitering terminal sensitive projectile, that is jointly affected by detection system errors and attitude disturbances, this study focuses on the dynamic factors which affect the dispersion of EFP warhead destruction element. In the study, a dynamic impact point model is established based on the seeker system detection error, airframe's spatial attitude disturbances, implicated motion deviation of the warhead destruction element and system latency, meanwhile the trajectory of missile target rendezvous and swinging attitude correction in the final guidance phase are utilized. By taking full advantage of airframe's attitude information and with the help of experimental flight data, the model can be used to analyze the influence of each factor on the point of impact. Simulation and analysis results demonstrate that, under certain random disturbances, the primary factor causing aiming error is the airframe's attitude disturbances and rendezvous trajectory deviation, so that gain the attitude angle control requirements necessary to achieve a desired hit rate on a typical armored target. This study provides a theoretical basis for the research on guidance law and trajectory prediction control of loitering terminal sensitive projectile.

Dynamic explosion point localization constitutes a pivotal element in damage assessment. The explosive sound wave, originating from the attenuation of a shock wave, is characterized by supersonic velocity and a defined sphere of influence. Accordingly, the accuracy of explosion point acoustic localization is susceptible to the influence of the explosive shock wave. To explore the impact of the explosive shock wave segment on the acoustic localization outcomes during the explosion point localization process, a wave arrival time-distance model for the attenuation of the shock wave into a sound wave was constructed, predicated on the pressure attenuation law of the explosive shock wave. This model was subsequently integrated into the localization methodology based on the time difference of arrival (TDOA). Utilizing a prototypical arrangement as a case study, the preset explosion point and the acoustic sensor array's position were delineated, and the corresponding wave arrival times were computed. The TDOA-based localization approach was employed to retrodict the position of the explosion point relative to the acoustic sensor array, and the acoustic localization error of the explosion point was ascertained by juxtaposing the retrodicted position with the preset position. Furthermore, by modulating the preset explosion point position and the charge mass, the influence of the explosive shock wave on the acoustic localization error was scrutinized. The research findings suggest that the charge mass and the distance from the explosion center are the predominant factors influencing the localization error. As the scaled distance escalates, the localization error induced by the shock wave exhibits a diminishing trend. When the scaled distance surpasses 18.45, the relative distance error can be mitigated to below 1%. It is anticipated that these findings will serve as a reference for the systematic error correction of explosion point acoustic localization, thereby enhancing the computational precision of explosion point acoustic localization.

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.

Aiming at the problems of large errors and insufficient adaptation to meteorological changes in existing ballistic drop prediction methods, a ballistic dataset containing meteorological conditions is established and a CNN-BiLSTM-BiGRU ballistic drop prediction method based on a self-attention layer is proposed in this paper. The self-attention layer and residual connection are introduced into the constructed combined model to strengthen the model's ability to dynamically focus on the information at different moments when processing the input sequences, and to alleviate problems such as gradient explosion in the network. It also uses the input representation of multi-dimensional time series data to reduce the ballistic drop prediction error by combining multiple information such as historical ballistic trajectory data and target characteristics. The simulation results show that the prediction effect of the CNN-BiLSTM-BiGRU network model based on the self-attention layer is better than the other models, and the maximum error of the range prediction accounts for 0.156% of the true value, and the maximum error of the lateral deviation prediction accounts for 5.904% as of the true value. The method provides an important reference for the field of ballistic drop prediction.

