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.

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.

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.

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.

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%.

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.

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.

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 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.

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.

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.

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.

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.

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.

Based on the study of the mechanism of high-temperature gas ablation on the surface of the flame guide groove during rocket launch, this paper proposes an innovative algorithm for predicting the maximum temperature of the guide groove surface. The algorithm uses Latin hypercube sampling technology to obtain sample data, and uses the Kriging fitting algorithm to construct a proxy model based on the sample data to predict the maximum temperature value of the guide groove surface under different profile parameters. Combining sample data analysis with optimized configuration calculation, the influence law of the structural size parameters of the guide groove on its surface temperature distribution is further analyzed. On the basis of the established proxy model, the genetic algorithm is used to calculate that the guide groove surface temperature reaches the lowest value when the basic straight section length is 2 084.2 mm, the guide surface curvature radius is 2 276.0 mm, and the guide surface inclination angle is 62.2°. By numerically simulating the gas flow field under different combinations of guide groove size parameters, the distribution characteristics of the high-temperature and high-pressure areas on the guide groove surface are studied, and the optimized configuration is compared with the original optimal working condition in terms of pressure gradient and other aspects. The results show that in the outer area of the guide groove, the high temperature and high pressure areas formed by the direct impact of the gas are more concentrated; compared with the optimized configuration, the pressure gradient distribution near the outer side of the original optimal working condition is more significant, which has a certain hindering effect on the gas flow.This causes the gas temperature to rise after compression, thereby maintains the surface temperature of the guide groove at a higher level. The research conclusion provides a theoretical basis and reference value for the optimization design of the guide groove of the rocket launcher.

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.

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.