Heterogeneous scene matching, as an important auxiliary navigation method, has been widely studied. However, due to the influence of nonlinear radiation distortion and geometric deformation between heterogeneous image pairs, achieving heterogeneous image matching remains a challenging task. To address these issues, a heterogeneous scene matching method with rotation and scale invariance is proposed to simultaneously estimate the rotation, scale, and displacement variations between heterogeneous image pairs. Firstly, based on the local structural relationships of the image, local self-similarity descriptors (LSS) are used for feature description to resist the influence of nonlinear radiation differences and local deformations. Combined with the logarithmic polar coordinate transformation, the overall rotation and scale changes of the image are orthogonally expanded and represented separately in the Cartesian coordinate system. Finally, by utilizing the continuity of displacement estimation, rotation, and scale estimation, a five-dimensional feature descriptor is constructed and using phase correlation method estimates the variation of image rotation, scale and displacement simultaneously. Experiments conducted on three common types of heterogeneous image matching tasks shows that the proposed method achieves a matching accuracy of at least 4.5% higher than existing state-of-the-art methods tech, that highlights its effectiveness in the field of heterogeneous scene matching.

In order to strengthen the detection capability of small infrared targets under complex background, an improved WSLCM(weighted local contrast measure) detection algorithm is proposed. In the pre-processing stage of WSLCM, adaptive curvature filtering is used to process the image in multi-scale, and the real object is not submerged in the process of background suppression. In the background suppression calculation, the maximum gray value of the background block is selected as the background estimation to reduce the false alarm rate. At the same time, target enhancement factor and background suppression factor are introduced to enhance the robustness of the algorithm, eliminate the influence of background noise, and enhance the detection ability of infrared small targets. The algorithm IP verification is carried out by the embedded ZYNQ platform, and the algorithm can recognize and detect the small target in the specific scene by using the hardware and software cooperation. Experiments show that compared with the traditional WSLCM algorithm, BSF and SCRG indexes are significantly improved, the continuous frame detection rate is 93.2%, and the detection efficiency of embedded platform is 17.6% higher than that of PC, which verifies the effectiveness of the algorithm and embedded system.

In order to investigate the effect of combined liner structure on the damage performance of shaped charge warhead, the arbitrary Lagrange-Eulerian fluid-structure coupling algorithm based on LS-DYNA software was used to study the formation of penetrator of single EFP shaped charge warhead and different structural combination of shaped charge warhead, as well as the damage of penetrator to target, after verified the effectiveness of the selected model and material parameters. The results indicate that with the addition of top liner before a single EFP liner, the penetrator with higher head velocity and longer configuration can be formed under the shaped charge. Under the condition that the bottom diameter and height of top liner are 0.2 times the diameter of the charge, when the top cone angle is 30°, the head velocity of the penetrator formed by the combined liner is higher than that formed by other combined liner, and the length of penetrator formed by the combined liner is longer than that formed by other combined liner. However, the diameter of the penetrator formed by the combined liner is the smallest, which is related to the propagation of stress waves in the interior of the liner and whether the microelements around the liner can satisfy the requirement of lengthening the penetrator. The hole aperture and through hole aperture under the penetration of combined liner are positively correlated, the hole aperture and through hole aperture are negatively correlated with the penetration depth under the penetration of combined liner. Under the condition that the bottom diameter and the height of the top liner are 0.2 times the diameter of the charge, the hole aperture of the target under the penetration of the combined liner is not significantly different from that under the penetration of the single EFP liner. The through hole aperture of the target under the penetration of the combined liner is about 0.17~0.2 times of the through hole aperture of the target under the penetration of the single EFP liner. The penetration depth of the target under the penetration of combined liner is about 2.7~3.7 times of the penetration depth of the target under the penetration of single EFP liner. According to the power requirements of warhead, the penetration depth of the target can be increased on the basis of ensuring the hole aperture by adding a proper top liner with a suitable structure.

Aiming at the problem of multi-missile cooperative attack on enemy air targets, a multi-missile and multi-constraint three-dimensional cooperative guidance method based on sliding mode control theory and artificial potential field method is deduced. Firstly, according to the relative motion of the missile and the target, a three-dimensional non-linear model of the relative motion between the missile and the target was established in the line-of-sight coordinate system. Then, taking the remaining flight time of multiple bombs as the coordination variable, based on the finite time consistency theory, the guidance law of the direction of sight of multiple bombs is designed, which realizes the coordinated attack on the target. On this basis, the normal direction of sight guidance law of multi-missile is designed through the sliding mode control theory, so that the multi-projectile can hit the enemy target at a given line of sight angle, and according to the convergence characteristics of the sliding mode surface, the parameters are adaptively updated to reduce the sliding mode parameter selection complexity. In addition, considering the difficulty in obtaining target maneuver information, an expanded state observer is designed to efficiently predict target acceleration information. Finally, combined with the idea of artificial potential field method, a multi-missile obstacle avoidance control command is designed to effectively avoid constraints such as avoidance zones during multi-missile flight. The simulation results verify the effectiveness of the proposed guidance method.

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%.

The influence of position effect and interface effect on the process of a tungsten alloy rod-shaped penetrator vertically penetrating a ceramic/aluminum alloy composite target plate is studied through experimental research and simulation analysis. The effect of the position effect on the protective performance of the target plate is examined by changing the impact point, while the effect of the interface effect on the protective performance of the target plate is investigated by varying the number of ceramic layers. Research on positional effects shows that when the impact point is within the range of 0 to 30 mm from the ceramic boundary, the penetration depth of the back-plane changes by 9 mm. When the impact point is within the range of 30 to 50 mm from the ceramic boundary, the penetration depth of the back-plane changes by 1 mm. It can be concluded that the protective performance of the target plate in this area is no longer influenced by the position of the impact point. In other words, the area that is greater than or equal to five times the bullet diameter from the ceramic boundary is the effective protection area of the target plate. Research on the interface effect indicates that the penetration depth of the composite target plate with three different structures at the same thickness is 40.32 mm, 47.47 mm, and 50.77 mm; that is, the ceramic protective performance decreases with the increase in the number of layers at the same thickness.

For missile's precision guidance task, an optimal leading angle sliding mode control algorithm based on GA-BP (genetic algorithm-back propagation) is proposed, in which the sliding mode control is improved segmentally. Aiming at the problem of the fixed leading angle sliding mode control depending on a leading angle value, but it is difficult to be predefined, a GA-BP neural network is developed and used to estimate the optimal leading angle for a specific task model. Subsequently, based on the estimated value of remaining time, a piecewise sliding mode reaching law is designed, which lead to a segment-improved sliding mode control algorithm that further enhances robustness during the missile guidance process. And then the complete optimal leading angle segment-improved sliding mode algorithm based on GA-BP is formed. After the algorithm is simulated and analyzed, Simulation results indicate that, compared with the fixed leading angle sliding mode control algorithms, the proposed algorithm can reduce task time approximately 5% in average, and the overload around 12%. As it can reduce the task time by up to 30% in the maximum, so it has a higher superiority.

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.

In view of the outstanding problems existing in the assessment model, method, system and process of assessment of target vulnerability, such as numerous and disorderly assessment models, methods and processes, and poor timeliness and accuracy of assessment system, it is necessary to deeply study the common process of assessment of target vulnerability including assessment model, method and system. This paper mainly uses theoretical analysis, example illustration and other methods to explore it. A common process of assessment of target vulnerability is proposed, it mainly includes characteristic analysis of target, characteristic analysis of damage element, classification of target damage levels, analysis of target critical components, analysis of target damage characteristic and assessment of target vulnerability. Based on this, a system framework for target vulnerability assessment is proposed, which includes the subsystem of target characteristics, the subsystem of warhead power and characteristics of damage element, the subsystem of warhead coincides with target, the subsystem of characteristics of target damage, the subsystem of vulnerability assessment, as well as the subsystems of preprocessing and postprocessing and data management. Integrate artificial intelligence into the assessment system to enhance the intelligence level of the assessment system of target vulnerability. At the same time, each subsystem should be open, allowing content to be added in a standardized format, continuously enhancing its standardization and versatility.

The aerodynamic loads, multi-body motion, and contact collisions involved in the separation process of rocket and missile launches are key factors affecting the reliability of launch separation. To adopt numerical methods for analyzing such coupled states of flow and motion, an in-depth study is conducted on the launch separation flow and collision coupling model based on a virtual contact approach. The coupling model starts from a transient flow field numerical model, that is introduced a dynamics model concerning the six degrees of freedom (6-DOF) multi-body motion, subsequently, a contact dynamics model characterized by virtual contact is used to examine the multi-body contact collision loads to obtain the coupled states of object motion and flow under various loads. The analysis of application examples shows that the virtual contact method can effectively simulate the coupled state of flow, motion, and collision during launch separation and the method can provide a reference for similar analyses.

In order to study the low energy impact resistance of re-entrant star-shaped (RES) and re-entrant circular star-shaped (RECS) honeycomb sandwich panels with negative Poisson ratio, a finite element model of honeycomb sandwich panels with composite laminates and ABS honeycomb is established by using commercial software ANSYS/LS-DYNA. Impact response of the composite sandwich panels with RES honeycomb and RECS honeycomb is analyzed numerically at low impact energies of 10 J,15 J,30 J. In this paper, the impactor displacement versus time curves、contact force versus time curves and energy absorption ratio of sandwich panels under different impact conditions are analyzed and compared. The simulation results show that under the same impact conditions, the damage depth of RECS honeycomb sandwich panels is greater than that of RES honeycomb sandwich panels, and the difference is more obvious when under in-plane impact. In addition, the impactor has a lower rebound speed after impacting RECS honeycomb sandwich panels, that is, the RECS honeycomb sandwich panel has a higher energy absorption ratio. According to the above results, it is concluded that RES honeycomb sandwich panel has better low energy impact resistance than RECS honeycomb sandwich panel, but the latter has better energy absorption performance.

For the segmented combustion instability phenomenon of slender solid rocket motors in the ground test, it can obtained the distribution range of pressure oscillation frequency in the combustion chamber by mean of acoustic cavity frequency analysis, analyzing the main factors of two-stage combustion instability from a perspective of acoustic-vortex coupling and propellant combustion response. Then through comparison experiment, comparing the combustion instability of the motor with the same grain configuration and different propellant formulations and combiningewith the acoustic cavity mode of the combustion chamber at the initial, middle and end time and the flow field vortex structure computation of the combustion chamber and the pressure coupled response function test results of the T-burner, the conclusion is that the combustion instability phenomenon at the initial stage of the motor operation is due to the fact that the vortex shedding frequency caused by the grain configuration is close to the frequency of the acoustic field of the chamber to generate coupling gain and induce the pressure oscillation. At the end of the operation of the motor, the combustion instability frequency is different from the initial frequency range, and there are frequency multiplication characteristics, the instability is caused by the combustion response of the propellant. The analysis unfolds from the two-stage combustion instability in the ground test andprovids a reference to optimize solid rocket design and effective avoidance of such problems in engineering 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.

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.

At present, scholars carry out anti-missile damage assessment work basically simplifies the assessment model and assumes the derivation conditions, which has the problem of insufficient credibility. It is an effective means to enhance the credibility by injecting external fitted attitude information into the semi-physical system to reproduce the flight process of the missile, so as to carry out the damage assessment under the real ballistic rendezvous conditions. In this paper, firstly, the equations of motion of the missile model in the relative coordinate system of the target are given through the ballistic fitting of the semi-physical simulation system, and then according to the empirical formulas of the probability of the distribution of the guidance error, the probability of the distribution of the fuse activation zone and the probability of the destruction of the combat unit, combined with the semi-physical projectile-eye rendezvous conditions, the method of the determination of the integrated probability of the destruction is given. Finally, this paper carries out simulation calculations for specific examples, and analyzes the parameter characteristics and damage classifications of the simulation results. The results show that the relationship between the distance of the meeting and the azimuth angle of the fragmentation incident and the damage result, and the established damage evaluation table can predict the interception result of the interceptor to a certain extent, which can provide support and reference for the subsequent military operational decision-making.